Hani Kamhawi, David H. Manzella, Timothy D. Smith, and George R. Schmidt

Glenn Research Center, Cleveland, Ohio

High-Power Hall Propulsion Development at

NASA Glenn Research Center

NASA/TM—2014-218084

April 2014

IAC–12–C4.4.6
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Hani Kamhawi, David H. Manzella, Timothy D. Smith, and George R. Schmidt

Glenn Research Center, Cleveland, Ohio

High-Power Hall Propulsion Development at

NASA Glenn Research Center

NASA/TM—2014-218084

April 2014

IAC–12–C4.4.6

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

63rd International Astronautical Congress (IAC 2012)

sponsored by the International Astronautical Federation

Naples, Italy, October 1–5, 2012
Acknowledgments

The authors would like to thank and acknowledge the Offi ce of the Chief Technologist (OCT) Game Changing Division for 

funding this work. Contributions in this paper came from various members of the OCT In-Space Propulsion (ISP) team at NASA 

Glenn Research Center including Li Chang, Lauren Clayman, Thomas Haag, Daniel Herman, Wensheng Huang, Luis Pinero, 

Rohit Shastry, George Soulas, Jonathan Van Noord, and John Yim. The authors would also like to thank that many technicians 

that supported the execution and completion of the many tasks highlighted in this paper including Kevin Blake, Mark Forrester, 

Larry Hambley, George Jacynycz, Terrell Jansen, John Lauerhahs, Kevin McCormick, Michael Pastel, Derrick Patterson, and 

George Readus.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2014-218084 1 

High-Power Hall Propulsion Development at 

NASA Glenn Research Center 

Hani Kamhawi, David H. Manzella, Timothy D. Smith, and George R. Schmidt 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The NASA Office of the Chief Technologist Game 

Changing Division is sponsoring the development and testing 

of enabling technologies to achieve efficient and reliable 

human space exploration. High-power solar electric 

propulsion has been proposed by NASA’s Human Exploration 

Framework Team as an option to achieve these ambitious 

missions to near Earth objects. NASA Glenn Research Center 

(NASA Glenn) is leading the development of mission 

concepts for a solar electric propulsion Technical 

Demonstration Mission. The mission concepts are highlighted 

in this paper but are detailed in a companion paper. There are 

also multiple projects that are developing technologies to 

support a demonstration mission and are also extensible to 

NASA’s goals of human space exploration. Specifically, the 

In-Space Propulsion technology development project at 

NASA Glenn has a number of tasks related to high-power Hall 

thrusters including performance evaluation of existing Hall 

thrusters; performing detailed internal discharge chamber, 

near-field, and far-field plasma measurements; performing 

detailed physics-based modeling with the NASA Jet 

Propulsion Laboratory’s Hall2De code; performing thermal 

and structural modeling; and developing high-power efficient 

discharge modules for power processing. This paper 

summarizes the various technology development tasks and 

progress made to date. 

Acronyms 

BN boron nitride 

DT,m mean thruster diameter, meter 

EDU engineering development unit 

EP electric propulsion 

ESMD Exploration Systems Development Directorate 

ESR&T Exploration System Research and Technology 

ETDD Engineering Technology Development and 

Demonstration 

GaN gallium nitride 

GCD Game Changing Division 

HARP high-speed axially reciprocating probe 

HEFT Human Exploration Framework Team 

HiVHAc High Voltage Hall Accelerator 

HEOMD Human Exploration and Operations Mission 

Directorate 

IB,250 intensity of boron at a wavelength of 250 μm 

Id discharge current, A 

IEPC International Electric Propulsion Conference 

ISP In-Space Propulsion 

Isp specific impulse, sec 

ISTP In-Space Transportation Program 

JIMO Jupiter Icy Moons Orbiter 

JPL Jet Propulsion Laboratory 

MFAM magnetic-field-aligned computational mesh 

MOSFET metal oxide semiconductor field effect transistors 

NASA National Aeronautic and Space Administration 

NEO near Earth objects 

OES optical emission spectroscopy 

OCT Office of the Chief Technologist 

PMAD power management and distribution 

PPU power processing unit 

RPA retarding potential analyzer 

SAA space act agreement 

SCCM standard cubic centimeter per minute 

SEP solar electric propulsion 

SiC silicon carbide 

TDM Technical Demonstration Mission 

Vd discharge voltage, V 

VF5 vacuum facility 5 

VF8 vacuum facility 8 

ηA anode efficiency 

I. Introduction 

High-power electric propulsion (EP) systems are enabling 

and enhancing for time-critical missions or missions requiring 

transportation of large payloads. A number of mission studies 

were performed over the last decade which highlight the 

enhancing and enabling features of high-power EP systems for 

reusable space tug applications for transfer of payloads from 

low-Earth-orbit to geosynchronous-Earth-orbit and for use in 

Mars mission scenarios (Refs. 1 to 3). 
NASA/TM—2014-218084 2 

National interest in high-power EP systems has been 

renewed. In 2010, NASA’s Human Exploration Framework 

Team (HEFT) concluded that the use of a high-power (i.e., on 

the order of 300 kW) solar electric propulsion (SEP) system 

could significantly reduce the number of heavy lift launch 

vehicles required for a human mission to a near Earth asteroid 

(Ref. 4). Hall thrusters are ideal for such applications because 

of their high-power processing capabilities and their efficient 

operation at moderate specific impulses (~2000 sec), which 

leads to reduced trip times for such missions (Ref. 5). 

NASA’s Human Exploration and Operations Mission 

Directorate (HEOMD) Enabling Technology Development 

and Demonstration (ETDD) Program was focused on 

developing, maturing, testing, and demonstrating the 

technologies needed to reduce the cost and expand the 

capability of future space exploration activities. The ETDD 

Program content included performing foundational research 

and studying of the requirements and potential designs of 

advanced, high-energy in-space propulsion systems. These 

high-energy propulsion systems were intended to support deep 

space human exploration and reduce travel time between 

Earth’s orbit and future destinations for human activity. This 

would enable a new space transportation capability via a SEP 

stage. The SEP stage could enable cost-effective missions 

within Earth orbit, near Earth objects (NEOs), and deep space 

robotic science missions. In 2011, the ETDD Program 

transitioned to NASA’s Office of the Chief Technologist 

(OCT) but the program content remained and is still focused 

on developing, maturing, and demonstrating the high-power 

propulsion technologies needed to enhance the Agency’s 

capabilities to explore and move large payloads in space. To 

that extent, there are several projects in the OCT Game 

Changing Division (GCD) that support technology 

development which will result in the SEP Technical 

Demonstration Mission (TDM) flight development. One of the 

key projects under OCT GCD is the In-Space Propulsion (ISP) 

project. This paper provides a brief overview of the SEP TDM 

project, additional details of the SEP TDM project are 

provided in a companion paper (Ref. 6). This paper provides a 

detailed summary of the OCT ISP technology tasks and 

activities over the past 2 years. 

II. Solar Electric Propulsion Technical 

Demonstration Mission 

The SEP TDM project is currently a pre-Phase A 

formulation activity focused on developing high value mission 

concepts for a demonstration mission that address vehiclelevel SEP-integrated system issues while simultaneously 

retiring technology risks that cannot be adequately simulated 

on the ground. The advanced technology utilized by any TDM 

concept maybe provided through the technology maturation 

activities projects being conducted by the OCT. These 

technologies may include but are not limited to high-power 

Hall thruster propulsion systems, advanced high-voltage solar 

array systems, and advanced power management and 

distribution topologies including direct drive. To date, the SEP 

TDM project has developed a series of seven mission concepts 

through contracts and Space Act Agreements (SAAs) with 

industry. These agreements were with the Aerojet 

Corporation, Analytical Mechanics Associates, Inc., the Ball 

Aerospace & Technologies Corporation, The Boeing 

Company, the Lockheed Martin Space Systems Company, the 

Northrop Grumman Systems Corporation, and Space Systems 

Loral. The four objectives being addressed by these mission 

concepts studies were: 

 Perform an in-space demonstration that advances the 

maturity of high-power EP technology and high-power 

solar array power system technology in relevant space 

environments and operational regimes. 

