Michael R. LaPointe and Pavlos G. Mikellides

Ohio Aerospace Institute, Brook Park, Ohio

High Power MPD Thruster Development

at the NASA Glenn Research Center

NASA/CR—2001-211114

August 2001

AIAA–2001–3499
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National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

37th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Salt Lake City, Utah, July 8–11, 2001

Michael R. LaPointe and Pavlos G. Mikellides

Ohio Aerospace Institute, Brook Park, Ohio

High Power MPD Thruster Development

at the NASA Glenn Research Center

NASA/CR—2001-211114 AIAA–2001–3499

Prepared under Cooperative Agreement NCC3–860
Acknowledgments

This research is supported through the Ohio Aerospace Institute under NASA Cooperative Agreement NCC3–860.

The authors express their gratitude to Gene Strzempkowski, Joseph Dick, and George Rodriguez for the design

and construction of the MW-class capacitor bank, pulse-forming network, and related control systems; to

Jim Coy for the modified design and construction of the pulsed propellant feed system; to Stan Grisnik

and Robert Roman for refurbishing the VF–1 pumping and control systems, and to Nicholas Iosue,

Derrick Patterson, Douglas Warren, Roberto DuBreuil, and their Akima Corporation colleagues

for the construction of the thrust stand and related test hardware. The successful construction

and operation of the high power MPD thruster test facility is a reflection of

their skill and dedication.

Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

This report is a formal draft or working

paper, intended to solicit comments and

ideas from a technical peer group.

This report contains preliminary

findings, subject to revision as

analysis proceeds.

Available electronically at http://gltrs.grc.nasa.gov/GLTRS
NASA/CR—2001-211114 1

Figure 1. MPD thruster diagram. (Courtesy J. MacNeill)

HIGH POWER MPD THRUSTER DEVELOPMENT

AT THE NASA GLENN RESEARCH CENTER

Michael R. LaPointe and Pavlos G. Mikellides 

Ohio Aerospace Institute

Brook Park, Ohio 44142

ABSTRACT

Propulsion requirements for large platform orbit raising,

cargo and piloted planetary missions, and robotic deep

space exploration have rekindled interest in the

development and deployment of high power

electromagnetic thrusters. Magnetoplasmadynamic

(MPD) thrusters can effectively process megawatts of

power over a broad range of specific impulse values to

meet these diverse in-space propulsion requirements. As

NASA's lead center for electric propulsion, the Glenn

Research Center has established an MW-class pulsed

thruster test facility and is refurbishing a high-power

steady-state facility to build and test efficient gas-fed

MPD thrusters. A complimentary numerical modeling

and thruster design effort based on the robust MACH2

code provides a well-balanced program of numerical

analysis and experimental validation leading to improved

high power MPD thruster performance. This paper

reviews the current and planned experimental facilities

and numerical modeling capabilities at the Glenn

Research Center, and outlines program plans for the

development of new, efficient high power MPD thrusters.

INTRODUCTION

High power electric propulsion offers significant

advantages over chemical systems for platform orbitraising, lunar and planetary cargo missions, and robotic

and piloted planetary and deep space exploration

missions.1,2,3,4 As the lead NASA center for electric

propulsion, the Glenn Research Center (GRC) is

developing high power electromagnetic thruster

technologies to meet these demanding propulsion

requirements. This paper presents an overview of the

GRC program to design and test new, efficient high

power (0.1-MW to 10-MW) magnetoplasmadynamic

(MPD) thruster technologies.

In its most basic form, the MPD thruster consists of a

central cathode surrounded by a concentric cylindrical

anode (Fig.1). A high-current arc is struck between the

anode and cathode, which ionizes and accelerates a gas

propellant. In self-field thrusters, an azimuthal magnetic

field produced by the return current flowing through the

cathode interacts with the radial discharge current density

flowing through the plasma to produce an axial body

force. In applied-field versions of the thruster, a solenoid

magnet surrounding the anode is used to provide

additional radial and axial magnetic fields that can help

stabilize and accelerate the plasma discharge.

The MPD thruster has had a long if intermittent history of

development, and several review papers have been

written over the past several years that chronicle both the

progress and challenge of this promising propulsion
NASA/CR—2001-211114 2

Figure 2. Diagram of capacitor bank and pulse forming network.

 technology.5,6,7 MPD thrusters have been operated with

a variety of propellants, with the most efficient thruster

performance to date obtained using vaporized lithium.

Lithium-fed MPD thrusters developed by the Moscow

Aviation Institute in Russia recently demonstrated 45%

efficiency at power levels exceeding 100-kW,8

 and

lithium-fed devices are currently being developed in the

United States by the NASA Jet Propulsion Laboratory

and Princeton University.9,10 A potential issue with

lithium is that it is a condensable propellant, and has the

potential to coat spacecraft surfaces and instruments.

