The 35-year history of the development and application of the pulsed plasma thruster (PPT) is reviewed. The PPT operates by creating a pulsed, high-current discharge across the exposed surface of a solid insulator, such as a Tee on t bar. The arc discharge ablates material from this surface, thereby providing propellant that is ionized, heated, and accelerated to high speed. Typically, the current pulse lasts for a few microseconds, driven by a capacitor that is charged and discharged approximately once per second. Exhaust speeds range from 3 to 50 km/s, depending on the details of the PPT design. We review the basic physics and types of PPTs, and discuss the performance of e ight and laboratory versions, with special attention to velocity and plume measurements. We also present the status of PPT theory and modeling, with emphasis on mass evolution and plasma acceleration, and describe recent variations on PPT operation, laboratory thrust measurement techniques, and the separate components of PPT efe ciency.

Pulsed Plasma Thruster

This dissertation presents research aimed at extending the efficient operation of 1600 s specific impulse Hall thruster technology to the 2000 to 3000 s range. Motivated by previous industry efforts and mission studies, the aim of this research was to develop and characterize xenon Hall thrusters capable of both high-specific impulse and high-efficiency operation. During the development phase, the laboratory-model NASA 173M Hall thrusters were designed and their performance and plasma characteristics were evaluated. Experiments with the NASA-173M version 1 (v1) validated the plasma lens magnetic field design. Experiments with the NASA 173M version 2 (v2) showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. Comparison of the thrusters showed that efficiency can be optimized for specific impulse by varying the plasma lens. During the characterization phase, additional plasma properties of the NASA 173Mv2 were measured and a performance model was derived. Results from the model and experimental data showed how efficient operation at high-specific impulse was enabled through regulation of the electron current with the magnetic field. The electron Hall parameter was approximately constant with voltage, which confirmed efficient operation can be realized only over a limited range of Hall parameters.

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Development and Characterization of High-Efficiency, High-Specific Impulse Xenon Hall Thrusters

Thruster Options for Microspacecraft: A Review and Evaluation of State-of-the-Art and Emerging Technologies

* Presented as Paper IEPC-01-036 at the 27 International Electric Propulsion Conference, Pasadena, CA, 15-19 October, 2001. † Copyright © 2001 by the Electric Rocket Propulsion Society. All rights reserved. Recent interest in extending the performance envelope of Hall thrusters has spurred the development of several new thrusters built specifically for high specific impulse operation. The Plasmadynamics and Electric Propulsion Laboratory in conjunction with the NASA Glenn Research Center have built and successfully tested one of these thrusters in singleand two-stage operation at flow rates of 5-10 mg/s and total voltages of 300-600 V. The thruster employs a plasma lens field line topography that extends the plume focal length and an intermediate electrode composed of lanthanum hexaboride for operation in twostage mode. Performance and plume measurements have been conducted using a thrust stand and ion current density probes. The results of these measurements indicate that the plasma lens optimizes thruster performance and extends the plume focal length beyond one meter. Operation of the thruster in two-stage mode generally resulted in increased thrust above single-stage operation at the expense of efficiency.

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A High Specific Impulse Two-Stage Hall Thruster with Plasma Lens Focusing

: Propellant inefficiency material in particulate form is characterized in a laboratory pulsed plasma thruster (PPT) operating at 1 Hz with a 204 discharge energy (20 W). Exhaust deposits are collected and analyzed using a combination of a scanning electron microscope with energy dispersive x-ray analysis and microscopic imaging. Teflon(trademark) particulates are observed with characteristic dianietens ranging from over 100 micrometers down to less than 1 micrometer.

Propellant Losses Because of Particulate Emission in a Pulsed Plasma Thruster

: The University of Michigan and United States Air Force Research Laboratory have jointly developed a 5 kW class Hall effect thruster. This thruster was developed to investigate, with a variety of diagnostics, a thruster similar to that specified by IHPRPT goals. The configuration of this thruster is adjustable so that diagnostic access to the interior of the thruster can be provided as necessary, and to allow for the exploration of various thruster geometries. At nominal conditions, the thruster was designed to operate at 5 kW with a predicted specific impulse of 2200 s. The actual operating parameters at 5 kW were 2326 s specific impulse, with 246 mN of thrust at an efficiency of 57%. These conditions are comparable to those of thrusters under commercial development, making the information learned from the study of this thruster applicable to the understanding of its commercial counterparts.

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Performance characteristics of a 5 kW laboratory hall thruster

: A pulsed plasma thruster (PPT) benefits from the inherent engineering simplicity-and reduced tankage fraction gained by storing the propellant as a solid. The solid is converted to the gaseous state and accelerated by an electric discharge across the propellant face. Previous research has concluded that as little as 10% of the consumed propellant is converted to plasma and efficiently accelerated. The remaining propellant is consumed in the form of late-time vaporization and particulate emission, creating minimal thrust. Critical to improving the PPT performance is improving the propellant utilization. The present work demonstrates one possible method of increasing the PPT propellant efficiency. By measuring the PPT thrust, propellant consumption, and propellant temperature while varying the power level, duration of the experimental run, and total propellant mass, a correlation is established between decreased propropellant temperature and increased propellant efficiency. The method is demonstrated by performance measurements at 60 W and S W, which show a 25% increase in thrust efficiency, while the propellant temperature decreases from 135 to 42 deg C. Larger increases in the efficiency may be realized on-orbit where operating temperatures are commonly subzero. The dependence of propellant consumption on temperature also creates systematic errors in laboratory measurements with short experimental runs, and orbit analyses where the PPT performance measured at one power level is linearly scaled to the power available on the spacecraft.

Effect of Propellant Temperature on Efficiency in the Pulsed Plasma Thruster

A time-resolved surface temperature diagnostic for ablation-controlled arcs is in development at the Air Force Research Laboratory at Edwards AFB. The diagnostic draws on heritage from the experimental dynamic crack propagation community which has used photovoltaic infrared detectors to measure temperature rise in materials in the process of fracture. The microsecond time-scales involved in the fracture process suggest that such detectors may be applicable to the ablation-controlled discharges in pulsed plasma thrusters as a direct measurement of surface temperature during and after the arc. HgCdTe detectors are evaluated for use on the surface of an AFRL micro-pulsed plasma thruster. Evaluation of the diagnostic focuses on development of valid calibration methodology and application of the detector in the presence of a plasma. Current calibration techniques are reviewed with physical limitations discussed. An estimate for the wavelength dependent emissivity of Teflon is determined and measurements are performed in real-time on the surface of a thruster. Teflon vapor pressure is calculated for post-pulse temperature measurements.

Ronald A. Spores

Papers

Propellant Losses Because of Particulate Emission in a Pulsed Plasma Thruster

Performance characteristics of a 5 kW laboratory hall thruster

Effect of Propellant Temperature on Efficiency in the Pulsed Plasma Thruster

Microsecond timescale surface temperature measurements in micro-pulsed plasma thrusters