 Demonstrate integrated SEP spacecraft design, 

fabrication, and test as well as operational modes 

associated with orbit transfer. 

 Demonstrate extensible high-power EP and solar array 

power system technologies and integrated SEP 

spacecraft operational modes that can be adapted for use 

in next-generation, higher power SEP systems. 

 Provide a SEP-based transportation capability with 

performance advancements over those previously 

demonstrated. 

 More details about these mission concepts and the SEP 

TDM project can be found in the companion paper by 

Smith et al. (Ref. 6). 

III. In-Space Propulsion Technology 

Project 

The NASA Glenn Research Center (NASA Glenn) OCT 

ISP technology program is testing, developing, and maturing 

high-power Hall thruster technologies that will enable the 

Agency to achieve its future objectives of robotic and human 

space exploration. To that extent, in 2011 and 2012, NASA 

Glenn high-power Hall development activities and tasks were 

focused on the following: 

 Evaluating performance of existing high-power Hall 

thrusters (NASA-300M and NASA-457Mv2) 

 Developing plasma diagnostics for Hall thrusters 

 Performing Hall thruster physics based modeling using 

the NASA Jet Propulsion Laboratory (JPL) Hall2De 

code 

 Performing Hall thruster structural and thermal 

modeling 

 Designing, fabricating, and testing high-current cathode 

assemblies with an emission current capability  100 A 

and lifetime capability greater than 20,000 hr 

 Developing and demonstrating high-power (300 V and 

7 kW) discharge modules for power processing units 
NASA/TM—2014-218084 3 

The following sections will summarize the results from 

these activities. 

A. High-Power Hall Thruster Testing 

NASA Glenn has a long history of researching and 

developing high-power Hall thrusters and ion engines. NASA 

Glenn’s high-power Hall thruster research and development 

efforts started in 1997 with the award of the 

T-220 10 kW Hall thruster development contract to TRW, 

Space Power Inc., and the Keldysh Research Center (Ref. 7). 

The T-220 thruster was fabricated and delivered to NASA 

Glenn in 1998 and the T-220 underwent extensive 

performance evaluation and was later subjected to a 1000 hr 

wear test (Ref. 8). In 2000, the In-Space Transportation 

Program (ISTP) funded the development of two 50 kW class 

Hall thrusters: one by the Aerojet Redmond Rocket Center and 

one by NASA Glenn. The thruster developed by NASA Glenn 

was designated the NASA-457Mv1 (Ref. 9). This thruster was 

tested at power levels up to 72 kW and results revealed key 

areas in the design that could be improved to increase the 

performance and reduce mass (Ref. 10). A new thruster 

designated the NASA-457Mv2 was designed and built. 

Unfortunately, in 2004 the high-power ISTP program was 

ended and the NASA-457Mv2 performance was never 

evaluated (Ref. 11). 

In 2003, NASA’s project Prometheus was initiated with the 

goal of developing nuclear power and propulsion technologies 

that would enable the exploration of the universe in search of 

life and resources. The Jupiter Icy Moons Orbiter (JIMO) 

mission would utilize these newly developed technologies. 

Mission analysis indicated the optimal high-power EP device 

for JIMO was a high-power, high-specific impulse Hall 

thruster. As such NASA Glenn initiated the development of 

the NASA-400M (Ref. 11). The thruster leveraged and 

incorporated lessons learned from the design and testing of the 

NASA-457Mv1. To be able to achieve high specific impulse 

magnitudes the thruster was designed to operate with krypton 

propellant at discharge voltages up to ~1000 V. A 300 hr wear 

test of the thruster indicated that the boron nitride (BN) 

discharge channel erosion rates were similar to rates measured 

during the wear test of the T-220 (Refs. 11 to 12). 

NASA Glenn’s long history of developing and testing of 

high-power Hall thrusters and its world-class vacuum facilities 

makes it the leading NASA center for developing high-power 

EP devices. In addition, complete evaluation of previously 

developed high-power thrusters provides a great start for 

NASA’s effort to develop a high-power SEP stage. As such, 

under the OCT ISP technology project, NASA Glenn was 

tasked with assessing the performance of the existing highpower Hall thrusters that were of interest to SEP TDM and 

were extensible to the future 300 kW NEO mission. The next 

two sections will summarize performance evaluation results 

from recent performance tests of the NASA-300M (20-kWclass) and NASA-457Mv2 (50-kW-class) thrusters. 

A.1. NASA-300M 20-kW Hall Thruster 

The NASA-300M 20-kW-class Hall thruster was designed, 

fabricated, and assembled at NASA Glenn in 2004 and 2005. 

The NASA-300M was designed and fabricated under the 

support of the Exploration Systems Mission Directorate 

(ESMD) Exploration System Research and Technology 

(ESR&T) program. The NASA-300M design is based on a 

scaled version of the NASA-457Mv2 and incorporated lessons 

learned from the development and testing of the NASA457Mv1, NASA-400M, and NASA-457Mv2 thrusters 

(Ref. 11). The goal of the design was to minimize the size of 

the design while optimizing the magnetic field and plasma 

lens to improve performance. The NASA-300M nominal 

design specifications were discharge power of 20 kW, 

discharge voltage range up to 600 V, discharge current up to 

50 A, and a magnetic circuit that has a similar magnetic field 

topology as the NASA-457Mv2. The emergence of the 

ETDD/OCT ISP programs resulted in renewed interest in the 

NASA-300M. 

Testing of the NASA-300M was performed in vacuum 

facility 5 (VF5) at NASA Glenn. The VF5 main cylindrical 

chamber is 4.6 m in diameter and 18.3 m long. Its main port 

(designated E55) is 1.8 m in diameter and 2.5 m long. The 

VF5 can be evacuated with cryopanels and oil diffusion 

pumps. Figure 1 shows the VF5 and the NASA-300M 

mounted inside the main port. 

Under ETDD, testing of the NASA-300M was performed to 

discharge power levels up to 20 kW with discharge voltages 

up to 600 V and discharge currents up to 50 A. Testing was 

performed with xenon and krypton propellants (Ref. 13). 

Xenon propellant testing was performed for power levels 

between 2.5 and 20 kW at 2.5 kW increments. Krypton 

propellant testing was performed for power levels between 5 

and 20 kW in 5 kW increments. Both propellants were tested 

at discharge voltages between 200 and 600 V. The anode mass 

flow rate (i.e., discharge current) was varied to attain the 

prescribed power level at a given discharge voltage; however, 

the discharge current was not allowed to exceed 50 A due to 

hollow cathode operational considerations. The cathode flow 

rate was maintained at approximately 8 percent of the anode 

flow rate. Peak facility pressure during thruster operation at 

50 A was 2.8×10–5 Torr (corrected for xenon). Testing at all 

power levels and discharge voltage conditions indicated that 

the thruster operation was stable and no anomalous behavior 

was observed. In addition, when testing at 20 kW, no thermal 

limitations on thruster operation were observed. Figure 2 

shows the thruster while operating on xenon and krypton at 

20 kW. 
NASA/TM—2014-218084 4 

 

Figure 1.—Photographs of VF5 (left) and the NASA-300M thruster (right) installed in VF5. 