Alternate non-condensable gas propellants have been

investigated, but the performance has typically been less

efficient than lithium propellant.6,7,11 However, recent

theoretical models indicate that significantly improved

performance can be obtained with gas-fed MPD thrusters

through the use of modified electrode geometries and the

proper application of applied magnetic fields.12,13 The use

of gas propellants mitigates potential handling and

spacecraft contamination issues associated with lithium,

and provides for thruster operation with various

propellants over a wide range of specific impulse values.

The remainder of this paper describes the theoretical and

experimental program currently underway at the NASA

Glenn Research Center to develop and test gas-fed MPD

thrusters. The GRC MW-class pulsed thruster facility is

described in the next section, followed by an overview of

the planned MW-class steady-state thruster test facility

anticipated to be online in FY03. Following the

experimental program description is an overview of the

MACH2 code numerical modeling effort, and recent

insights into MPD thruster operation provided through

MACH2 simulations. The paper concludes with a brief

summary of planned Glenn Research Center activities to

develop efficient, high power MPD thrusters for near and

mid-term mission applications.

EXPERIMENTAL FACILITIES

Prior MPD thruster experiments at NASA GRC were

discontinued in the early 1990s, coinciding with the

termination of the Space Exploration Initiative and a

dramatic decrease in program funds for high power

electric propulsion system development. Power and

control systems for a 600-kW steady-state thruster test

facility were mothballed and eventually removed and

reused in other programs. With the advent of renewed

interest in high power thrusters, GRC embarked on the

development of a pulsed test facility to evaluate

MW-class MPD thruster designs. This facility is now

online, and is described in detail in the sections below.

Plans to bring the previous steady-state MPD thruster test

facility back online, with an increase in power to 1-MW,

are also described.

Pulsed Thruster Facility

Planning for the GRC MW-class pulsed thruster test

facility was initiated in FY99, with the facility becoming

fully operational in FY01. A high-energy capacitor bank

and pulse-forming network were designed and built, the

vacuum test facility was refurbished, and an

impulse-based thrust stand was fabricated for pulsed

MPD thruster operation.

Capacitor Bank. The capacitor bank consists of

46 capacitors and 7 inductors arranged in a 7-element

Guillemin pulse-forming network (PFN; Fig 2). Each

capacitor is rated for 10-kVDC, with an expected shot life
NASA/CR—2001-211114 3

I 

V 

Figure 3. Voltage and current waveforms for 13-m


resistive load. Bank charged to 3-kV. Terminal voltage

= 54V, current = 8400A. 500μs/division

Figure 4. Partial view of capacitor bank. Inductors

mounted along top rack. Solid-state switch is not shown.

Figure 5. VF-1 vacuum chamber and propellant plenum

test cart.

of 104

 discharges at maximum voltage. The total bank

capacitance is 4.88-mF, providing a maximum stored

energy of 250-kJ at full charge. The discharge period is

approximately 2x10-3 s, with less than 10% current ripple

over the discharge plateau. Switching is performed by a

solid-state thyristor rated for 15-kV and 50-kA peak

current, although the bank can also be fired using an

available triggered spark-gap switch. The inductors

consist of insulated 3/8" copper tubing wound around

short sections of 10.75" (27.3 cm) OD pvc pipe. A 0.12


matching resistor, rated for 10-kV and 250-kJ, is located

in series between the switch and load. Representative

current and voltage traces obtained with a 0.13-


resistive load used to simulate a typical MPD thruster

discharge are presented in Figure 3. A 1000-
, 250-kJ

resistor tied across the bank is used to discharge the

capacitors in the event of a thruster malfunction or safety

interlock violation. The approximate physical dimensions

of the capacitor bank are 4.5 m (15 ft) long by 2.1 m (7 ft)

high by 0.9 m (3 ft) wide. A portion of the capacitor bank

and PFN are shown in Figure 4. 

Vacuum Chamber. Pulsed MPD thruster testing

is conducted in GRC Vacuum Facility 1 (VF1). The

vacuum chamber, shown in Figure 5, is 1.5-m (5 ft) in

diameter by 4.5-m (15 ft) long. Access to the chamber

interior is provided through a flat endcap mounted on a

balanced swing-arm. The tank is pumped by two 0.8-m

(32") oil diffusion pumps backed by a mechanical blower,

and is capable of reaching an unloaded base pressure of

10-6 torr (8x10-9 Pa). The vacuum tank is located on the

floor immediately above the capacitor bank, and power is

cabled to the tank through high voltage feedthroughs

located on a 1-m (36") bottom flange. Instrumentation

and gas feedthroughs are located on a 0.4-m (16") side

port, with additional access available through several

0.1-m (4") ports mounted to the facility endcap. A

separate 0.4-m side port provides power cable access for

a magnetic field coil, described in a later section of the

paper.

Propellant Plenum. The propellant gas plenum

is mounted to a test cart located beside VF1 (Fig. 5). The

cylindrical plenum chamber is constructed of stainless

steel, with an interior volume of 2.5x10-2 m3

 (1.5x103 in3).

