.T 

d-, 

NASA Technical Memorandum 1023 12 

AIAA-89-27 10 

Performance of a 100 l(w Class 

Applied Field MPD Thster 

Maris A. Mantenieks and James S. SOV& 

National Aeronautics and Space Administration 

Lewis Research Center ‘t 

Cleveland, Ohio I <. 

Roger M. Myers 

Sverdrup Technology, Inc. 

NASA Lewis Research Center Group I 

Cleveland, Ohio ‘Z 

Thomas W. Haag and Paul Raitano 

National Aeronautics and Space Adnain&ration 

Lewis Research Center 

Cleveland, Ohio 

* 

James E. Parkes 

Sverdrup Technology, Inc. 

NASA Lewis Research Center Group 

Cleveland, Ohio 

Prepared for the 

25th Joint Propulsion Conference 

cosponsored by the AIAA, ASME, SA.$, and ASEE 

Monterey, California, July 10-12, 15” 

- 

___ -~ - 

I.-- .O (aasa-iin-tam2) PEBPOR~IAMCE OF A 100 JCY 189-277 01 

. Levis Research Center) 18 p CSCL 21H , ’‘ CUSS APPLIED FIELD HPD fIiRUSTER [NAS-A, 

Unclas 

G3/20 0224989 

-* AI. 
-- 

PERFORMANCE OF A 100 kW CLASS APPLIED FIELD MPD THRUSTER 

M.A. Mantenieks, J.S. Sovey, T.W. Haag, and P. Raitano 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135 

R.M. Myers and J.E. Parkes 

Sverdrup Technology, Inc. 

NASA Lewis Research Center Group 

Cleveland, Ohio 44135 

ABSTRACT 

Performance of a 100 kY class, applied field ?IPD thruster was evaluated and 

sensitivities of discharge characteristics to arc current, mass flow rate, and 

applied magnetic field were investigated. Thermal efficiencies as high as 

60%. thrust efficiencies up to 218, and specific impulses of up to 1150 s were 

attained with argon propellant. Thrust levels up to 2.5 N were directly 

measured with an inverted pendulum thrust stand at discharge input powers up 

to 57 kY. It was observed that thrust increased monotonically with the 

product of arc current and magnet current. 

NOUENCLATURE 

J arc current, A 

Jm magnet current. A 

m mass flow rate, kg/sec e 

Pe electric power, W (excludes magnet power) 

Pe/m specific power. J/kg a 

P1 power loss to the electrodes, Y 

PO inlet cold yes power. V 

T thrust,, N 

V arc voltage, V 

Ve 

rl thrust efficiency, T2/2&, 

‘Ith thermal efficiency 

effective anode fall voltage, V 

1 
INTRODUCTION 

The magnetoplasmadynamic (MPD) thruster has potential propulsion applications 

for Earth-orbit transfer, maneuvering of large space systems, and 

interplanetary missions. The high thruster specific impulse (1000 to 5000 s) 

minimizes propellant requirements and the demands for mass transfer to 

low-Earth-orbit. MPD thruster missions in the next decades will probably 

employ solar or nuclear electric power systems in the 30 to 300 kW range. 

Farther term missions with interplanetary propulsion applications could use 

megawatt class nuclear power systems. 

Recent interest in high power and interplanetary missions has reactivated the 

studies of steady-state MPD thruster technologies. Summaries of the early 

phases of MPD thruster work can be found in References 1-3. Most researchers 

have used pulsed or quasi-steady state MPD thrusters because of thruster/power 

system simplicity and modest vacuum facility requirements. Pulsed thrusters, 

operating for about 1 msec, are the only source of data concerning the 

discharge phenomena and performance limits of 1 to 10 MW class MPD arcjets. 

The most attractive specific impulse and thrust efficiency data to date have 

resulted from thrusters using hydrogen or lithium propellants. Since lithium 

propellant may not be desirable because of potential spacecraft and groundbased facility contamination issues, it is likely that steady-state thruster 

applications will involve the use of hydrogen or hydrogenic propellants. The 

power, thermal, and vacuum system requirements for megawatt class, 

steady-state thrusters operating with hydrogenic propellant exclude the use of 

most of the existing vacuum facilities in the United States. For example, a 

vacuum environment of less than 4~10-~ Pa (3x10-4 torr) is generally 

required to minimize the background gas entrainment effects on thruster 

performance (Ref. 4). Prior to 1972, a large percentage of the data for 10 to 

100 kW inert gas, ammonia, or hydrogen thrusters were obtained at vacuum 

facility pressures in excess of 1.3 Pa (1.0~10-~ torr). 

thrust uncertainties of 50 to 100 percent could exist because of gas 

entrainment in applied field devices (Refs. 3,4). More recently, self-field, 

steady state MPD thrusters using argon and nitrogen have been evaluated at 

vacuum tank pressures between 1 and 2.7 Pa (Ref. 5). These results showed 

that, for powers up to 75 kW, reliable self-field thruster data could be 

obtained at facility pressures less than 1.3 Pa. 

