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  • 1
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    Radioisotope electric propulsion (REP): A near-term approach to nuclear propulsion (2010)
    Elsevier
    Details
    Publication Date: 2010-02-01
    Print ISSN: 0094-5765
    Electronic ISSN: 1879-2030
    Topics: Mechanical Engineering, Materials Science, Production Engineering, Mining and Metallurgy, Traffic Engineering, Precision Mechanics
    Published by Elsevier
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    PAPER CURRENT
  • 2
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    High-Power Krypton Hall Thruster Technology Being Developed for Nuclear-Powered Applications (2004)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: The NASA Glenn Research Center has been performing research and development of moderate specific impulse, xenon-fueled, high-power Hall thrusters for potential solar electric propulsion applications. These applications include Mars missions, reusable tugs for low-Earth-orbit to geosynchronous-Earth-orbit transportation, and missions that require transportation to libration points. This research and development effort resulted in the design and fabrication of the NASA-457M Hall thruster that has been tested at input powers up to 95 kW. During project year 2003, NASA established Project Prometheus to develop technology in the areas of nuclear power and propulsion, which are enabling for deep-space science missions. One of the Project-Prometheus-sponsored Nuclear Propulsion Research tasks is to investigate alternate propellants for high-power Hall thruster electric propulsion. The motivation for alternate propellants includes the disadvantageous cost and availability of xenon propellant for extremely large scale, xenon-fueled propulsion systems and the potential system performance benefits of using alternate propellants. The alternate propellant krypton was investigated because of its low cost relative to xenon. Krypton propellant also has potential performance benefits for deep-space missions because the theoretical specific impulse for a given voltage is 20 percent higher than for xenon because of krypton's lower molecular weight. During project year 2003, the performance of the high-power NASA-457M Hall thruster was measured using krypton as the propellant at power levels ranging from 6.4 to 72.5 kW. The thrust produced ranged from 0.3 to 2.5 N at a discharge specific impulse up to 4500 sec.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2003; NASA/TM-2004-212729
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Preliminary plume characteristics of an arcjet thruster (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An experimental program initiated to characterize the near field of an arcjet plume is described. The complete emission spectrum from 3200 to 7200 A at the nozzle exit plane detected the electronically excited species N2, N2(+), NH, and H, indicating excitation, dissociation, ionization, and recombination in the nozzle. Axial intensity profiles indicated an exponential decay in excited state population for H(alpha), H(beta), and NH. The rate of axial decay indicated lower velocities for NH than H in the plume and population of the third excited energy state of hydrogen from the decay of higher energy levels. Rotational temperatures ranged from 750 K for N2 to 2500 K for NH. Based on these results, the arcjet plume is found to be a highly nonequilibrium plasma. Anode electrical configuration is found to have a large effect on the spectral intensities measured in the plume.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 90-2645
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Performance characterization of a segmented anode arcjet thruster (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Results are presented from a continuing experimental program aimed at providing insight into arc energy deposition in the nozzle, the nature of the arc attachment, and its effects on performance characteristics of the device. A modular, 1-2 kW class arcjet thruster incorporating a segmented anode/nozzle was run on a thrust stand to determine performance characteristics under a number of experimental conditions. The nozzle comprised five axial conducting segments isolated from each other by boron nitride spacers. The electrical configuration permitted the current delivered to the arcjet to be collected at any combination of segments. It is concluded that the changes in the electric field in the nozzle that occur as a result of the changes in the current distribution do not significantly affect the momentum transfer or loss mechanisms in the type of nozzle investigated. Performance characteristics show that the segmented anode reasonably simulates the behavior of solid anodes of similar geometry.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 90-2582
    Format: text
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    NASA TECHNICAL REPORTS
  • 5
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    Vacuum ultraviolet absorption in a hydrogen arcjet (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Atomic absorption spectroscopy was utilized to measure the ground state atomic hydrogen number density in the plasma produced in a low power hydrogen arcjet. A microwave driven hydrogen plasma was used as the source of radiation resonant with the vacuum ultraviolet Lyman alpha transition. The suitability of this radiation source is discussed. The optical depth of this transition prevented measurements at locations where the ground state atomic hydrogen number density was larger than 3 x 10 exp 19/cu m. These results indicate that other single-photon optical diagnostic techniques are equally ineffective in locations of higher hydrogen number density unless the spectral line shape of the atomic hydrogen absorbers is known.
