High Power Electrical Propulsion Laboratory

Laboratory, United States

High Power Electrical Propulsion Laboratory

Laboratory, United States
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Frieman J.D.,Georgia Institute of Technology | Frieman J.D.,High Power Electrical Propulsion Laboratory | Walker J.A.,Georgia Institute of Technology | Walker J.A.,High Power Electrical Propulsion Laboratory | And 4 more authors.
Journal of Propulsion and Power | Year: 2016

The impact of facility conductivity on Hall effect thruster cathode coupling is experimentally investigated. The 3.4 kW Aerojet Rocketdyne T-140 Hall effect thruster operating at a discharge voltage of 300 V, a discharge current of 10.3 A, and an anode flow rate of 11.6 mg/s serves as a representative Hall effect thruster test bed. The nominal facility operating pressure during thruster operation is 7.3 × 10-6 Torr corrected for xenon. Two 0.91 x 0.91 m square aluminum plates are placed adjacent to, but electrically isolated from, the walls of the conductive vacuum chamber at two locations with respect to the center of the thruster exit plane: 4.3 m axially downstream along the thruster centerline, and 2.3 m radially outward centered on the exit plane. The plates and body of the Hall effect thruster are configured in three distinct electrical configurations with corresponding measurements: 1) electrically grounded with measurements of currents to ground, 2) electrically isolated with measurements of floating voltages, and 3) isolated from ground but electrically connected with measurements of the current conducted between the plates. Measurements are taken as the radial position of the cathode is varied from 0 to 129 cm with respect to the nominal cathode location. Measurements of the current collected by the plates and thruster body indicate that cathode electrons preferentially travel to the thruster body, Hall effect thruster plume, and radial facility surfaces for cathode locations in the near field, midfield, and far field, respectively. These results indicate that cathode position alters the recombination pathways taken by cathode electrons in the Hall effect thruster circuit. Copyright © 2015 by Jason D. Frieman. Published by the American Institute of Aeronautics and Astronautics, Inc.


Walker J.A.,Georgia Institute of Technology | Walker J.A.,High Power Electrical Propulsion Laboratory | Frieman J.D.,Georgia Institute of Technology | Frieman J.D.,High Power Electrical Propulsion Laboratory | And 5 more authors.
Journal of Propulsion and Power | Year: 2016

The physical mechanisms that govern the electrical interaction between the Hall-effect-thruster electrical circuit and the conductive vacuum-facility walls are not fully understood. As a representative test bed, an Aerojet Rocketdyne T-140 Hall-effect thruster is operated at 3.05kWand a xenon mass flow rate of 11.6 mg/s with a vacuum facility operating neutral pressure of 7.3 × 10-6 torr, corrected for xenon. Two electrical witness plates, representative of the facility chamber walls, are placed 2.3 m radially outward from thruster centerline and 4.3 m axially downstream from the thruster exit plane. The cathode is radially translated from 18.1 to 77.8cmaway from the thruster centerline. At each cathode position, the discharge current and the electrical waveform of the radial and axial plates are simultaneously measured. As the cathode radial position changes from 18.1 to 77.8 cm from the thruster centerline, the discharge-current oscillation frequency decreases between 17 and 35% for the electrically grounded thruster-body case, and between 15 and23%for the electrically floating thruster-body case. The analysis of the electron current collected by the radial plate suggests that electrons directly sourced from the cathode impinge on the radial plate at large cathode positions. Overall, the results of this work show that the chamber walls act as an artificial electrical boundary condition that keeps the Hall-effect-thruster plume plasma potential to within certain bounds. © Copyright 2015 by Jonathan A. Walker, Jason D. Frieman, and Mitchell L.R. Walker.


Williams L.T.,Georgia Institute of Technology | Williams L.T.,High Power Electrical Propulsion Laboratory | Walker M.L.R.,Georgia Institute of Technology | Walker M.L.R.,High Power Electrical Propulsion Laboratory
Journal of Propulsion and Power | Year: 2014

Helicon plasma sources are devices that are capable of efficiently producing high-density plasmas. There is growing interest in using a helicon plasma source in space propulsion as a replacement to the direct current plasma discharge in ion engines. A radio frequency ion engine is developed that combines a helicon plasma source with electrostatic grids and a magnetically shielded anode. Thruster performance evaluation includes estimation of the discharge plasma ion density and electron temperature and measurement of the plume current density, grid currents, and thrust. The radio frequency ion engine is tested across the following operating parameter ranges: 343-600 Wradio frequency power, 50-250 G magnetic field strength, 1.0-2.0 mg/s argon flow rate, and 100-600 V discharge voltage range at an operating pressure of 1.6 × 10-5 torr Ar:, discharge ion density and electron temperature range 1.5-9.3 × 1016 m-3 and 5.4-9.9 eV, respectively. Maximum beam current extracted is 120 mA at a 600 V discharge and is primarily limited by ion impingement on the grids due to insufficient grid potentials. The maximum measured thrust is 2.77 mN with an average uncertainty of ±3.6 mN. Further optimization of the thruster operating conditions and grid assembly can improve thruster performance. Copyright © 2013 by the American Institute of Aeronautics and Astronautics, Inc.