It was shown that 

This paper describes initial test results, including direct thrust 

measurements, in a MPD thruster test facility capable of providing reliable 

performance data for a steady-state, applied-field thruster operating up to 

250 kW. A brief description of the test facility, including the vacuum 

chamber, power supplies, a closed cycle thruster water cooling system, and 

propellant metering is provided. Complete details of the test facility may be 

found in Reference 6. 

Preliminary performance measurements are presented for a number of MPD 

thruster configurations operating at arc currents of up to 3000 A and arc 

powers of over 100 kW. For the convenience of the reader, some relevant data 

from Reference 6 are included in this report. Argon was mainly used as the 

propellant, but thermal efficiencies of helium, neon, and nitrogen are also 

reported. Basic thruster performance parameters discussed are discharge 

2 
I. 

characteristics, thrust measurements, thermal efficiencies, and magnetic field 

effects . 

Diagnostic methods and instrumentation capable of measuring plasma parameters, 

current distribution, plasma species identification, and distribution of the 

cathode temperatures are described in Reference 7. The structure, operation, 

and performance characteristics of an inverted pendulum thrust stand are 

reported in Reference 8. 

APPARATUS AND TEST PROCEDURE 

The test facility consisted of the vacuum chamber, a WD thruster and magnet, 

a propellant metering panel, water-cooling apparatus, the thrust stand, the 

power system, and a data acquisition system. This section provides 

descriptions of each of these systems. 

MPD THRUSTER TEST FACILIX 

The MPD thruster used in this investigation was a modified commercial coaxial 

plasma generator that is commonly used for plasma spraying (Ref. 9). This 

commercial device was selected because it was capable of routinely operating 

at 30 to 60 kW for extended periods of time and could be easily retrofitted 

with different anode and cathode designs. The various thruster configurations 

tested are shown in Fig. 1. All of the thrusters contained a copper anode 

assembly in which the water circuit was connected in series with the cathode 

water cooling hardware. The 13 mm diameter, 2 percent thoriated tungsten 

cathode was brazed to a water-cooled copper retainer. The anode and cathode 

were isolated using fluoropolymer and alumina insulators. Table I lists the 

relevant dimensions of the six thruster configurations that were evaluated. 

Thrust data were obtained with Thrusters D, E, and F. 

In order to enhance the electromagnetic thrust component, the anodes were 

designed with larger channels (Refs. 10,111 and the current range was extended 

to 3000 A. The mass flow rates were also reduced to 5rovide inlet pressures 

between 1.3 and 13 kPa (10 to 100 torr). Values of J /$ ranged from lxlOIO to 

3x1Ol1 A2-.s/kg, which generally implied a significant electromagnetic thrust 

contribution (Refs. 3,121. 

The magnet used in the present investigation (Fig. 2) was designed to provide 

a magnetic field of approximately 0.3 Tesla at the center of the magnet for a 

magnet current of 1500 A. It was a 14 cm I.D., 15 cm long solenoid consisting 

of 46 coils of copper tubing covered with glass sheath insulat!.on. The 

magnet had four layers of ten coils each, and a final layer with six coils. 

Measured and calculated axial magnetic field strengths are shown in Fig. 3. 

The axial field strengths measured in the region of the cathode tip and the 

thruster exit plane were about 0.23 and 0.16 T, respectively, for a magnet 

current of 1400 A. 

The thrust stand, fully described in Reference 8,and shown in Fig. 41 was an 

inverted pendulum whose upright flexures were designed to deflect about 5 cm 

at a thrust of 5 N. The pivoting structure of the thrust stand transferred 

the weight of the thruster to a system of flexures anchored to the lower 

3 
frame. The flexures pivoted about an axis 1.5 m below the thruster 

centerline. The major restoring force was provided by the 16 nun diameter 

primary flexures. The time derivative of the thrust stand displacement was 

used to drive an electromagnetic damper. Thrust stand displacement was 

measured by a movable iron core in a linearly variable differential 

transformer (LVDT) which was mounted to a fixed reference structure. The 

thrust stand assembly contained a remotely controlled leveling mechanism to 

vary the angle of the stand and adjust the LVDT signal. Remote calibration of 

the stand was obtained by applying small weights along the thrust axis using 

monofilament line and a precision pulley. 