Figures 3 and 4 present the xenon propellant total specific 

impulse and thrust efficiency results, respectively, for the 

NASA-300M. These results indicate that the total specific 

impulse and thrust efficiency increased with increasing power 

at a given discharge voltage. At 20 kW, NASA-300M 

performance results indicate that it is able to achieve a peak 

specific impulse of about 2900 sec at a discharge voltage of 

600 V and a peak thrust efficiency of 67 percent at a discharge 

voltage of 500 V. 

Figures 5 and 6 present the krypton propellant total specific 

impulse and total thrust efficiency results, respectively. 

Results in Figures 5 and 6 indicate that, in general, total 

specific impulse and thrust efficiency increased with 

increasing power at a given discharge voltage (except for Vd

of 200 V). The highest attained total specific impulse and total 

thrust efficiency were 3223 sec and 63 percent, respectively, 

when operating at a discharge voltage of 600 V at 20 kW. 

Detailed performance evaluation results of the NASA-300M 

can be found in Reference 13. 

Figure 3.—NASA-300M xenon total specific impulse 

magnitudes. 

Figure 4.—NASA-300M xenon total thrust efficiency 

magnitudes. 

Figure 2.—Photograph of the NASA-300M operating on 

xenon (left) and krypton (right) at a discharge power 

level of 20 kW. 
NASA/TM—2014-218084 5 

Figure 5.—NASA-300M krypton total specific impulse 

magnitudes. 

Figure 6.—NASA-300M krypton total thrust efficiency 

magnitudes. 

A.2. NASA-457Mv2 50 kW Hall Thruster 

A photograph of the higher fidelity 50 kW Hall thruster 

(i.e., the NASA-457Mv2), is shown in Figure 7. This highpower thruster has a solid BN discharge chamber with a 

457 mm outer diameter and a centrally mounted high-current 

hollow cathode. The development goal for the NASA-457Mv2 

was to advance its technology readiness level from 4 to 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. 11). 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 

457Mv2 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 

Figure 7.—Photograph of the NASA-457Mv2 thruster. 

amp-turns (Ref. 11). The magnetic field produced by the 

NASA-457Mv2 thruster was measured and compared to its 

laboratory model predecessor. Although measured magnetic 

fields for the laboratory model thruster were already reported 

in References 9 and 16, they were remeasured 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. 11). 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. 

The operating conditions 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. 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. Performance was 

measured in 100 V increments; it was also 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. 
NASA/TM—2014-218084 6 

TABLE 1.—DISCHARGE THROTTLE LEVELS FOR THE 

NASA-457M v2 THRUSTER PERFORMANCE TEST 

Discharge 

power, 

kW 

Discharge current, A versus 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 --- 25 --- 

15 a75 60 50 --- 37.5 30 

20 a100 80 66.7 57.1 50 40 

25 100 83.3 --- 62.5 50 

30 100 85.7 75 60 

35 - 87.5 70 

40 100 80 

45 90 

50 100 a

d means discharge (1-e, e referrers to exponent). 

Figure 8.—Thrust as a function of total thruster input power at 

various discharge voltages for the NASA-457Mv2 thruster 

operating with high anode efficiency. 

Thrust as a function of total thruster input power is plotted 

in Figure 8 for the NASA-457Mv2 thruster operating with 

high anode efficiency. The uncertainty in total power for all 

results reported herein was estimated to be 0.3 percent of the 

indicated values. The peak thrust demonstrated by the NASA457Mv2 was 2.3 N at a total thruster input power of 50 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. 9 and 16). This was because the total 

input power was dominated by discharge power. 

Specific impulse as a function of total thruster input power 

is shown in Figure 9 for the NASA-457Mv2 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 NASA-457Mv2 

thruster. The thruster demonstrated specific impulse range of 

Figure 9.—Specific impulse as a function of total thruster input 

power at various discharge voltages for the NASA-457Mv2 

thruster operating with high anode efficiency. 

1420 to 2740 s and also 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. 17). 

The thrust efficiency as a function of total thruster input 

power is shown in Figure 10 for the NASA-457Mv2 thruster 

operating with high anode efficiency. The uncertainty in thrust 

efficiency was estimated typically to 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 NASA-457Mv2 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 NASA-457Mv2 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 

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

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
NASA/TM—2014-218084 7 

Figure 10.—Thrust efficiency as a function of total thruster 

input power at various discharge voltages for the NASA457Mv2 thruster operating with high anode efficiency. 

 

Figure 11.—Photographs of the probe array used in this study. Left: Photograph showing order of four probes 

within array, prior to operation. Right: Array sweeping near the thruster exit plane of the 300M (array is 

shown to the left of thruster). 

specific impulses would be about 2000 s, thrust efficiencies 

varied from 0.57 to 0.63. Detailed performance evaluation 

results of the NASA-457Mv2 can be found in Reference 18. 

B. Near-Field Plasma Measurements 

A number of near-field plasma measurements were taken in 

order to better understand high-power Hall thruster physics 

and characterize performance. The purpose of these 

measurements is to provide data that will help elucidate the 

physics governing high-power Hall thruster operation and to 

characterize thruster performance and efficiency. Another 

purpose is to provide experimental data to validate existing 

modeling efforts at Glenn and the JPL (Ref. 19). The 

experimental data along with model results will be used to 

refine thruster design and will help develop a life qualified 

thruster. This will ultimately lead to the design of a higher 

performance high-power Hall thruster with a long lifetime at a 

lower cost. 

To reach these goals, an electrostatic probe array was used 

to measure the near-field plume of the 20-kW NASA-300M 

and the NASA-457Mv2 from approximately 0.1 to 2 mean 

thruster diameters (DT,m) downstream of the thruster exit 

plane. Similar measurements have been taken before on the 

6-kW H6 Hall thruster at JPL (Ref. 20). The probe array 

consisted of a near-field Faraday probe, a single Langmuir 

probe, and two emissive probes (one for redundancy). 

Figure 11 shows the electrostatic probe array. Measurements 

at these distances were facilitated by a high-speed linear 

motion table typically run between 250 and 

375 mm/s, helping to ensure probe survivability in the harsh 

near-field environment. In addition, optical emission 

spectroscopic measurements were also performed. The next 

two sections will highlight findings from those studies. 

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—2014-218084 8 

Figure 12.—Operating conditions for the ion current 

density measurements for the NASA-300M and 

NASA-457Mv2 thrusters. 

B.1.Ion Current Density Measurements 

Ion current density plume measurements provide an 

accurate estimate of the plume divergence in a Hall thruster. 

These measurements are typically performed with a Faraday 

probe (see Fig. 11). Details of the probe design, motion table 

setup, and the data acquisition electrical circuit setup is 

detailed in References 21 to 23. For this test, the NASA-300M 

was operated at seven conditions spanning discharge voltages 

from 200 to 500 V, discharge currents from 20 to 67 A, and 

discharge powers of 10 and 20 kW. The nominal operating 

condition was 500 V and 40 A. The BN channel walls of the 

NASA-300M were largely uneroded for this present test. 

These channel walls have accumulated at most 70 hr prior to 

the start of the tests described in this section. 

For this test, the NASA-457Mv2 was operated at 14 

conditions spanning discharge voltages of 200 to 500 V, 

discharge currents of 40 to 100 A, and discharge powers of 10 

to 50 kW. The nominal operating condition was 500 V and 

100 A. The anode efficiency of the NASA-457Mv2 varies 

from 55 to 70 percent over the tested operating conditions. 