Stainless steel tubing connects one end of the plenum to

a regulated gas bottle, while the other end is connected to

a pneumatic actuator powered by a fast acting solenoid

valve. Gas flow from the actuator can be directed either

through a side flange into the VF-1 test chamber or to a

small calibration chamber located on the test cart. The
NASA/CR—2001-211114 4

Figure 6. Applied-field quasi-steady MPD thruster

mounted to flexure-based thrust stand. 

cylindrical calibration chamber, also constructed of

stainless steel, has an internal volume of 6.1x10-3 m3

(3.7x102 in3) and is used to measure the amount of gas

discharged during pulsed plenum operation. Both the

propellant plenum and the calibration chamber are

outfitted with wide-ranging pressure gauges (dual pirani

gauge and diaphragm manometer), capable of measuring

pressure values from 10-3 torr to 1500 torr. During

calibration, the propellant plenum pressure is set to a

given value, and the calibration chamber is evacuated to

vacuum. Gas from the propellant plenum is discharged

into the calibration plenum, and the resulting pressure rise

is recorded. Knowing the pressure change and the

chamber volume allows the propellant mass injected into

the calibration chamber to be calculated, and dividing this

mass by the total response period of the pneumatic

actuator provides the propellant flow rate for the pulsed

discharge. The response time of the actuator can be set

between 60-ms and 180-ms through the use of a variable

resistor in the control circuit; the latter value has been

used to reduce the effect of uncertainty in the actuator

opening and closing response times. The calibrated mass

flow rates are estimated to have maximum uncertainties

of ± 5%.

Applied-Field Magnet. Prior research has

shown that it is often advantageous to operate MPD

thrusters with applied magnetic fields.6,7,12,13 To provide

this capability, 4-AWG insulated cabling was wound on

a 0.2-m (8") OD pvc pipe to form a solenoid test magnet.

The magnet consists of 7 layers with 18 turns per layer.

The magnet cabling is uncooled, limiting both the

maximum current and period of operation, but the

simplicity of operation and the attainable field strength

over the required thruster pulse period outweigh these

disadvantages. Power is supplied by a 50-kW

constant-current arc welder, capable of delivering up to

1000 A to the magnet. The axial magnetic field strength

at the center of coil as a function of current is

approximately 5x10-4 T/A, while at the end of the coil the

axial field is along the centerline is approximately

3.3x10-4 T/A.

Control System. The events leading up to a

pulsed thruster discharge are controlled by an automated

sequencing system mounted in the MPD control rack. If

the applied-field magnet is to be used, the arc welder is

set to the desired current level. The capacitor bank is then

charged to the desired voltage using a 4-kW, 10-kV

power supply located in the control rack. Once the bank

is charged to the desired voltage, the charging supply is

disengaged and the test sequence is initiated. The magnet

power is turned on for a period of 4.5 s, sufficient to

establish a constant field distribution within the thruster

and allow any residual thrust stand disturbances due to

turning on the magnet to damp away. Toward the end of

this period, the propellant plenum solenoid is activated to

provide a gas pulse to the thruster; as noted, the duration

of the gas pulse can be set from 60-ms to 180-ms, with

the present value set at 180-ms to reduce uncertainties in

the pulsed propellant mass flow rate. The capacitor bank

switch is triggered toward the end of the gas flow period,

providing a 2-ms, high current discharge to the thruster.

Once the bank has fired, the control sequence turns off

the gas flow and magnet current, and the logic system is

then reset to initiate another test sequence.

Thrust Stand. The impulse thrust stand is based

on a simple flexure design (Fig. 6). The thruster is

mounted to a stiff plate, which in turn is supported by

four thin, contoured aluminum flexures. Current is fed to
NASA/CR—2001-211114 5

the thruster through flexible braided copper cables

mounted on each side of the stand. Upon discharge, the

thrust stand is horizontally displaced by the thruster

impulse. An LVDT transducer is used to measure thrust

stand displacement for the larger oscillations occurring at

high discharge powers, while strain gages mounted to the

flexures are used to measure smaller displacements. The

thrust stand is calibrated using a fixed-length pendulum

that provides non-elastic momentum transfer to the stand.

Varying the angle at which the pendulum is released

varies the impulse delivered to the stand, allowing

displacement to be measured as a function of impulse.

The present thrust stand does not allow in-situ calibration

under vacuum conditions; however, a solid-state

inclinometer mounted beneath the top plate of the thrust

stand is used to provide information on the initial and

final orientation (tilt) of the stand. No significant

differences have been observed between thrust stand

orientations at rest in air or under vacuum, providing

some assurance that the measured impulse values will

closely correspond to the calibrated thrust stand values.

Nevertheless, the angular orientation of the stand before

and after each shot is recorded to allow refined

calculations of the delivered impulse, if required.

Data Collection. The primary diagnostics at this

stage of facility development include thruster voltage and

current, thrust stand displacement, and magnet coil

temperature. Measurements of the thruster voltage and

current allow the thruster power to be calculated;

combined with thrust measurements supplied by the thrust

stand displacement, the efficiency of the thruster can then

be determined. These basic measurements serve the main

purpose of the pulsed thruster facility: to develop an

efficient MPD thruster design prior to undertaking more

extensive steady-state development and testing.