Tares due to thruster discharge current were shown to be negligible. 

due to the applied magnetic field were reduced, after considerable effort, to 

values less than 3.0 percent of measured thrust for arc discharge currents 

between 1000 to 2500 A and magnet currents up to 1000 A. The tares due to the 

applied field could be determined independently and were subtracted from the 

indicated thrust measurement. Thermally induced drift of the thrust signal 

was on the order of 50 mN during test runs of about an hour and this drift 

could also be acccunted for by an in-situ thrust calibration procedure. 

Tares 

The thruster and thrust stand were mounted to a vacuum chamber bulkhead and 

were inserted into a 3 m diameter by 3 m long port. The vacuum chamber was 

evacuated using nineteen, 0.8 m diameter hydrocarbon oil diffusion pumps with 

freon traps cooled to 236 K. 

nitrogen, and argon were about 4.5~15-~, 8.0~15-~, and 15~10-~ kg/s, 

resepctively, for a main chamber pressure of 7~10-~ Pa (5~10~~ torr). 

Maximum allowable flow rates for helium, 

The propellant feed system was designed to provide three parallel gas paths 

such that one or more gases might be available for thruster operation. 

Flowmeter ranges were 0.5, 5, and 50 standard liters of nitrogen per minute. 

Flowmeter accuracy was estimated to be within 2 percent over the range of 10 

to 100 percent of full-scale. The flowmeters were directly calibrated with 

argon and helium. Measured values of nitrogen and neon flow rates were based 

on the vendor's calibration factors. 

The thruster inlet pressure was sensed using a capacitance manometer. While 

the thruster chamber pressure could not be measured directly, the inlet 

pressure provided an upper bound of the chamber pressure, as a significant 

pressure drop probably existed across the propellant injector. 

Water cooling to the arc and magnet was provided by two separate pwnp/heat 

exchanger assemblies capable of supplying 0.8 liter/s of water at a pressure 

of 1 ma. The MPD thruster cathode and anode were cooled by the same 

pump/heat exchanger using de-ionized, distilled water in the closed-loop 

system. Water flow rate data were obtained from a calibrated turbine meter. 

The power lost to the electrodes was obtained by measuring the water flow rate 

and the change in water temperature. Digital readouts were used to indicate 

the water temperature to the nearest degree Fahrenheit. 

The power system was composed of the arc supply, a high voltage arc starter, 

the magnet supply, and safety sensors and interlocks. The arc supply 

consisted of four welding power supplies connected in a series-parallel 

4 
arrangement which were isolated from the vacuum facility ground to minimize 

extraneous arcing. 

69 V. The open circuit voltage was 160 V. Arc voltage was directly measured 

with potential leads across the MPD thruster. The arc was started using a 

950 V, half-wave rectifier connected to the thruster via a 100 ohm resistor. 

A single welding supply provided power to the electromagnet. Maximum steady 

state current was 1500 A. The power supply was isolated from ground as well 

as from the arc supply. 

Maximum output power capability was 264 kY at 3000 A and 

PROCEDURE 

The arc was started by setting the argon flow rate to approximately 5.0~10-4 

kg/sec, initiating a low current, glow discharge from the starter supply, and 

activating the arc supply which was preset for 2000 A operation. Approximately 

four seconds after arc ignition, the arc current was reset to 1000 A and the 

arc starter was turned off. The flow rate was then reduced to about 

9.0~10-~ kg/sec which was more typical of the low pressure MPD arc 

operation. In all cases, the arc was started with argon propellant. 

Transition to another propellant was accomplished by closing the argon 

isolation valve and opening the valve for the desired propellant. 

The major thruster variables were mass flow rate, arc current, and magnet 

current. The propellant flow rate was chosen to be high enough to provide 

stable dischar e operation but low enough to yield values of J2/; in the 

lx109 to lxl0lq A2-s/kg, which are representative values for low 

pressure MPD thrusters (Ref. 3). The upper limit of the arc current was 

selected to prevent thermal failure of the water coolant O-rings in the anode 

assembly. The range of magnet currents was defined by a minimum current 

(-500 A) to provide a stable discharge and by the magnet power supply limit 

1500 A. 

During thruster operation, the anode potential was found to float within 10 V 

of facility ground at argon flow rates in excess of 4.0~10-5 kg/sec. 

However. the anode floated to about 40 V above ground at very low flow rates 

of 8.0~10-6 kg/sec. 

external to the thruster as long as the arc supply was floating, the power 

cables were insulated, and the facility pressure was less than 7x10-* Pa. 

In all cases, there was very little evidence of arcing 

Thrust measurements were obtained after approximately 20 minutes of thruster 

operation at constant operating conditions. The thruster was then turned off 

to obtain a zero thrust reference level and to measure tares due to the 

applied magnetic field. 