The test matrixes of the near-field Faraday probe test were 

chosen to target a variety of discharge voltages in order to 

characterize the thrusters for both high thrust-to-power (low 

specific impulse) and high specific impulse operation. Many 

of the conditions chosen also enable constant-current and 

constant-power trend studies. Figure 12 shows all operating 

conditions tested for the present study for both the NASA300M and the NASA-457Mv2. 

To aid in the creation of the data location matrix, predicted 

near-field plume characteristics were drawn from a Hall2De 

simulation study on the NASA-300M performed by 

Mikellides, et al. (Ref. 19) From prior tests on other Hall 

thrusters of similar design, some ion acceleration was 

expected to take place downstream of the exit plane. Thus, the 

axial spatial resolution of the ion current density data is 

highest in the region nearest the thruster channel exit. The 

spatial resolution decreases as the probe travels axially away 

from the thruster. The spacing is based on the predicted 

plasma density and the assumption that changes in the plasma 

should become more gradual as the plasma density falls. The 

data location matrix for the NASA-457Mv2 was created based 

on geometric scaling from that of the NASA-300M. Figure 13 

shows a diagram of the spatial resolution of the radial sweeps 

for the NASA-300M and NASA-457Mv2 thrusters. 

Figures 14 and 15 present typical NASA-300M ion current 

density profiles and a contour map at 500 V and 20 kW. 

Additional NASA-300M and NASA-457Mv2 ion current 

density profiles and contour maps can be found in 

Reference 21. Total ion current and divergence angles were 

calculated based on the methodology described Reference 21. 

For each operating condition, calculations were carried out for 

each of the 1-e, 2-e, and 3-e integration limits. The reported 

divergence angle is calculated from the radial position of the 

point of origin up to the outer integration limit. Thus, the 

reported angle is strictly an outer-side divergence angle. 

Table 2 shows a summary of the results from carrying out total 

ion current and outer-side divergence angle calculations for 

the NASA-300M and NASA-457Mv2 thrusters. The first five 

columns of the table list basic operating parameters including 

thruster, discharge voltage, discharge current, anode mass 

flow rate, and Faraday probe bias voltage, which is measured 

with respect to the facility ground. The sixth and seventh 

columns list the total ion current integrated within the 1-e 

limits and 2-e limits, respectively. The eighth and ninth 

columns list the outer-side divergence angle calculated within 

the 1-e limits and 2-e limits, respectively. Calculations with 

3-e limits were carried out but are not shown because the 

resulting total ion current varied from 83 to 92 percent of the 

discharge current, these were deemed too large to be real. For 

1-e and 2-e limits, this percentage is 63 to 75 percent and 78 to 

87 percent, respectively. For state-of-the-art Hall thrusters, 

this percentage is typically 70 to 80 percent, (Refs. 24 to 27) 
NASA/TM—2014-218084 9 

Figure 13.—A diagram of the spatial resolution of radial sweeps taken for the NASA-300M and 

NASA-457Mv2. 

Figure 14.—Ion current density profiles for the NASA-300M 

operating at 500 V, 40 A. 

Figure 15.—Flooded contour of ion current density for the 

NASA-300M operating at 500 V, 40 A. 
NASA/TM—2014-218084 10 

TABLE 2.—SUMMARY OF TOTAL ION CURRENT AND OUTER-SIDE DIVERGENCE ANGLE ANALYSIS 

 Vd,

V 

Id, 

A 

Anode mass flow, 

mg/s 

Probe bias,

V 

Total ion current, 

A 

Outer-side divergence angle,

degree 

1-e 

limits 

2-e 

limits 

1-e 

limits 

2-e 

limits 

300M 200 50.3 39.9 –60 37.2 42.7 6.3 12.4 

300 33.3 32.3 –60 22.6 27.6 1.9 4.7 

301 67.3 52.2 –60 45.0 52.8 3.6 9.3 

401 24.9 25.2 –60 17.2 20.9 2.0 4.9 

399 50.0 43.8 –60 33.9 40.9 1.8 4.7 

501 19.9 20.7 –60 14.4 17.1 2.9 5.7 

501 40.6 37.4 –60 28.6 33.8 2.9 5.8 

457Mv2 202 49.1 46.1 –60 31.0 38.3 4.1 7.3 

200 99.8 77.8 –60 71.0 82.1 8.1 14.9 

250 80.2 66.4 –60 54.8 66.4 4.7 9.2 

250 99.8 76.7 –60 66.1 78.7 4.8 10.2 

301 66.4 59.7 –60 44.1 54.3 3.5 6.1 

300 99.8 79.5 –60 72.9 86.4 5.2 9.7 

400 49.7 48.4 –60 33.7 41.8 3.1 5.8 

400 74.8 65.9 –60 51.9 63.1 3.5 5.7 

401 98.2 79.1 –90 68.2 82.4 4.3 6.9 

501 40.5 38.8 –80 28.9 34.7 4.0 6.9 

500 60.1 56.6 –80 41.6 50.9 3.0 5.5 

501 79.7 69.0 –80 56.7 67.9 3.7 6.4 

501 101.1 80.5 –90 73.5 87.8 4.4 7.3 

hi-mag 201 99.6 84.9 –60 74.6 86.4 8.7 15.9 

so the correct choice of integration limit appears to lie 

somewhere between 1-e and 2-e. Since no further information 

is available to help determine the right integration limits, both 

1-e and 2-e values are reported. 

One very notable feature of the outer-side divergence angle 

result is how small they are compared to typical far-field 

measurement results. Reducing far-field ion current density 

data typically yield divergence angles in the range of 15° to 

25° for state-of-the-art Hall thrusters (Refs. 25 and 28). Yet, 

the above near-field measurement show that the discharge 

beam exits the thruster with an average of 4° and 8° 

divergence as calculated with the 1-e and 2-e limits, 

respectively. Lastly, a short study of the NASA-457Mv2 

plume while roughly doubling the magnetic field strength 

reveal that the overall plume shape does not change greatly 

but the cathode plume becomes more distinctly separated from 

the overall plume structure. While the anode mass flow rate 

was increased by about 9 percent to keep the discharge current 

constant, the total ion current increased by approximately 

5 percent when operating with the higher magnetic field strength. 

B.2.Plasma Potential Measurements 

Plasma potential measurements were performed with an inhouse designed and built emissive probe (shown in Fig. 11). 

Details of the probe design and the electrical circuit used 

during data acquisition are provided in References 22 and 23. 

For the NASA-300, plasma potential measurements were 

performed at 20 kW and 500 V operating condition. Figure 16 

shows the contour of the of plasma potential at this operating 

condition. It is evident from the figure that the plasma 

potential is fairly uniform throughout most of the near-field 

plume. Most of the acceleration zone that was captured lies 

within 0.2 DT,m of the thruster exit, while the remainder of the 

plume is at 20 to 30 V above cathode potential. The maximum 

measured plasma potential at 0.1 DT,m from the exit plane is 

approximately 96 V. This confirms that while near-field plume 

studies provide critical information about the plasma just 

downstream of the exit plane, internal measurements are 

necessary to capture the majority of the acceleration zone. 