Digital oscilloscopes are used to measure and record

voltage and current waveforms at the capacitor bank and

at the thruster or resistive load. Differential voltages are

read across the thruster electrodes and across the

solid-state switch. Voltage attenuation (resistor network)

circuits are used to scale the voltage channels to prevent

waveform clipping by the oscilloscopes. Current

monitors (transformers) are used to measure the pulsed

current at the output of the capacitor bank and at the input

to the thruster, providing independent measurements of

the discharge current. The coils are rated for 50-kA peak

current, a current-time product of 65 A-s, and a rise time

of 2.5x10-7 s. Voltage and current waveforms are

transferred from the oscilloscope to a stand-alone desktop

computer for later analysis. 

As previously noted, the impulsive force delivered to the

thrust stand during a thruster discharge is measured by a

LVDT transducer at high impulse values and strain-gages

mounted to flexure supports at low impulse values.

Transducer or strain-gage voltage signals are sent through

a 9-pin vacuum feedthrough to a desktop computer data

acquisition system, where the data is recorded for later

analysis. The thruster temperature is not expected to rise

significantly during each discharge, and the period

between discharges is sufficiently long that thruster

electrode temperature measurements are not required.

However, the magnet coil current will resistively heat the

applied-field magnet, and thermocouples have been

embedded in the coil to monitor the temperature rise and

signal when testing should be suspended to cool the

magnet. Hall and Langmuir probe plasma diagnostics

will be installed in the pulsed thruster facility at the

beginning of the next fiscal year.

Steady-State Thruster Facility

To evaluate steady-state thruster performance, as well as

investigate a number of related material and life issues, a

MW-class steady-state MPD thruster test facility is

currently being refurbished at the NASA Glenn Research

Center. Prior steady-state, 100-kW class MPD thruster

experiments were performed at the Glenn Research

Center during the late 1980’s through the early 1990’s,

when MPD research was terminated due to NASA

budgetary constraints. Results of that effort have been

described elsewhere, and the reader is referred to the

original papers and the references therein.7,14,15,16,17,18

Several major pieces of test equipment remain from the

earlier program, and the primary power and control

systems are currently undergoing repair or replacement to

provide up to 1-MW of uninterrupted power to the test

facility. The result of this multi-year upgrade will be a

world-class test chamber capable of supporting steadystate MW-class MPD thruster experiments. The following

sections provide a brief overview of the major

components of this facility, which is slated to begin high

power MPD thruster operation in 2003.

Vacuum Chamber. The vacuum chamber,

designated VF-5, is a 4.6-m (15 ft) diameter by 18.3-m

(60 ft) facility pumped by twenty 0.8-m (32 in) diffusion

pumps, which in turn are backed by four root blowers and

four mechanical pumps. The facility is capable of

attaining an unloaded base pressure of 4x10-7 torr (3x10-9

Pa), and can pump 2.5x104 l/s of air while maintaining a

chamber pressure of 10-5 torr. In addition to the diffusion

pumps, the facility is outfitted with a 33.5-m2

 helium
NASA/CR—2001-211114 6

cryopanel that increases the total pumping capability to

1.5x106

 l/s of air at a chamber pressure at 10-6 torr. The

vacuum facility is fully automated, and provides a number

of gated test ports and instrumentation feedthroughs to

support a wide variety of experimental requirements. 

Infrastructure. Although most of the MPD

thruster control systems were removed in the early 1990s,

some power supplies and water-cooling lines remain

intact. Additional power capability will be added to the

facility, bringing the total available power to 1-MW.

Water-cooling pumps will be reinstalled to provide heat

rejection for steady state, high power thruster operation.

The thruster control system will be replaced, as most of

the original components were cannibalized for use in

other electric propulsion test programs. An existing thrust

stand, based on an inverted pendulum design14 and used

in the prior MPD thruster test program, will be

refurbished and modified for MW-class operation. A

calibrated propellant mass flow system will be retrieved

from storage and refurbished, along with several

equipment racks and control consoles. Instrumentation,

such as terminal voltage and current monitors, will need

to be replaced, as will more advanced diagnostic systems

such as motion-controlled Langmuir and Hall probes.

Fortunately the facility and control schematics from the

previous MPD thruster test program were well

documented, and will allow a fairly smooth transition to

higher power operation.

Pulsed vs. Steady State Operation

The pulsed thruster test facility provides an economical

testbed for the evaluation of MW-class MPD thruster

designs prior to more extensive testing under steady-state

conditions. Novel thruster designs can be quickly

constructed and tested to gauge their potential

performance during high power operation, with only the

most promising thruster concepts promoted for further

steady-state development and testing. Although pulsed

operation does not capture all of the performance

variables encountered in steady-state thruster operation,

the pulsed high power tests are sufficient to determine

viable designs for efficient MW-class operation. In

support of this approach, MPD thruster experiments

performed at the University of Stuttgart demonstrated that

pulsed thrusters operated at up to several hundred

kilowatts provided lower thrust and higher voltage than

identical thrusters operated at comparable steady-state

power levels, although the discrepancies became less

significant as the discharge power was increased.19,20

These results indicate that quasi-steady MPD thruster

performance will probably be slightly worse than steadystate performance under the same operating conditions.