RESULTS AND DISCUSSION 

Preliminary experiments, with several MPD thruster configurations (Fig. l), 

were directed at determining the sensitivity of discharge parameters to 

geometry and propellant flow rate (Ref. 6). The following discussion will 

focus on experiments to gain insight into MPD thruster performance by 

examining: 1) thermal efficiency, 2) effective anode voltage drops, and 31, 

thrust measurements over a range of thruster configurations, operating 

conditions, and applied magnetic fields. 

5 
THERMAL EFFICIENCY 

It has been observed that a large fraction of the MPD thruster input power was 

deposited in the electrodes at input power levels up to about 300 kW. For 

example, in previous work the maximum power available for thrust production 

with hydrogen. ammonia, and argon HPD arcjets was typically 80, 70, and 60 

percent of the input power. respectively (Refs. 5, 13, and 14). These data 

were obtained from water-cooled devices which are less efficient and are 

presumed to have larger anode voltage drops than radiation-cooled UPD 

thrusters (Ref. 2). The thermal efficiency is defined as: 

nth = P, iePY 1 - P1/Pe (1) 

The cold gas power, Po, was neglected in the calculations since it was less 

than 0.3 percent of the electric power for all data reported. To simplify 

comparison of data, the magnet power was not included in the parameter, Pe. 

An effective anode fall voltage may be defined as: 

and has been used by others in describing anode loss phenomena (Ref. 15). In 

Fig. 5, the effective anode fall voltage is plotted against specific power for 

argon, neon, and helium propellants. In this figure, the data from thruster 

configurations E and F have been added to data found in Reference 6. 

Thrusters, operating at specific powers of 37 to 100 MJ/kg, had effective 

anode fall voltages between 10 to 15 V. 

inlet ressures in excess of 13 kPa and values of J2/m between 1.8 and 7.1 

x109 A%-s/kg. The MPD thrusters operating at higher specific powers and 

higher values of J2/$ had effective anode fall voltages between 15 to 30 V 

which were a major fraction of the arc voltages measured using Thrusters A to 

F. 

Thrusters opsrating in this range had 

The Z1I2V parameter was used in Reference 6 to correlate the thermal 

efficiency data for argon, helium, nitrogen, and neon using Thrusters A-D. 

The data for thruster configurations E and F were added to the data of 

Reference 6 and are shown in Fig. 6. 

The thermal efficiency was shown to be very sensitive to the i1/2V 

parameter in the 0.13 to 0.32 V-kg1/2s'1/2 range. 

efficiency appears to approach an asymptote are thermal arc results (high mass 

flow rates and dominant Joule heating). The conical anode thrusters (E and F) 

exhibited higher anode losses and these data did not follow the thermal 

efficiency trends of the straight channel thrusters (A-D). A clear parametric 

relationship between the thermal efficiency data and arc current, J2/~, 

power, or specific power was not evident. 

The data were taken between 1 and 3 kA. 

The data where the thermal 

THRUST MEASUREMENTS 

Thrust measurements, as a function of arc power, are shown in Fig. 7. 

Thruster configurations D,E, and F were evaluated with argon propellant. The 

power was increased at a given arc current by increasing the applied magnetic 

6 
field which produced an increase in arc voltage. The measured thrust 

increased nearly linearly with power at constant arc current. Thrust levels 

between 0.8 and 2.5 N were obtained over a 20 to 62 kW range of input electric 

power. 

Fig. 8 shows the specific impulse as a function of specific power for the 

various thruster configurations using argon. The specific impulse also 

increased almost linearly with increasing specific power at constant arc 

current. The highest values of specific impulse (-1100 s) were obtained with 

thruster configurations E and D at discharge currents of 1000 and 1500 A, 

respectively. 

The data of a thermal arcjet and a self-field WD thruster from References 16 

and 17, respectively, are included in Fig. 8 for comparison. The specific 

impulse of 720 s, attained with an argon thermal arcjet is the highest value 

of specific impulse found in the available literature for that operating 

mode. The specific impulse levels of Thrusters D and E exceeded the thermal 

arcjet data by more than 50% implying electromagnetic field thrust 

augmentation was significant. 

Arc currents of 1000 to 1500 A (J2/i of 6.3~109 and 1.4~10~~ A2-s/kg) 

were required to attain 1100 s specific impulse using Thrusters E and D, 

respectively, with argon propellant. These results compare favorably to the 

self-field MPD thruster which provided a specific impulse of 1200 s at a 

J2/m of 2.8~10~~ A2-s/kg. 

efficiency was between 0.21 and 0.24. 