For the NASA-457Mv2, plasma potential contours are 

shown in Figure 17. It shows that most of the near-field plume 

is at a fairly uniform potential of approximately 20 to 25 V 

above cathode potential for both operating conditions. Most of 

the acceleration zone appears to be contained within 0.4 DT,m

of the thruster exit plane. The maximum measured plasma 

potentials at 0.08 DT,m from the exit plane are 52 V at 300 V 

and 30 kW and 55 V at 400 V and 30 kW. This confirms that 

while such near-field studies provide critical information 

regarding the plasma downstream of the exit plane, the 

majority of the acceleration zone lies upstream of this region 

and requires internal measurements to fully capture it. From 

Figure 17, the plasma potential appears to drop off radially 

from channel centerline more rapidly at 400 V compared to 

300 V. This is an indication that the plasma is more 

concentrated around channel centerline at 400 V, thus 

implying the beam is more collimated. This trend is confirmed 

with ion current density measurements made in the near-field 

at these conditions (Ref. 5). 
NASA/TM—2014-218084 11 

Figure 16.—Contour map of plasma potential in the 

near-field plume at 500 V and 20 kW. Most of the 

acceleration zone captured is within 0.2 DT,m of the 

thruster exit plane. All values reported are with 

respect to cathode potential. 

Figure 17.—Plasma potential contours in the near-field plume of the NASA-457Mv2. Left: 300 V at 30 kW. Right: 400 V 

at 30 kW. The acceleration zone for both conditions appears to be contained within 0.4 DT,m of the thruster exit plane. 

All reported potentials are with respect to cathode potential. 

 
NASA/TM—2014-218084 12 

Detailed discussions of the plasma potential results of the 

NASA-300M and NASA-457Mv2 are provided in 

References 22 and 23, respectively. 

B.3.Electron Temperature Measurements 

Electron temperature measurements were performed with an 

in-house NASA Glenn designed and built single Langmuir 

probe (shown in Fig. 11) that was housed inside a double-bore 

alumina tube. An identical wire was sleeved in the adjacent 

bore and encapsulated by high-temperature Ceramabond. A 

“null” probe was biased in the same manner as the “active” 

probe, but since it was isolated from the plasma, it was used to 

characterize the capacitive currents in the electrical lines that 

can be subtracted out during plasma processing and reduction. 

Data analysis of the Langmuir probe I-V characteristics 

followed the simple Langmuir probe theory (Refs. 29 and 30). 

Details of the probe design, raw data analysis procedures, and 

the electrical circuit used during data acquisition are provided 

in References 22 and 23. 

The near-field NASA-300M electron temperature profiles 

were obtained for various discharge voltage and power 

operating conditions. In Reference 22, results are presented for 

discharge operating conditions of 300 V and 10 kW, 300 V 

and 20 kW, 400 V and 20 kW, and 500 V and 20 kW. Results 

presented in this paper are for operation at a discharge voltage 

of 500 V and 20 kW. Figure 18 left shows that the high 

temperature region is contained within 0.3 DT,m of the thruster 

exit plane. Afterwards, the electron temperature is fairly 

uniform in the near-field at 2 to 5 eV. The maximum 

measured electron temperatures at 0.07 DT,m from the exit 

plane were 27 eV at 500 V and 20 kW. As reported in 

Reference 22, higher electron temperatures are expected at 

higher discharge voltages due to increased ohmic heating. In 

addition, changes in the location of the ionization and 

acceleration zones introduce additional factors that affect the 

maximum electron temperature seen in the near-field plume. 

Detailed discussion of the electron temperature measurements 

at the various operating conditions of the NASA-300M was 

presented in Reference 22, but in general, the results indicate 

that there exists a lobe-like structure that is strongly 

reminiscent of the magnetic field topology in the near-field 

region of the NASA-300M. 

For the NASA-457Mv2, electron temperature profiles were 

obtained for various discharge voltage and power operating 

conditions. In Reference 23, results are presented at discharge 

operating conditions of 300, 400, and 500 V at 30 kW, and 

500 V at 50 kW. Figure 18 (right) shows the electron 

temperature contours at 500 V and 50 kW. For all operating 

conditions, the region of high temperature is contained within 

0.25 DT,m from the thruster exit plane. The remainder of the 

near-field plume is at roughly constant temperature ranging 

from 2 to 5 eV. The maximum measured electron temperatures 

at 0.06 DT,m from the thruster exit plane are 26 eV for 300 V at 

30 kW, 24 eV for 400 V at 30 kW, 27 eV for 500 V at 30 kW, 

and 32 eV for 500 V at 50 kW. The NASA-457Mv2 electron 

temperature contours at 500 V are very similar to results found 

for the NASA-300M. 

Finally, both the NASA-300M and NASA-457Mv2 electron 

temperature contours provided great insights as to the location 

and structure of the acceleration zone, but detailed probing of 

the discharge chamber is required to assess and determine the 

precise location of the acceleration zone and lifetime of the 

thruster. As is highlighted in Section B.5, measurements of the 

floating potential, electron temperature, and number density 

were performed on the NASA-300M thruster. At the present 

time, there are no plans to perform such measurements on the 

NAS-457Mv2, but if that need arises such measurements 

could be performed using the existing setup. 

Figure 18.—Contour maps of electron temperature in the near-field plume of the NASA-300M and NASA-457Mv2. Left: 

NASA-300M at 500 V and 20 kW. Right: NASA-457Mv2 at 500 V and 50 kW. The two thrusters contour profiles exhibit 

similar structure in their isothermal lines.
NASA/TM—2014-218084 13 

B.4.Optical Emission Spectroscopy 

Another goal of the advanced plasma diagnostics 

development under the OCT ISP technology project is to 

demonstrate nonintrusive diagnostics capability that can assess 

thruster life under various operating conditions. The ultimate 

goal is to demonstrate and integrate a nonintrusive diagnostic 

capability that provides a real-time erosion measurement that 

mitigates or eliminates the need for very long duration testing 

of developmental hardware (at least from a life-limiting wear 

perspective). The technique will have the potential to measure 

not only insulator erosion but also cathode surface erosion and 

efflux of cathode emitter material. It may also have the ability 

to characterize the interaction of the thruster plume with 

spacecraft surfaces. The following summarizes the first phase 

in the development of this capability, which now has the 

ability to provide a real-time assessment of the relative 

insulator erosion rate with thruster operating condition. 

The evaluation and demonstration of the optical emission 

spectroscopy (OES) erosion monitoring technique took 

advantage of opportunities to test with different Hall thrusters. 

In addition to providing erosion data on these different 

thrusters, the ability to use the technique with a variety of 

thrusters of significantly different powers was demonstrated. 

One objective of this investigation was to compare the erosion 

trends of the small, albeit High Voltage Hall Accelerator 

(HiVHAc) engineering development unit (EDU) thruster with 

the higher power NASA-300M and NASA-457Mv2 thrusters. 

All thrusters operated on xenon propellant in either vacuum 

facility 8 (VF8) or VF5 at NASA Glenn. Magnetic field 

settings were selected to optimize performance of each 

thruster. 

The use of fiber optic probes to characterize the wear and 

operation of high-power Hall thrusters has been demonstrated. 

Figure 19 shows the testing configuration. The emission of 

boron atoms has been normalized (NB values) using an 

actinometrical method that mitigates the need to know the 

electron temperature in the discharge or the detailed excitation 

and quenching mechanisms of the boron I transitions. The 

technique yielded a real-time, nonintrusive diagnostic that 

captured trends in BN insulator erosion as various thrusters 

were operated. In this regard, it demonstrated the ability to 

provide at least a real-time relative assessment of the impact 

of changes in the operating conditions. Also, the clear 

differences in IB,250 (intensity of a boron transition with 

wavelength of 250 μm) collected from probes focused a few 

millimeters apart on the BN walls demonstrated the ability to 

measure with spatial resolutions on the order of a few 

millimeters within the thruster. 