Williams L.T.,Georgia Institute of Technology | Williams L.T.,High Power Electrical Propulsion Laboratory | Walker M.L.R.,Georgia Institute of Technology | Walker M.L.R.,High Power Electrical Propulsion Laboratory
Journal of Propulsion and Power | Year: 2013

There is interest in the use of a helicon plasma source in propulsive applications as both an ion source and a thruster. Development of a helicon thruster requires a performance baseline as a basis for future optimization and modification. For the first time, the thrust of a helicon plasma source is measured using a null-type inverted pendulum thrust stand at an operating pressure of 2 × 10-5 torr through the operating range of 215-840WRF power, 11.9 and 13.56 MHz RF frequency, 150-450 G magnetic field strength, and 1.5-4.5 mg/s propellant flow rate for argon. Maximumthrust is found to be 6.3mNat a specific impulse of 140 s and a maximumspecific impulse of 380 s at 5.6 mN. Thrust efficiency is less than 1.4% and demonstrates very-low-power coupling to ion acceleration. Copyright © 2013 by Logan T. Williams.


Langendorf S.,Georgia Institute of Technology | Langendorf S.,High Power Electrical Propulsion Laboratory | Xu K.,University of Alabama in Huntsville | Walker M.,Georgia Institute of Technology | Walker M.,High Power Electrical Propulsion Laboratory
Physics of Plasmas | Year: 2015

This paper investigates the physical mechanisms that cause beneficial and detrimental performance effect observed to date in Hall effect thrusters with wall electrodes. It is determined that the wall electrode sheath can reduce ion losses to the wall if positioned near the anode (outside the dense region of the plasma) such that an ion-repelling sheath is able to form. The ability of the wall electrode to form an ion-repelling sheath is inversely proportional to the current drawn-if the wall electrode becomes the dominant sink for the thruster discharge current, increases in wall electrode bias result in increased local plasma potential rather than an ion-repelling sheath. A single-fluid electron flow model gives results that mimic the observed potential structures and the current-sharing fractions between the anode and wall electrodes, showing that potential gradients in the presheath and bulk plasma come at the expense of current draw to the wall electrodes. Secondary electron emission from the wall electrodes (or lack thereof) is inferred to have a larger effect if the electrodes are positioned near the exit plane than if positioned near the anode, due to the difference in energy deposition from the plasma. © 2015 AIP Publishing LLC.


Xu K.G.,Georgia Institute of Technology | Xu K.G.,High Power Electrical Propulsion Laboratory | Walker M.L.R.,Georgia Institute of Technology | Walker M.L.R.,High Power Electrical Propulsion Laboratory
Journal of Propulsion and Power | Year: 2012

The T-220HT Hall-effect thruster is modified to include in-channel electrodes and additional magnetic coils to study ion focusing. The goal of this work is to decrease energy losses from ion-wall neutralization and plume divergence to increase the thrust-to-power ratio. In this paper, thrust and plume measurements on xenon are presented. The thruster was tested from 125 to 300 V at 9 A discharge, with the electrodes either floating or biased to 10 or 30 V above anode potential. The mass flow rate was varied from 9.8 to 10:4 mg=s to maintain constant discharge current. The maximum operating chamber pressure was 7:7 × 10 -6 Torr-Xe. Performance measurements on xenon show the best overall increase in performance at 150 V discharge and 10 V electrodes with an increase of 7.69 mN of thrust, 4:6 mN=kW thrust-to-power ratio, 123 s I SP, and 5.3% anode efficiency. The plume ion energy distribution function indicates an ion energy increase up to 25 V. The ion number density and propellant efficiency both show increases. The different electrode currents and ion energy distribution functions at 10 V compared with 30 V electrodes suggest different electrode-plasma interaction at the two electrode biases. Copyright © 2012 by Kunning G. Xu. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


Langendorf S.,Georgia Institute of Technology | Langendorf S.,High Power Electrical Propulsion Laboratory | Walker M.,Georgia Institute of Technology | Walker M.,High Power Electrical Propulsion Laboratory
Physics of Plasmas | Year: 2015

In this experiment, plasma sheath potential profiles are measured over boron nitride walls in argon plasma and the effect of secondary electron emission is observed. Results are compared to a kinetic model. Plasmas are generated with a number density of 3 × 1012 m-3 at a pressure of 10-4 Torr-Ar, with a 1%-16% fraction of energetic primary electrons. The sheath potential profile at the surface of each sample is measured with emissive probes. The electron number densities and temperatures are measured in the bulk plasma with a planar Langmuir probe. The plasma is non-Maxwellian, with isotropic and directed energetic electron populations from 50 to 200 eV and hot and cold Maxwellian populations from 3.6 to 6.4 eV and 0.3 to 1.3 eV, respectively. Plasma Debye lengths range from 4 to 7 mm and the ion-neutral mean free path is 0.8 m. Sheath thicknesses range from 20 to 50 mm, with the smaller thickness occurring near the critical secondary electron emission yield of the wall material. Measured floating potentials are within 16% of model predictions. Measured sheath potential profiles agree with model predictions within 5 V (∼1 Te), and in four out of six cases deviate less than the measurement uncertainty of 1 V. © 2015 AIP Publishing LLC.

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