This provides some assurance that an efficient pulsed

thruster design may provide better performance under

continuous operation, and justifies the use of the pulsed

thruster facility as a first step in the development of

efficient MW-class thrusters.

MPD THRUSTER MODELING

Over the last thirty years, development of MPD thrusters

has mostly proceeded empirically, thus only pertinent to

the specific geometries and operating conditions

examined by those experiments. The lack of complete

theoretical tools has impeded the general understanding

of the interdependent physical processes in self-field

thrusters. Several numerical models7

 have been

constructed to study MPD operation, but due to

constrained computational resources they have generally

approached the problem within over-simplifying

assumptions. In the early ‘90s, a more sophisticated MHD

code was employed to study kW-level, applied-field

operation with great success.12 The current thruster

development program at the Glenn Research Center will

once again utilize such a tool along with the insights

gained from the previous effort to provide experimental

guidance in the design of efficient self-field and appliedfield MPD thrusters operating at MW power levels. 

The MACH2 Code

In the mid-1980s the Mission Research Corporation

developed a magnetohydrodynamics (MHD) code under

Air Force contract to study collisional plasmas for

problems of complex geometries.21 Over the years, this

Multi-block Arbitrary Coordinate Hydromagnetic

(MACH) simulation tool has been utilized to model a

diverse range of plasma problems and has been

substantially upgraded through contributions from a

variety of scientists. It is currently being maintained and

upgraded to a three-dimensional solver (MACH3) by The

Center for Plasma Theory and Computation at the Air

Force Research Laboratory.

MACH2 is a time-dependent, two-dimensional,

axisymmetric, multi-material code that can be applied to

problems of complex geometries due to its multi-block

structure.22 The computational mesh can move in an
NASA/CR—2001-211114 7

Figure 7. Schematic of the multi-megawatt, MY-II

MPD thruster used for the MACH2 simulations.

Dimensions in mm: rc=4.75, ra=25, re=52, Lc=45, 

Ls-a=75, La=104.

Arbitrary-Lagrangian-Eulerian (ALE) fashion allowing

applicability to both diffusive- and dispersive-dominated

problems as well as code validation. The mesh can be

refined via a variety of adaptive schemes to capture

regions of varying characteristic scale. The set of the

single-fluid, MHD equations is time-advanced with

finite-volume spacial differencing, and the boundary

conditions are applied via the ghost-cell technique so that

no special conditional statement is necessary at the

boundaries. 

The mass continuity and momentum equations assume a

compressible, viscous fluid with the latter including both

real and artificial viscosity effects. The stress-deviator

can be chosen to evolve under elastic stress for strength

of material calculations23 or modeled as a Newtonian fluid

to upgrade the code to a Navier-Stokes solver.24 The code

includes two ablation models that allow mass addition

due to solid evaporation and have been successfully

employed to model ablation-fed pulsed plasma thrusters.25

The electrons, ions and radiation field are in thermal nonequilibrium, so MACH2 solves three energy equations.

These include thermal conduction with anisotropic

transport26 and three different models for radiation

cooling.21,27 Evolution of the magnetic field is prescribed

by the induction equation that includes resistive diffusion,

the Hall effect and the thermal source for magnetic fields.

Various models for the plasma resistivity are available.

They comprise classical anisotropic resistivity,26 several

anomalous resistivity models and contributions from

electron-neutral collisions applicable to weakly ionized

gases.21 In many engineering applications the source of

magnetic flux is applied currents produced from

externally-applied voltage differentials. For this, the code

includes a variety of circuit models such as LRC, PulseForming-Networks, sine-waveforms and several others.

Other additional physical models include laser-pulse

energy deposition and a detonation package.

The set of the MHD equations is completed by an

equation of state that can be either analytic or tabular. The

latter is provided by the SESAME library that includes

semi-empirical models for the thermodynamic properties,

transport coefficients, (including opacities) and average

ionization state under local thermodynamic equilibrium.

These models have been constructed and are being

maintained by the T-1 and T-4 groups at Los Alamos

National Laboratory.28

The level of sophistication and capability of the MACH2

code has been instrumental in providing invaluable

insights to a variety of plasma problems. Some of these

include plasma opening switches,29 inertial-confinement

fusion and alternative concepts,30 compact toroid

formation and acceleration schemes,31 gas and solid

density z-pinch implosion physics,32 laser-target

interactions,33 high-power plasma source design,34

magnetic nozzles35 and a variety of plasma thrusters.12,18

Its diverse success establishes the code as the primary

numerical tool in the development of MW-class MPD

thruster designs at NASA GRC.

MW-class Self-field MPD Simulations

In an effort to provide code validation and useful insights,

MACH2 was employed to model real MW-class thrusters

in order to compare to experimental data. The particular

thruster selected for the simulations was one of the

Japanese MY36,37 family of MPD configurations that was

operated with and without external magnetic fields. In

this manner, MACH2 could be compared to both selffield and applied-field operation within the power level of

interest; 0.5-6MW. 