In all three cases, the overall thrust 

Figure 9 depicts the thrust to power ratio as a function of specific impulse 

for the various thrust configurations. As mentioned previously, most of the 

thrusters did not operate stably below a magnet current of 500 A; however, on 

several occasions Thruster D was stable long enough to permit a thrust 

measurement. As the data from Thruster D at an arc current of 1000 A 

indicate, there was a steep increase in the ratio of thrust to power as the 

magnetic field current was increased from zero to approximately 500 A. The 

increase in the thrust to power ratio was 50% with an associated 64% increase 

in the specific impulse over the applied magnetic field increment. Further 

increases in magnetic field resulted in increases of specific impulse without 

a noticeable thrust to power ratio decrease. This trend was observed for both 

thruster configurations (D and F). 

The thermal and thrust efficiencies of the thruster configurations D,E, and F 

are shown in Fig. 10. Thermal efficiencies as high as 0.47 were obtained with 

the flared anode thruster using argon propellant. Generally, lower thermal 

efficiencies were observed as arc current was increased from 1000 to 1500 A. 

This change usually resulted in lower arc voltage and a higher effective anode 

voltage drop as shown in Figure 5. Both thrusters D and E, operating at 1000 

A, converted 39 to 47 percent of available power (Pe- P1) for propulsion 

to thrust power. This is a rather high fraction since approximately 15 to 20 

percent of the power available for propulsion must be consumed to fully ionize 

the argon. 

7 
The highest thrust efficiencies (21%) were obtained with thruster 

configurations D and E. Because of arc instabilities and thermal problems, 

thruster configuration F was constrained to operate at higher propellant mass 

flow rates, thus limiting the maximum attainable specific impulse to 700 s. 

Thrust efficiencies of 20 to 25% with argon have been reported with watercooled, self-field, and pulsed thrusters (Refs. 10,131. The thrust 

efficiencies of the three thruster configurations (D,E, and F) operating at 

the various arc currents tended to increase as the specific impulse was 

increased by an increasing magnetic field. 

It should be mentioned that discharge characteristics were not always 

repeatable from day to day while operating at the same thruster parameters. A 

possible reason for this may be that the cathodes, while being of the same 

geometry, may not have had similar emission characteristics due to thoria 

depletion or due to cooling water leaks because of failure of water seals. 

Current distribution measurements also indicated different profiles for 

similar thruster configurations and identical operating parameters (Ref. 7), 

indicating that the observed voltage differences are not totally attributable 

to electrode fall voltages. Cathode deterioration was also evidenced by 

difficulty in establishing an arc discharge. More detailed studies of MPD 

thruster discharge characteristics and the plasma exhaust are required to 

understand this behavior. 

MAGNETIC FIELD EFFECTS 

The application of a magnetic field to the MPD thruster had two primary 

effects: first, it enlarged the operating envelope of the thruster in which 

stable operation was achieved; second, the performance increased with 

increasing magnetic field. The first result has been observed by previous 

investigators (Refs. 12.14,15). The second result also was to be expected, 

because the applied field introduces additional force terms besides the 

electrothermal term. and the "pumping" and "blowing" force terms due to 

self-induced fields (Ref. 2). Due to the complications arising from the 

interactions of these forces, no universally accepted model for applied field 

MPD thruster operation exists. 

Following the practice of previous investigators, the thrust in Fig. 11 was 

plotted as a function of the product of the arc current and magnet current for 

thruster configurations D, E, and F. All of the thrusters exhibit a monotonic 

dependence with arc current-magnet current product. This result is similar to 

previous investigations (Refs. 17-21). 

FoI thruster configuration D, the thrust increased by 64% as the applied 

magnet current increased from 0 to 1500 A. These results indicate that the 

thrust should increase monotonically with product of the arc current and 

magnet currents up to 2.2~10~ A2.

saturation effect with increasing magnet current was found. 

In contrast to References 22 and 23, no 

8 
CONCLUDING REMARKS 

Performance of 100 kW class, applied field MPD thrusters were evaluated. Six 

thruster configurations were characterized as functions of discharge current, 

mass flow rate, and applied magnetic field. Thrust measurements were obtained 

with three of the thruster configurations. 

Higher arc voltages and thermal efficiencies were obtained using the larger 

diameter anode channel (26 mm) thruster with a flared anode. Representative 

effective anode fall voltages, estimated by anode power losses ranged from 15 

to 30 V. 

as the propellant: however, measured thrust efficiencies were limited to only 

21%. A thrust of 2.5 N was measured at an arc discharge power of 57 kW. The 

highest measured specific impulse was 1150. 

compares favorably with values found in the literature using self-field MPD 

thursters at comparable J2/& parameters. 

nearly constant with increasing specific impulse at magnetic current values of 

greater than 500 A. Finally the measured thrust was shown to be monotonically 

inceasing with the product of the arc and magnet currents. 

