THE APPLICABILITY OF MPD THRUSTERS TO SATELLITE POWER SYSTEMS* R. M. Jones and L. K. Rudolph Jet Propulsion Laboratory, Pasadena, California The magnetoplasmadynamic (MPD) thruster is currently under development at JPL for a range of applications including deep space propulsion, near Earth payload transportation, and stationkeeping and attitude control of large space structures. Recent experiments tend to confirm past projections that specific impulses from 1000 to 5000 seconds at efficiencies exceeding 50% can be obtained with argon propellant. The high power self-field MPD thruster is fundamentally different than an ion thruster in that it uses electromagnetic forces rather than electrostatic to accelerate a neutral plasma. The MPD thruster has a cylindrically symmetric geometry with an annular anode ring placed at the downstream end of a discharge chamber. The discharge current flows from this anode to a centrally located cathode which extends upstream to the discharge chamber backplate. The propellant is injected through the backplate and flows through the discharge current pattern where it is ionized and accelerated by a selffield Lorentz body force (jXB). The resulting thrust and specific impulse both depend quadratically on the discharge current, while the thrust efficiency increases in a more linear fashion. For reasonable specific impulse and efficiency levels, discharge currents of order tens of kiloamperes are necessary, leading to power levels of order megawatts. At this power, one MPD thruster can develop over 150 N of thrust in a volume similar to that of one 30-cm ion thruster. This high thrust density and the overall simplicity of the MPD thruster system lead to a low system specific mass and high reliability. The projected thruster efficiency for an argon MPD thruster is compared to that of an argon ion thruster in Fig. 1. The attainable MPD thruster specific impulse depends on the inverse square root of the propellant atomic weight; hence much higher specific impulses can be attained by using lighter propellants. Using helium or hydrogen the attainable specific impulse may be well above 10,000 sec. This specific impulse at thrust levels of tens of newtons makes a MPD propulsion system a candidate for stationkeeping and attitude control of large space structures such as a SPS. The most attractive application of MPD thrusters to satellite power systems is in the area of electric propulsion for a cargo orbit transfer vehicle (COTV). Calculations have been performed in order to compare the performance of a COTV using an ion or MPD propulsion system. It was assumed that the COTV carried an SPS size payload (millions of kilograms) and that a large solar array supplied power (^hundred megawatts) to the electric propulsion system. The LEO to GEO trip time was estimated by using a closed form analytical approximation which included factors for steering and drag losses and losses due to Earth shadowing and degradation of the solar array. The propellant for both the MPD and Ion thruster propulsion systems was assumed to be argon. The performance of the ion thrusters was that of a projected 120-cm thruster operating at 5000 and * This paper presents the results of one phase of research carried out at the Jet Propulsion Laboratory, California Institute of Technology under contract NAS7-100, sponsored by the National Aeronautics and Space Administration.
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