    Keywords: PLASMA PHYSICS
    Type: AIAA PAPER 92-3564
    Format: text
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    NASA TECHNICAL REPORTS
  • 6
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    High-Power Hall Thruster Technology Evaluated for Primary Propulsion Applications (2003)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: High-power electric propulsion systems have been shown to be enabling for a number of NASA concepts, including piloted missions to Mars and Earth-orbiting solar electric power generation for terrestrial use (refs. 1 and 2). These types of missions require moderate transfer times and sizable thrust levels, resulting in an optimized propulsion system with greater specific impulse than conventional chemical systems and greater thrust than ion thruster systems. Hall thruster technology will offer a favorable combination of performance, reliability, and lifetime for such applications if input power can be scaled by more than an order of magnitude from the kilowatt level of the current state-of-the-art systems. As a result, the NASA Glenn Research Center conducted strategic technology research and development into high-power Hall thruster technology. During program year 2002, an in-house fabricated thruster, designated the NASA-457M, was experimentally evaluated at input powers up to 72 kW. These tests demonstrated the efficacy of scaling Hall thrusters to high power suitable for a range of future missions. Thrust up to nearly 3 N was measured. Discharge specific impulses ranged from 1750 to 3250 sec, with discharge efficiencies between 46 and 65 percent. This thruster is the highest power, highest thrust Hall thruster ever tested.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 2002; NASA/TM-2003-211990
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    The effect of electrode configuration on arcjet performance (1989)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A segmented anode/nozzle for a low-power (1 kW) arc jet thruster was fabricated to investigate the effect of electrode configuration on the arc jet performance. The five segments of the nozzle, which could be isolated individually or in groups made it possible to observe current distribution from the constrictor through the diverging section of the nozzle; measurements of the potential difference between the cathode and any of the individual segments were possible. Results showed that the discharge initiates in the high pressure region of the nozzle upstream of the diverging section, and then rapidly moves to the diverging section of the nozzle. When the arc was allowed to seat in the diverging section, the anode fall voltage was between 10 to 20 volts. When the current was forced to the high-pressure section of the constrictor, the anode fall voltage increased to more than 40 volts.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 89-2722
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    Stationary Plasma Thruster Ion Velocity Distribution (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A nonintrusive velocity diagnostic based on laser induced fluorescence of the 5d4F(5/2)-6p4D(5/2) singly ionized xenon transition was used to interrogate the exhaust of a 1.5 kW Stationary Plasma Thruster (SPT). A detailed map of plume velocity vectors was obtained using a simplified, cost-effective, nonintrusive, semiconductor laser based scheme. Circumferential velocities on the order of 250 m/s were measured which implied induced momentum torques of approximately 5 x 10(exp -2) N-cm. Axial and radial velocities were evaluated one mm downstream of the cathode at several locations across the width of the annular acceleration channel. Radial velocities varied linearly with radial distance. A maximum radial velocity of 7500 m/s was measured 8 mm from the center of the channel. Axial velocities as large as 16,500 m/s were measured.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-195379 , E-9094 , NAS 1.26:195379 , AIAA PAPER 94-3141 , Joint Propulsion Conference; Jun 27, 1994 - Jun 29, 1994; Indianapolis, IN; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Stationary Plasma Thruster Plume Characteristics (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Stationary Plasma Thrusters (SPT's) are being investigated for application to a variety of near-term missions. This paper presents the results of a preliminary study of the thruster plume characteristics which are needed to assess spacecraft integration requirements. Langmuir probes, planar probes, Faraday cups, and a retarding potential analyzer were used to measure plume properties. For the design operating voltage of 300 V the centerline electron density was found to decrease from approximately 1.8 x 10 exp 17 cubic meters at a distance of 0.3 m to 1.8 X 10 exp 14 cubic meters at a distance of 4 m from the thruster. The electron temperature over the same region was between 1.7 and 3.5 eV. Ion current density measurements showed that the plume was sharply peaked, dropping by a factor of 2.6 within 22 degrees of centerline. The ion energy 4 m from the thruster and 15 degrees off-centerline was approximately 270 V. The thruster cathode flow rate and facility pressure were found to strongly affect the plume properties. In addition to the plume measurements, the data from the various probe types were used to assess the impact of probe design criteria
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-194454 , E-8406 , NAS 1.26:194454 , International Electric Propulsion Conference; Sep 13, 1993 - Sep 16, 1993; Seattle, WA; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    Plasma contactor development for Space Station (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Plasma contactors have been baselined for the Space Station (SS) to control the electrical potentials of surfaces to eliminate/mitigate damaging interactions with the space environment. The system represents a dual-use technology which is a direct outgrowth of the NASA electric propulsion program and, in particular, the technology development effort on ion thrustor systems. The plasma contactor subsystems include the plasma contactor unit, a power electronics unit, and an expellant management unit. Under this pre-flight development program these will all be brought to breadboard or engineering model status. Development efforts for the plasma contactor include optimizing the design and configuration of the contactor, validating its required lifetime, and characterizing the contactor plume and electromagnetic interference. The plasma contactor unit design selected for the SS is an enclosed keeper, xenon hollow cathode plasma source. This paper discusses the test results and development status of the plasma contactor unit subsystem for the SS.
    Keywords: PLASMA PHYSICS
    Type: NASA-TM-106425 , IEPC-93-246 , E-8267 , NAS 1.15:106425 , International Electric Propulsion Conference; Sep 13, 1993 - Sep 16, 1993; Seattle, WA; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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