Thermal efficiencies as high as 60 percent were obtained using argon 

This specific impulse 

The thrust to power ratio was 

9 
REFERENCES 

1. Nerheim, N.M. and Kelly, A.J., "A Critical Review of the State-of-the-Art 

of the MPD Thruster," AIM Paper 67-688, Sept. 1967. 

2. Seikel, G.R., York, T.M., and Condit, W.C., "Applied-Field 

Magnetoplasmadynamic Thrusters for Orbit-Raising Missions," Orbit-Raising 

and Maneuvering Propulsion: Research Status and Needs. (Progress in 

Astronautics and Aeronautics, Vol. 89), L.H. Caveny, Ed., AIAA, New York, 

1984, pp. 260-286. 

3. Sovey, J.S. and Mantenieks, M.A., "Performance and Lifetime Assessment of 

MPD Arc Thruster Technology," AIAA Paper 88-3211, July 1988 (NASA 

TM-101293). 

4. Sovie, R.J. and Connolly, D.J., "Effect of Background Pressure on 

Magnetoplasmadynamic Thruster Operation," Journal of Spacecraft and 

Rockets, Vol. 7, No. 3, Mar. 1970, pp. 255-258. 

5. Kurtz, H.L., Auweter-Kurtz, M., and Schrade, H.O., "Self-Field MPD 

Thruster Design - Experimental and Theoretical Investigations," AIAA Paper 

85-2002, Sept. 1985. 

6. Sovey, J.S., Mantenieks, M.A., Haas, T.W., Raitano, P., and Parkes, J.E., 

"Test Facility and Preliminary Performance of a 100 kW Class MPD Thruster," 

NASA TM-102021, 1989. 

7. Myers, R., "Plume Characteristics of MPD Thrusters: A Preliminary 

Examination," AIAA Paper 89-2832, July 1989. 

8. Haag, T.W., "Design of Thrust Stand for High Power Electric Propulsion 

Devices," AIAA Paper 89-2829, July 1989. 

9. Product Data Sheet 387-1, Electro-Plasma, Inc., Irvine, CA, 1988. 

10. Jahn, R.G., Physics of Electric Propulsion, McGraw-Hill Inc., New York, 

1968. 

11. Wolff, M., Kelly, A.J., and Jahn, R.G., "A High Performance 

Magnetoplasmadynamic Thruster," 17th International Electric Propulsion 

Conference, Japan Society for Aeronautical and Space Sciences, Tokyo, 

Japan, 1984, pp. 206-212. 

12. Malliaris, A.C., John, R.R., Garrison, R.L., and Libby, D.R., ''QuasiSteady MPD Propulsion at High Power," AVSD-O146-71-RR, Avco Corp., 

Wilmington, MA, Feb. 1971, NASA CR-111872. 

13. Ducati, A.C., Muehlberger, E., and Todd, J.P., "Design and Development of 

a Thermo-Ionic Electric Thruster," NASA CR-59804, 1964. 

10 
14. Bennett, S., Enos, G., John, R., and Powers, W., "Magnetoplasmadynamic 

Thrustor Research," AVSSD-272-67-RR9 AVCO Corp., Space Systems Div., 

Lowell, MA, May 1967, NASA CR-72345. 

15. Cam, G.L., Harder, R.L., Moore, R.A., and Lenn, P.D., "Hall Current 

Accelerator," ESOC-5470-Fina1, Electro-Optical Systems Inc., Pasadena, CA, 

1966, NASA CR-54705. 

16. Kurtz, M.L., Auweter-Kurtz, M., Glocker, B., Merke, W., and Schrade, 

H.O., "A 15-kW Experimental Arcjet," IEPC Paper 88-107, Oct. 1988. 

17. Merke, W.D., Auweter-Kurtz, M., Habiger, H., Kurtz, H., and Schrade, H.O., 

"Nozzle Type MPD Thruster Experimental Investigations," IEPC Paper 88-028, 

Oct. 1988. 

18. Patrick, R.M. and Schneiderman, A.M., "Performance Characteristics of a 

Magnetic Annular Arc," AIM Journal, Vol. 4, No. 2, Feb. 1966, pp. 283-290. 

19. Kagaya, Y., Yoshikawa, T., and Tahara, H., "Quasi-Steady MPD Arcjets with 

Applied Magnetic Fields," AIAA Paper 85-2001, Oct. 1985. 

20. Fradkin, D.B., Blackstock, A.W., Roehling, D.J., Stratton, T.F., Williams, 

M., and Liewer, K.W., "Experiments Using a 25-kW Hollow Cathode Lithium 

Vapor MPD Arcjet," AIAA Journal, Vol. 8, No. 5, May 1970, pp. 886-894. 