Specifically, the simulated cases addressed the MY-II

version of the thruster (Fig. 7) that was operated in the

aforementioned power range at a Hydrogen mass flow

rate of 1.37 g/s. The physics included viscosity effects

and real equation of state for the propellant gas along

with thermal non-equilibrium. Even though available,

these simulations did not include non-uniform massinjection schemes as it was shown that such effects were

negligible in this regime of operation. The computational
NASA/CR—2001-211114 8

&DWKRGH

&DWKRGH

&DWKRGH

Figure 8. Current Distribution comparisons at 10kA. 

Top: MACH2 simulation with electron-neutral collisions

included in the resistivity model. Thrust=36.9N,

Voltage=183V. 

Middle: Experiment, Thrust=34N, Voltage=200V,

efficiency=21%.

Bottom: MACH2 simulation with classical resistivity,

Thrust=32.4N, Voltage=108V.

 ̄ 
1.0328x10 	4  ̄

 ln

μoT3/2

e

 1e

μoe

1	



meTe

e

m2 /s (1)

domain extended to 300 mm in order to capture plume

processes such as current distension and thus additional

acceleration. For this, calculation of thrust and deposition

to various energy modes was performed at the outlet

boundaries of the computational region. The solidelectrode boundary conditions were no-slip, electrical

conductors with thermal conduction modeled at a

constant wall temperature. These temperature values were

chosen at 1000K for the anode and 2000K for the cathode

consistent with experimental measurements.36

Comparisons of the steady state simulations to the

experimental data for a discharge current of 10 kA

(Fig. 8) were quite favorable and provided new insights

that demanded further interrogation. Principally, they

have identified the importance of electron-neutral

collisions in this regime of operation, especially with

respect to energy deposition. In particular, the plasma

electrical diffusivity in MACH2 can be modeled to

include contributions from such collisions as follows

(shown in scalar form):

where Te is the electron temperature in eV, e is the

elementary charge, me is the electron mass,  is the mean

ionization ( ̄ 1), μo is the permeability of free space, ln

is the Coulomb logarithm and 1e is an appropriate

collision cross-section. For Hydrogen this electronneutral collision cross section was taken at 5x10-19 m2

 as

a reasonable average over a range of electron energies.38

The first term represents the classical Spitzer-Härm

formulation while the second is additive and accounts for

electron-atom collisions.

Exclusion of such a contribution implied that the major

energy losses are due to fall voltages and other external

circuit losses as it dramatically underestimated the total

experimental voltage (Fig. 8). Specifically, the difference

of 92V between the calculated plasma voltage and the

experimental total voltage suggests that such losses are of

the order of 0.92MW, in other words almost half the

available power is not deposited into the plasma and thus

has no opportunity to be converted to useful thrust power.

On the other hand, inclusion of such physics substantially

elevated the plasma voltage (and thus plasma power

deposition) implying that the predominant inefficiencies

are due to frozen-flow losses. Furthermore, this is

accurately illustrated by a power budget accounting for

individual energy deposition to different modes and

external losses. This is given below with the value in

parenthesis corresponding to calculated values under

classical transport:

Axial Thrust Power = 0.5 MW (0.377 MW)

Radial Thrust Power = 0.123 MW (0.08 MW)

Ionization Power = 0.608 MW (0.352 MW)

Thermal Power = 0.445 MW (0.251 MW)

Electrode Conduction = 0.154 MW (0.027 MW)

Fall Voltage Power = 0.17MW ( 0.92 MW)

The comparisons of the two-dimensional current

distribution to experiment shown in Fig. 8 leave no doubt

as to the validity of the formulation. Further support is

provided by a simple calculation of the total power

necessary to fully ionize the hydrogen propellant at

1.37g/s. This requirement is of the order of 2 MW which

in turn insinuates that the propellant may be highly

incompletely ionized; interrogation of the MACH2

calculated average degree of ionization did indeed

confirm the implication. Specifically, such ionization was

less than 10-3 in the straight section of the discharge

chamber, implying that electron-ion collisions do not

dominate. It is noteworthy that comparisons of the total
NASA/CR—2001-211114 9

1. Gilland, J., "Mission and System Optimization of

Nuclear Electric Propulsion Vehicles for Lunar and Mars

Missions," NASA CR–189058, Dec 1991.

2. Gilland, J. and Oleson, S., "Combined High and Low

Thrust Propulsion for Fast Piloted Mars Missions,"

NASA CR–190788, Nov 1992.

3. Filliben, J.D., "Electric Propulsion for Space

Applications," CPTR96–64, Chemical Propul-sion

Information Agency, Columbia, MD, 1996.

4. Oleson, S., "Advanced Electric Propulsion for Space

Solar Power Satellites," NASA/TM—1999-209307, Aug

1999.

5. Nerheim, N. and Kelly, A., "A Critical Review of the

Magnetoplasmadynamic Thruster for Space

Applications," NASA CR–92139, Feb 1968.