Gold and silver foil layers were used to demonstrate the 

potential of the OES technique incorporating actinometry. 

While not directly applicable to the life assessment of the 

thruster itself, the technique demonstrated the ability to 

correlate the OES signal with the overall erosion rate of the 

Figure 19.—Superposition of the sampling volume onto the test 

configuration of the NASA-300M. Note that the regions of 

relatively high sampling (red-green regions) are in the plume 

and the regions of lower sampling are largely outside of it. 

target materials. Note that there is qualitatively good 

agreement. Thus, the technique demonstrated the ability to 

measure BN erosion rates if those rates can be correlated with 

IB,250. 

Insulator erosion trends characterized by the normalized 

boron neutral atom spectral emission at 250 nm were obtained 

on the 3.9 kW HiVHAc EDU thruster, the 20 kW NASA300M thruster, and the 50 kW NASA-457Mv2 thruster. The 

NB values were linearly proportional to power and discharge 

voltage for the HiVHAc thruster and showed little difference 

between the inner and outer walls as seen in Figure 20. NB 

values for the NASA-300M thruster showed a greater than 

linear dependence on power, primarily trending with discharge 

current. The inner wall of this thruster had much higher NB 

values than the outer wall. NB values of the NASA-457Mv2 

thruster also showed an unexpected dependence on discharge 

current as illustrated in Figure 21. The NB values varied 

linearly with discharge current with a slope roughly equal to 

(1-ηA). Why a stronger dependence on discharge voltage was 

not observed for the high-power thrusters remains unclear but 

may result from the averaging of the xenon lines by the OES 

probes used to normalize (correct) the raw IB,250. More 

analysis and data are required to confirm these trends or to 

resolve why the OES technique is not capturing the proper 

dependence of the erosion rate on the operating condition. 

Details of the OES technique development can be found in 

Reference 31. 
NASA/TM—2014-218084 14 

Figure 20.—Variation in NB normalized to its maximum value for the HiVHAc EDU thruster 

operating at a discharge current of 5 A. 

Figure 21.—Trends in normalized NB values as a function of axial location for 100 A 

operation of the NASA-457Mv2 thruster. Note that the signal decreases for 

increasing Vd. 

 
NASA/TM—2014-218084 15 

C. Far-Field Plume Measurements 

Far-field plasma measurements were performed within the 

far-field plume of the NASA-300M Hall thruster. Data were 

collected up to 500 V discharge voltage and 20 kW discharge 

power. The far-field measurements incorporated the use of a 

Faraday probe, Langmuir probe, a retarding potential analyzer 

(RPA), and an E×B probe shown on the photograph in 

Figure 22 as mounted in VF5. Spatial mappings of the beam 

current density, as well as single point measurements of the 

ion charge state and acceleration potential were made. 

Measurements were performed to gain insights into the 

various efficiency losses associated with Hall thrusters and 

how they vary with operating conditions. In addition, the farfield measurements provide valuable insights into the 

thruster/spacecraft interactions. The measured plasma 

properties will be used to evaluate how the plume from highpower Hall thrusters interact with the various spacecraft 

surfaces. The far-field data is being analyzed and will be 

presented at the 2013 International Electric Propulsion 

Conference (IEPC 2013). 

D. Internal Discharge Chamber Plasma 

Measurements 

Internal discharge chamber plasma measurements were 

obtained for the NASA-300M. The test provides critical 

reference data for attaining a better understanding of highpower Hall thruster physics, provides data for validating/ 

improving a JPL Hall thruster simulation code, and will help 

guide the design of future high-power, long-life, and highperforming Hall thrusters. Additionally, the newly assembled 

plasma diagnostics setup enhances NASA Glenn’s plasma 

diagnostic capabilities for future EP testing. 

Tests were performed in VF5. The tests were carried via a 

newly developed high-speed axially reciprocating probe 

(HARP) system that can rapidly inject and retract probes from 

the channel of the Hall thruster. Figure 23 shows the setup 

inside VF5. 

For this test series the NASA-300M was operated at five 

different conditions spanning discharge powers of 10, 15, and 

20 kW, and discharge voltages of 300, 400, and 500 V. These 

operating conditions cover the power and specific impulse 

range of interest to OCT missions. Floating potential, electron 

temperature, and number density measurements were 

performed at various radial locations inside the NASA-300M 

discharge chamber. Initial data inspection revealed that the 

data will provide great insights as to the location of the 

ionization and acceleration zones and will be critical in 

validating the JPL Hall2De code. Data analysis of the acquired 

data is ongoing and results will be presented in an upcoming 

paper. The internal, near-field, far-field data, thruster 

performance, and numerical modeling results will greatly 

enhance our understanding of high-power Hall thruster 

physics and lifetime and will contribute to future design of 

high-power and high-fidelity Hall thrusters. 

E. Physics Based Modeling 

To reduce flight risk as well as time and cost during the 

development and flight qualification of high-power Hall 

thrusters, NASA JPL and Glenn are employing physics-based 

numerical simulation, plasma and wear diagnostics, and 

thruster testing by combining unique capabilities at the two 

centers. 

A first series of numerical simulations of a 20-kW class 

Hall thruster have been completed to establish the groundwork 

for the development of high-power, high-performance, and 

long-life Hall thrusters. The numerical simulations have been 

performed with Hall2De, a two-dimensional axisymmetric, 

magnetic-field-aligned computational mesh (MFAM) plasma 

solver developed at NASA JPL (Refs. 32 and 33). Detailed 

results from the computational work were presented in 

Reference 19. Simulation results provided a first assessment of 

the erosion rates in the design of the NASA-300M. Figure 24 

shows the two-dimensional contours of the number density in 

the thruster’s acceleration channel. The simulation results 

indicated that the ionization zone is located far upstream of the 

LP 

RPA 1 

RPA 2 

E×B 

Figure 22.—Photograph of the far-field probes. 
NASA/TM—2014-218084 16 

Figure 23.—HARP stage setup inside VF5 with NASA-300M 20 kW thruster. 

Figure 24.—Two-dimensional contours of the number density 

in the NASA-300M discharge channel. 

channel exit and that significant ion focusing is achieved due 

largely to the applied magnetic field topology. This 

combination leads to erosion rates along the channel insulators 

that do not exceed 1 mm/kh. Recent findings on erosion 

physics and related wear mitigation techniques suggest it may 

be possible to modify the magnetic circuit and channel 

geometry of this thruster to achieve magnetic shielding, a 

technique that seeks to reduce erosion rates at the channel 

walls by more than one order of magnitude. Reduced erosion 

rates are attained by modifying the thruster’s magnetic circuit 

to reduce ion bombardment on the discharge channel walls 

(Refs. 32 and 33). Recent simulations have been performed on 

a magnetically shielded version of the NASA-300M thruster. 

Extended simulations that include assessment of thermal loads 

to the thruster surfaces, plasma diagnostics and model 

validation, thruster modifications, and testing constitute the 

focus of this effort in the near future. 

F. High-Current Hollow Cathode Development 

As part of the OCT ISP technology project, the NASA 

Glenn is developing and testing new high-current hollow 

cathode assemblies that can meet and exceed the required 

discharge current and life-time requirements (20,000 hr) of 

high-power Hall thrusters test results of three high-current 

hollow cathode configurations were reported in Reference 34. 