21. Fradkin, D.B., "Analyses of Acceleration Mechanisms and Performance of an 

Applied Field MPD Arcjet," Ph.D. Thesis, Princeton University, 1973. 

22. Kimura, I. and Arakawa, Y., "Effect of Applied Magnetic Fields on Physical 

Processes in an MPD Arcjet," AIAA Journal, Vol. 15, No. 5, May 1977, 

pp. 721-724. 

23. Connolly, D.J., Bishop, A.R., and Seikel, G.R., "Tests of Permanent Magnet 

and Superconducting Magnet MPD Thrusters," AIAA Paper 71-696, June 1971, 

(NASA TM X-67827). 

11 
TABLE I - THRUSTER CONFIGURATIONS 

THRUSTER ANODE CHANNEL CHANNEL 

DESIGNATIONS DIAMETER LENGTH, ( a 

mm cm 

A 16.5 5.3b 

B 16.5 4.1b 

C 26 4.2b 

D 26 4.2b 

E 26d 5.4b 

F 26d 5.4= 

a Upstream end of cathode to anode exit plane 

b Cathode diameter 1.27 cm, cathode length 2.2 cm 

c Cathode diameter 1.27 cm, cathode length 3.8 cm 

d Inlet diameter 

ANODE FLARE 

HALF ANGLEILENGTH, 

deg /cm 

010 

010 

010 

41°/1. 1 

10.5O/5.4 

10.5OI5.4 

12 
PROPELLANT 

c .4- 

I 

IINJECTOR CATHODE 

CATHODE 

MAGNET CURRENT, A 

0 1000 

0 1400 

OPEN SYMBOLS DENOTE 

SOLID SYMBOLS DENOTE 

MEASURED FIELD STRENGTH 

CALCULATED FIELD STRENGTH 

rCATHODE TIP 

LOCATION 

 WATER 

\\LTHORIATED TUNGSTEN 

INLET 

INSULATOR 

(FLUOROPOLYMER)’ WATER 1 CATHODE 

OUTLET-’ 

H 

2 CM +CHANNEL 

LENGTH -+I 

(A) MPD THRUSTER (A OR B) WITH CYLINDRICAL ANODE. (REF. 6). 

PROPELLANT 

INSULATOR I N JECTOR 

(FLUOROPOLYMER)7\ (ALUM1NA)l 4 WATER-COOLED 

/I! ANODE (COPPER) 1 + +---1.1 CM 

:%E 

I ) ~ I ‘\,I \ 

WATER LTHORIATED TUNGSTEN 

+----I 2 CM OUTLET-’’ CATHODE 

+CHANNEL+ 

LENGTH 

(B) MPD THRUSTER (D) WITH FLARED ANODE. (REF. 6). 

+CHANNEL+ 

LENGTH 

(C) MPD THRUSTER (E OR F) WITH CONICAL ANODE. 

FIGURE 1. - SKETCHES OF W’D THRUSTERS. 

e 15.2 CM 

000000 

11 

(MAGNET TUBING: 1.27 

0.d.. 0.89 MM WALL, 

COVERED WITH GLASS 

CLOTH SHEATH.) 

/ 

-THRUSTER D 

-8.5 CM TYP. 

0000000000 

0000000000 

000000 

FIGURE 2. - SKETCH OF THRUSTER AND MAGNET ARRANGEMENT. 

13 
3 fl TEST SECTION 7 

\ 

3 M GATE VALVE 

(SHOWN CLOSED) - 

REFERENCE STRUCTURE - 

CALIBRATION MECHANISM-\ 

DISPLACEMENT 

rAPPLIED FIELI I MAGNET 

I 

rMPD 

THRUSER 

I 

IM 

FIGURE 4. - SCHEMATIC OF MPD THRUSER WITH THRUS STAND. (REF. 6). 

PROELANT THRUSER 

0 Ar B 

0 Ar C 

A Ar D 

0 He A 

30 

15 

IU W 

U 

101 102 103 104 

SPECIFIC POWER. MJ/kg 

FIGURE 5. - EFFECTIVE ANODE VOLTAGE DROP VERSU SPECIFIC 

POWER. 

I 

\ 

14 
THRUSTER PROPELLANl 

OB Ar 

oc AT 

AD Ar 

OA He 

n~ He 

OB Ne 

.5 

n~ AT 

VF AT MO - REF. 6 00 

- INCREASING PIAGNET CURRENT 

1 I I I I I I 

3 kA=-J=-l kA b 4- b 

0 .2 .4 .6 .8 1.0 

CORRELATION PARAMETER h%, VJk;;/h 

FIGURE 6. - THERMAL EFFICIENCY CORRELATION. 