6. Sovey, J. and Mantenieks, M., "Performance and

Lifetime Assessment of Magnetoplasmadynamic Arc

Thruster Technology," NASA TM–101293, 1988.

7. Myers, R., Mantenieks, M., and LaPointe, M., "MPD

Thruster Technology," NASA TM–105242, 1991.

8. Tikhonov, V., Semenikhin, S., Brophy, J., and Polk, J.,

"Performance of 130-kW MPD Thruster with an External

Applied Field and Lithium as Propellant," IEPC–97–120,

25th International Electric Propulsion Conference,

Cleveland, OH 1997.

9. Kudysis, A., Emsellem, G., Cassady, L., Polk, J., and

Choueiri, E., "Lithium Mass Flow Control for High

Power Lorentz Force Accelerators", AIP Conference

Proceedings 552, Space Technology and Applications

International Forum (STAIF) 2001, M. El-Genk (ed),

American Institute of Physics, Feb 2001, pp. 908–915.

10. Polk, J., Tikhonov, V., Semenikhim, S., and Kim, V.,

"Cathode Temperature Reduction by Addition of Barium

in High Power Lithium Plasma Thrusters", AIP

Conference Proceedings 504, Space Technology and

Applications International Forum (STAIF) 2000, M.

El-Genk (ed), American Institute of Physics, Feb 2000,

pp. 1556–1563.

thrust alone could not have offered such understanding or

suggested inclusion of the improved model for the

electrical diffusivity as both calculated thrust values can

not be distinguished within experimental error. 

Based on the above understanding the MACH2

simulations will proceed to model the range of current

levels performed by the MY-II experiments. This should

provide elevated confidence in the code’s capabilities as

well as generalize the aforementioned deductions for this

regime of operation. Concurrently, MACH2 will be

utilized to guide the design toward proper expansion

nozzles that can recover some of the frozen-flow losses

calculated above. In addition, code improvements are

underway that include sheath models to calculate fall

voltage and study electrode erosion when utilized in

conjunction to the already existing ablation models.

CONCLUDING REMARKS

A MW-class pulsed MPD thruster test facility is currently

operational at the NASA Glenn Research Center. This

facility is used as a testbed for the economical evaluation

of new high power MPD thruster designs, with the goal

of transitioning the most promising designs to a MWclass steady-state experimental facility. The steady-state

facility, currently being refurbished and upgraded for

high power operation, will be used to evaluate thruster

performance, material and life issues at power levels from

0.1-MW to 1-MW. In concert with the experimental

program, the MACH2 code is being used to understand

the basic physics of high power MPD thruster operation,

and has already led to new insights and new directions

toward efficient gas-fed thruster designs.

Near-term plans for the NASA GRC high power MPD

thruster program include use of the pulsed test facility to

validate specific MW-class thruster designs developed

using the MACH2 numerical simulations. These designs

include self-field devices with contoured expansion

nozzles to recapture frozen flow losses, and applied-field

thrusters operated with and without expansion nozzles.

Efficient quasi-steady thruster designs will be fabricated

for continuous operation and tested in the 1-MW steadystate facility. Concurrent research efforts will be

undertaken to improve cathode life and reduce anode

wear, leading to efficient, long-life electromagnetic

thrusters. The strong synergy between numerical

modeling and experimental testing at GRC provides a

solid basis for the development and eventual deployment

of high power MPD thrusters capable of meeting the

demanding propulsion requirements of future NASA

missions.

REFERENCES
NASA/CR—2001-211114 10

11. Choueiri, E. and Zeimer, J., "Quasi-Steady

Magnetoplasmadynamic Thruster Measured Performance

Database", AIAA–98–3472, 34th Joint Propulsion

Conference, Cleveland, OH, July 1998.

12. Mikellides, P. G., Turchi, P. J., and Roderick, N. F.,

"Applied-Field Magnetoplasmadynamic Thrusters, Part

1: Numerical Simulations Using the MACH2 Code", J.

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13. Mikellides, P. G., and Turchi, P. J "Applied-Field

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Power, 16 (5), Sept–Oct 2000, pp. 894–901.

14. Sovey, J., Mantenieks, M., Haag, T., Raitano, P., and

Parkes, J., "Test Facility and Preliminary Performance of

a 100-kW Class MPD Thruster", NASA TM–102021,

May 1989.

15. Myers, R., "Scaling of 100-kW Class Applied-Field

MPD Thrusters", AIAA–92–3462, presented at the 28th

Joint Propulsion Conference, Nashville, TN, July 1992.

16. Myers, R., Mantenieks, M., and Sovey, J., "Geometric

Effects in Applied-Field MPD Thrusters",

AIAA–90–2699, presented at the 21st International

Electric Propulsion Conference, Orlando, FL, July 1990.

17. Myers, R., "Applied-Field MPD Thruster Geometry

Effects", NASA CR–187163, August 1991.

18.Myers, R., "Applied-Field MPD Thruster Performance

with Hydrogen and Argon Propellants", J. Propulsion and

Power, 9 (5), Sep–Oct 1993, pp. 781–784.