Figure 25 shows configurations 1 and 2. Test results indicated 

that both configurations were able to attain lower peak emitter 

temperatures compared to state-of-the-art hollow cathode 

assemblies. Figure 26 compares the cathode orifice plate tip 

temperatures for configurations 1 and 2 at 

50 and 100 A. Test results at 100 A and a xenon flow rate of 

37 sccm indicate that configuration 2 cathode orifice plate tip 

temperature is 70 °C lower than that for configuration 1. Also, 

comparing configurations 1 and 2 cathode orifice plate tip 

temperatures at 100 A to test results of previously tested highcurrent cathode configurations at NASA Glenn shows that 

configurations 1 and 2 cathode orifice plate tip temperatures 

were 200 to 270 C lower than what was measured by 

Carpenter et al. in 2001 (Ref. 35) and were approximately 100 

to 170 C lower than what was measured by John et al. in 2005 

(Ref. 36). Configuration 2 cathode orifice plate tip temperature of 

1132 C at 100 A is consistent with the temperature magnitude 

that is required for emitter life exceeding 20,000 hr. A 

preliminary analysis of configuration 2 test results indicated that a 

uniform emitter temperature was realized along the emitter 

surface. This results in uniform impregnate deletion rates across 

the emitter and long life (Ref. 34). 

 
NASA/TM—2014-218084 17 

Figure 25.—Photograph of high-current hollow assemblies 

configurations 1 and 2. 

Figure 26.—Configurations 1 and 2 cathode orifice plate tip 

temperature variation with flow rate for discharge currents 

of 50 and 100 A. 

Near-term tests will include operating configuration 2 for 

extended durations. In addition, detailed emitter temperature 

and plasma measurements inside the hollow cathode emitter 

region have to be performed. These measurements will be 

supplemented by extensive physics-based hollow cathode 

modeling. 

G. Thermal and Structural Modeling of HighPower Hall Thrusters 

Thermal modeling of the NASA-300M and NASA-457Mv2 

thrusters was performed to gain insight into their thermal 

characteristics and to improve future designs of flight 

thrusters. The models were created using commercially 

available software. Simulations were performed for the 

thruster operation under deep space cold conditions and under 

worst-case hot conditions—in low-Earth-orbit with solar and 

Earth albedo heat fluxes while firing. In addition, simulations 

were also performed for a ground test environment to evaluate 

the model results against experimentally measured values. 

Both the NASA-300M and NASA-457Mv2 were instrumented 

with thermocouples during performance and plume testing to 

provide experimental data for model correlation. Figures 27 

and 28 plot the experimental temperature data with the thermal 

modeling results for the NASA-300M and NASA-457Mv2 

thrusters, respectively. Since the data were not obtained from 

dedicated thermal characterization tests, but simply the 

temperatures collected during performance and plume testing, 

thermal steady state was not achieved for all of the test data. 

Curve fits to the semi-empirical estimates of the steady state 

temperatures based on the experimental data are also shown in 

Figures 27 and 28. Simulation results indicate a fair 

comparison to the steady state temperature curves. 

Structural modeling of the NASA-457Mv2 Hall thruster 

was also performed with commercially available software to 

evaluate the response of the thruster under launch acceleration 

and vibration loads. The calculated stresses of the thruster 

components were well within material margins. A preliminary 

assessment of thermally-induced stress was also performed 

with inputs from the thermal modeling results. The primary 

thruster modes were also evaluated. The NASA-457Mv2 

thruster, designed to incorporate structural improvements over 

the previous version, will undergo vibration testing at NASA 

Glenn to provide data to correlate to the modeling efforts. The 

lessons learned from the thermal and structural analyses and 

testing will aid in the understanding and improvement of 

existing thruster designs and also provide guidelines for future 

thruster development. 

H. High-Power Discharge Module Development 

for High-Power Hall Thrusters 

The development of a high-power SEP system for HEOMD 

missions will require the development of power electronics. 

As power level increases, it becomes more difficult to develop 

the power converters because of component availability, the 

increase in current stress, and conduction losses. The increase 

in power level impacts not only the EP system but also the 

power management and distribution (PMAD) systems and the 

solar arrays. For future high-power spacecraft, it may be 

advantageous to implement high-voltage power systems to 

mitigate these effects. However, one issue with this 

implementation is that space qualified, radiation-hardened, 

and high-voltage semiconductors are very scarce. Since part 

qualification will likely be necessary for a power processing 

unit (PPU) with high power, consideration of emerging 

semiconductor technologies like silicon carbide (SiC) and 

gallium nitride (GaN) should be made. Recently, practical 
NASA/TM—2014-218084 18 

Figure 27.—Test temperatures, including steady state curve 

fits, and model results for the NASA-300M Hall thruster. 

Figure 28.—Test temperatures, including steady state curve 

fits, and model results for the NASA-457Mv2 Hall thruster. 

commercial transistors became available for both of these 

technologies. SiC is especially appealing for high-power 

applications because the available devices are rated for high 

voltage (1200 VDC) and should also have good immunity to 

radiation effects. 

A comprehensive trade study was conducted to look into a 

variety of semiconductor technologies and transistor devices 

Figure 29.—Breadboard discharge module. 

for development of a high-power PPU. Models to estimate 

performance and power losses as a function of converter 

topology, voltage, current, and frequency were generated. The 

models were validated with data from component testing. SiC 

metal oxide semiconductor field effect transistors (MOSFETs) 

were selected as the best candidates for high-power PPUs with 

input voltages higher than 200 V. SiC Schottky diodes were 

selected for high-voltage, high-current rectification. 

The effort then focused on designing and building a highpower module to process power for a Hall thruster main 

discharge. A breadboard module is shown in Figure 29. An 

input voltage of 300 V 10 percent was chosen for this design 

anticipating it would operate of a high-voltage power bus. An 

output voltage range of 300 to 400 V was chosen because it 

will result in a specific impulse that would satisfy most 

missions of interest. The module was designed to process as 

much power as possible while meeting the derating 

requirements for the power semiconductors. A full bridge 

topology was chosen because it has low voltage stress on the 

power transistors and good transformer utilization making it a 

good choice for high-power applications. A switching 

frequency of 20 kHz was used to maintain switching losses as 

low as possible and to maximize efficiency. At full power the 

module processes as much as 7.5 kW. A new circuit was 

developed for this design to provide gate drives with the high 

voltage and speed required by the SiC MOSFETs to operate 

with very low switching losses. Preliminary measurements 

revealed efficiencies around 97 percent. 
NASA/TM—2014-218084 19 

Future work will focus on refining and optimizing the 

module design so that it has full functionality of a flight unit. 

A vacuum compatible brassboard unit will be designed and 

built to start addressing the thermal and packaging issues 

while using a highly modular slice architecture. 

IV. Summary 

NASA’s Human Exploration Framework Team identified 

high-power electric propulsion (EP) as an enabling technology 

for NASA human exploration missions. NASA’s Office of the 

Chief Technologist (OCT) Game Changing Division (GCD) is 

funding a solar EP technical demonstration mission and 

technology development projects that are developing and 

maturing and will eventually demonstrate high-power Hall 

thruster technologies in space. 