THRUSTER ARC MASS FLOW 

CURRENT, RATE, 

OD 1000 0.158 

OD 1500 ,158 

DE 1000 .158 

A vs 

F 1500 .320 

F 2000 .320 

F 2500 .320 

ARGON PROPELLANl 

THRUSTER ARC MASS FLOW 

CURRENT. RATE, 

OD 1500 0.157 

OD 1000 .165 

AE 1000 ,158 

A g/s 

'loo - V REF. 17 4500- 

co 

j 900 - 

0 100 200 300 

SPECIFIC POWER. MJ/kg 

FIGURE 8. - SPECIFIC IMPULSE AS A FUNCTION OF SPECIFIC POWER. 

15 
6C 

2 

s 

\ 

= 5c E 

c i 

Y 

CT 40 

0 L

0 + 

c 

w 8 + 

g 30 

20 

THRUSTER ARC MASS FLOW 

CURRENT, RATE, 

OD 1000 0.165 

OD 1500 ,157 

AF 1500 .316 

A g's 

EFFICIENCY ,- THRUST 

400 800 n = 0.2 

1400 

0 200 400 600 800 1000 1200 1400 

SPECIFIC IMPULSE, s 

FIGURE 9. - THRUST-TO-POWER RATIO AS A FUNCTION OF SPECIFIC 

IMPULSE. 

2.0 

r 1.5 

THRUSTER ARC 

CURRENT, 

A 

OD 1000 

AE 1000 

OD 1500 

THRUSTER ARC 

CURRENT, 

A 

OD lo00 

AE lo00 

OF 1500 

OD 1500 

OE 1530 

MASS FLOW 

RATE, 

g/s 

0.165 

.158 

,158 

.320 

.320 

SOLID SYMBOLS MNOTE THRUST EFFICIENCY 

OPEN SYMBOLS DENOTE THERMAL EFFICIENCY 

- 

00 A 

0 

mA A 

'8 

-0 

0 ARGON PROPELLANT 

I I I 

600 800 1000 1200 

SPECIFIC IMPULSE. s 

FIGURE 10. - THERMAL AND THRUST EFFICIENCY VERSUS SPECIFIC 

IMPULSE. 

~ 

0 50 100 150 200 250 

PRODUCT OF ARC AND MAGNET CURRENT. 104 ~2 

FIGURE 11. - THRUST AS A FUNCTION OF THE PRODUCT OF ARC 

CURRENT AND MAGNET CURRENT. 

16 
National Aeronautics and 

Space Administration 

7. Key Words (Suggested by Author(s)) 

MPD thrusters 

Electric propulsion 

Report Documentation Page 

18. Distribution Statement 

Unclassified - Unlimited 

Subject Category 20 

NASA TM-102312 1. Report No. 

9. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No of pages 

Unclassified Unclassified 18 

I 2. Government Accession No. 

22. Price’ 

A03 

AIAA-89-27 10 

4. Title and Subtitle 

Performance of a 100 kW Class Applied Field MPD Thruster 

~~ ~~ 

7. Author(s) 

Maris A. Mantenieks, James S. Sovey, Roger M. Myers, 

Thomas W. Haag, Paul Raitano, and James E. Parkes 

9. Performing Organization Name and Address 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 

2. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, D.C. 20546-0001 

3. Recipient’s Catalog No. 

5. Report Date 

6. Performing Organization Code 

8. Performing Organization Report No. 

E-5006 

10. Work Unit No. 

506-42-3 1 

11. Contract or Grant No. 

13. Type of Report and Period Covered 

Technical Memorandum 

14. Sponsoring Agency Code 

5. Suplemntary Notes 

Prepared for the 25th Joint Propulsion Conference cosponsored by the AIAA, ASME, SAE, and ASEE, 

Monterey, California, July 10-12, 1989. Maris A. Mantenieks, James S. Sovey, Thomas W. Haag, and Paul 

Raitano, NASA Lewis Research Center; Roger M. Myers and James E. Parkes, Sverdrup Technology, Inc., 

NASA Lewis Research Center Group, Cleveland, Ohio 44135. 

6. Abstract 

Performance of a 100 kW, applied field MPD thruster was evaluated and sensitivities of discharge characteristics 

to arc current, mass flow rate, and applied magnetic field were investigated. Thermal efficiencies as high as 60% 

thrust efficiencies up to 21%, and specific impulses of up to 1150 s were attained with argon propellant. Thrust 

levels up to 2.5 N were directly measured with an inverted pendulum thrust stand at discharge input powers up 

to 57 kW. It was observed that thrust increased monotonically with the product of arc current and magnet 

current.