19. Auweter-Kurtz, M., Epple, J., Habiger, H., Kurtz, H.,

Loesener, O., Messerschmidt, E., and Mulzer, D., "Steady

State MPD Devices for Reentry Simulation",

IEPC–88–114, presented at the 20th International Electric

Propulsion Conference, Garmisch-Partenkirchen, W.

Germany, Oct 1988.

20. Auweter-Kurtz, M., Boie, C., Kaeppeler, H., Kurtz,

H., Schrade, H., Sleziona, P., Wager, H., and Wegmann,

T., "Magnetoplasmadyanmic Thrusters—Design Criteria

and Numerical Evaluation", presented at the International

Symposium on Simulation and Design of Applied

Electromagnetic Systems, Sapporo, Japan, Jan 26–30,

1993.

21. Peterkin, R.E. and Frese, M.H., MACH: A Reference

Manual—First Edition, Air Force Research Laboratory,

Kirtland AFB, New Mexico, September 14, 1998.

22. Frese, M.H., “MACH2: A Two-Dimensional

Magnetohydrodynamics Simulation Code for Complex

Experimental Configurations,” AMRC–R–874,

September 1986.

23. Peterkin, R.E., Jr., and Frese, M.H., “A Material

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24. Mikellides, P.G., “A Theoretical Investigation of

Magnetoplasmadynamic Thrusters,” Ph.D. Dissertation,

Department of Aeronautical and Astronautical

Engineering, The Ohio State University, 1994.

25. Turchi, P.J. Mikellides, I.G. Mikellides, P.G. and H.

Kamhawi, "Pulsed Plasma Thrusters for Microsatellite

Propulsion: Techniques for Optimization," MicroPropulsion for Small Spacecraft, Progress in

Astronautics and Aeronautics, Hardcover ISBN:

1-56347-448-4 , Copyright 2000.

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27. Douglas, M.R., “Radiation Production from

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Properties Data Base.. Vol Ic: EOS,” LA-1160-MS, Los

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29. Buff, J. et. al., “Enhancement of the Radiation Yield

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30. Peterkin, R.E., Jr., et. al., “Simulations of Staged

Solid Shell Implosions to Produce Fusion,” AIAA Preprint 95–2899, 31st AIAA Joint Propulsion Conference

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31. Degnan, J.H., Peterkin, R.E., Jr. et. al., “Compact

Toroid Formation, Compression, and Acceleration,”

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32. Degnan, J.H., et. al., “Electromagnetic Implosion of

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33. Stamper, J.A., et. al., “Aneurysms in laser-driven last

waves,” Physics of Fluids, 31 (11), 3353, 1988.

34. Mikellides, P.G., Turchi, P.J., and Mikellides, I.G.,

"Design of a Fusion Propulsion System, Part 1:

Gigawatt-Level Magnetoplasmadynamic Source,”

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35. Mikellides, I.G., Mikellides, P.G., Turchi, P.J. and

York, T.M."Design of a Fusion Propulsion System, Part

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36. Tahara, H., Kagaya, Y. and T. Yoshikawa, “QuasiSteady MPD Arcjets with Applied Magnetic Fields,”

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37. Tahara, H., Kagaya, Y. and Yoshikawa, T. “QuasiSteady Magnetoplasmadynamic Thrusters with Applied

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AIAA–2001–3499

E–12961

WU–755–B4–07–00

NCC3–860

17

High Power MPD Thruster Development at the NASA Glenn Research Center

Michael R. LaPointe and Pavlos G. Mikellides

Propulsion; Electric propulsion; Electromagnetic propulsion;

Plasma propulsion; Magnetoplasmadynamic thruster

Unclassified -Unlimited

Subject Category: 20 Distribution: Nonstandard

Ohio Aerospace Institute

22800 Cedar Point Road

Brook Park, Ohio 44142

Prepared for the 37th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE, and ASEE,

Salt Lake City, Utah, July 8–11, 2001. Project Manager, Dhanireddy Reddy, Power and On-Board Propulsion

Technology Division, NASA Glenn Research Center, organization code 5430, 216–433–8133.

Propulsion requirements for large platform orbit raising, cargo and piloted planetary missions, and robotic deep space

exploration have rekindled interest in the development and deployment of high power electromagnetic thrusters. Magnetoplasmadynamic (MPD) thrusters can effectively process megawatts of power over a broad range of specific impulse

values to meet these diverse in-space propulsion requirements. As NASA’s lead center for electric propulsion, the Glenn

Research Center has established an MW-class pulsed thruster test facility and is refurbishing a high-power steady-state

facility to design, build, and test efficient gas-fed MPD thrusters. A complimentary numerical modeling effort based on

the robust MACH2 code provides a well-balanced program of numerical analysis and experimental validation leading to

improved high power MPD thruster performance. This paper reviews the current and planned experimental facilities and

numerical modeling capabilities at the Glenn Research Center and outlines program plans for the development of new,

efficient high power MPD thrusters.