Under NASA’s OCT GCD In-Space Propulsion technology 

program, NASA Glenn has been engaged in developing and 

testing of high-power Hall thrusters. Performance evaluations 

of the NASA-300M (20 kW) and NASA-457Mv2 (50 kW) 

thrusters indicated that both thrusters achieved peak total 

thrust efficiencies above 65 percent. Detailed plasma 

measurements performed in the near-field of the NASA-300M 

and NASA-457Mv2 thrusters, provided great insights as to the 

location of the ionization and acceleration zones and indicated 

close correlation between the plume structure and the 

magnetic field topology. Internal discharge chamber and farfield mappings were also performed on the NASA-300M and 

analysis of acquired data is ongoing. The internal discharge 

chamber measurements will further help us evaluate and 

assess thruster lifetime and can be used to further anchor the 

NASA Jet Propulsion Laboratory (JPL) physics-based Hall 

thruster model. The far-field data will be used to assess 

thruster efficiency loss mechanisms and thruster/spacecraft 

interactions. Thermal modeling of the NASA-300M and 

NASA-457Mv2 thrusters was also performed and compared 

well with measured temperatures. Detailed structural modeling 

of the NASA-457Mv2 was also performed. Model results will 

be compared to upcoming measurements from a random 

vibration test of the NASA-457Mv2. All the above mentioned 

tests and modeling efforts have greatly improved our 

understanding of high-power Hall thruster physics and will be 

leveraged in the design of high-efficiency long-life Hall 

thrusters for NASA missions. 

In addition to testing and developing high-power Hall 

thrusters, the NASA Glenn is also developing a high-current 

hollow cathode for integration with high-power Hall thrusters. 

A hollow cathode assembly was operated at 100 A and at 

emitter temperatures that are consistent with long life. Finally, 

the NASA Glenn is also developing high-power, highperformance discharge modules for power processing units 

that will be used to power the high-power Hall thrusters. 

References 

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5. Brophy, J.R., et al., “300-kW Solar Electric Propulsion 

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8. Mason, L.S., et al., “1000 Hours Testing of a 10 KW Hall 

Effect Thruster,” AIAA Paper 2001-3773, July 2001. 

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2002,” AIAA Paper 2002-3675, July 2002. 

10. Manzella, D.H., et al., “Laboratory Model 50 kW Hall 

Thruster,” AIAA Paper 2002-3676, July 2002. 

11. Jacobson D.T., et al., “NASA’s 2004 Hall Thruster 

Program,” AIAA Paper 2004-3600, July 2004. 

12. Peterson, P.Y., “Performance and Wear Characterization 

of a High Power High–Isp Hall Thruster,” AIAA Paper 

2005-4243, July 2005. 

13. Kamhawi, H., Haag, T.W., Jacobson, D.T., and Manzella, 

D.H., “Performance Evaluation of the NASA-300M 

20 kW Hall Effect Thruster,” AIAA Paper 2011-5521, 

July 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, October 

2001. 

15. Hofer, R.R., and Gallimore, A.D., “The Role of Magnetic 

Field Topography in Improving the Performance of HighVoltage Hall Thrusters,” AIAA Paper 2002-4111, July 

2002. 

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Ph.D. Dissertation, Department of Mechanical 

Engineering, Stanford University, Stanford, CA, June 

2005. 

17. 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. 

 
NASA/TM—2014-218084 20 

18. Soulas, G.C., Haag, T.W., Herman, D.A., Huang, W., 

Kamhawi, H. et al., “Performance Test Results of the 

NASA-457M v2 Hall Thruster,” AIAA Paper 2012-3940, 

July 2012. 

19. Mikellides, I.G., Katz, I., Kamhawi, H., and Vannoord, 

J.L., “Numerical Simulations of a 20-kW Class Hall 

Thruster Using the Magnetic-Field-Aligned-Mesh Code 

Hall2De,” IEPC Paper 2011-244, September 2011. 

20. Jameson, K.K., Goebel, D. M., Hofer, R.R., and Watkins, 

R.M., “Cathode Couplings in Hall Thrusters,” IEPC Paper 

2007-278, September 2007. 

21. Huang, W., Shastry, R., Herman, D.A., Soulas, G.C., and 

Kamhawi, H., “Ion Current Density Study of the NASA300M and NASA-457M v2 Hall Thrusters,” AIAA Paper 

2012-3870, July, 2012. 

22. Herman, D.A., Shastry, R., Huang, W., Soulas, G.C., and 

Kamhawi, H., “Plasma Potential and Langmuir Probe 

Measurements in the Near-field Plume of the NASA 

300M Hall Thruster,” AIAA Paper 2012-4115, July 2012. 

23. Shastry, R., Huang, W., Herman, D.A., Soulas, G.C., and 

Kamhawi, H., “Plasma Potential and Langmuir Probe 

Measurements in the Near-field Plume of the NASA 

457Mv2 Hall Thruster,” AIAA Paper 2012-4196, July 

2012. 

24. Reid, B.M., “The Influence of Neutral Flow Rate in the 

Operation of Hall Thrusters,” Ph.D. Dissertation, 

Aerospace Engineering, University of Michigan, Ann 

Arbor, MI, 2008. 

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Hall Thruster Characteristics and Evaluation of Loss 

Mechanisms,” Ph.D. Dissertation, Aerospace 

Engineering, University of Michigan, Ann Arbor, MI, 

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determining the performance of stationary plasma 

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5, 1998, pp. 736-743. 

27. Hofer, R.R., “Development and Characterization of HighEfficiency, High-Specific Impulse Xenon Hall Thrusters,” 

Ph.D. Dissertation, Aerospace Engineering, University of 

Michigan, Ann Arbor, MI, 2004. 

28. Huang, W., Kamhawi, H., and Shastry, R., “Farfield Ion 

Current Density Measurements before and after the 

NASA HiVHAc EDU2 Vibration Test,” AIAA Paper 

2012-4195, July 2012. 

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Parameters and Chemistry, edited by Flamm, D.L. 

(Academic Press, Inc., 1989), pp. 113-181. 

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“Advanced Diagnostic Characterization of High-Power 

Hall Thruster Wear and Operation,” AIAA Paper 2012-

4036, July 2012. 

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Channel Walls in a Hall Plasma Accelerator,” Phys. 

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Thruster With Magnetically Shielded Channel Walls, 

Phase I: Numerical Simulations,” AIAA Paper 2011-

5809. 

34. Kamhawi, H., and Van Noord, J., “Development and 

Testing of High Current Hollow Cathodes for High Power 

Hall Thrusters,” AIAA Paper 2012-4080, July 2012. 

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Hollow Cathode Development,” IEPC Paper 01-274, 

October, 2001. 

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Current Cathode Development for 50 kW Class Hall 

Thruster,” AIAA Paper 2005-4244, July 2005. 

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High-Power Hall Propulsion Development at NASA Glenn Research Center

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

The NASA Office of the Chief Technologist Game Changing Division is sponsoring the development and testing of enabling technologies to

achieve efficient and reliable human space exploration. High-power solar electric propulsion has been proposed by NASA’s Human

Exploration Framework Team as an option to achieve these ambitious missions to near Earth objects. NASA Glenn Research Center is

leading the development of mission concepts for a solar electric propulsion Technical Demonstration Mission. The mission concepts are

highlighted in this paper but are detailed in a companion paper. There are also multiple projects that are developing technologies to support a

demonstration mission and are also extensible to NASA’s goals of human space exploration. Specifically, the In-Space Propulsion

technology development project at the NASA Glenn has a number of tasks related to high-power Hall thrusters including performance

evaluation of existing Hall thrusters; performing detailed internal discharge chamber, near-field, and far-field plasma measurements;

performing detailed physics-based modeling with the NASA Jet Propulsion Laboratory’s Hall2De code; performing thermal and structural

modeling; and developing high-power efficient discharge modules for power processing. This paper summarizes the various technology

development tasks and progress made to date.

15. SUBJECT TERMS

Hall thruster; Stationary plasma thruster; Hollow cathode; Plasma diagnostics; Channel erosion

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