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  • 1
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    The discharge plasma in ion engine neutralizers: Numerical simulations and comparisons with laboratory data (2010)
    American Institute of Physics
    Details
    Publication Date: 2010-12-01
    Print ISSN: 0021-8979
    Electronic ISSN: 1089-7550
    Topics: Physics
    Published by American Institute of Physics
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    PAPER CURRENT
  • 2
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    An ion thruster internal discharge chamber electrostatic probe diagnostic technique using a high-speed probe positioning system (2008)
    American Institute of Physics
    Details
    Publication Date: 2008-01-01
    Print ISSN: 0034-6748
    Electronic ISSN: 1089-7623
    Topics: Electrical Engineering, Measurement and Control Technology , Physics
    Published by American Institute of Physics
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    PAPER CURRENT
  • 3
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    Wear Trends of the HERMeS Thruster as a Function of Throttle Point (2019)
    In:  CASI
    Details
    Publication Date: 2019-07-20
    Description: A series of short-duration (200 h) wear tests were conducted with two Hall Effect Rocket with Magnetic Shielding (HERMeS) technology demonstration units. Front pole covers, cathode keeper, and discharge channel wear were characterized as a function of discharge voltage, magnetic field strength, and chamber pressure. No discharge channel erosion was observed. Inner pole cover erosion was shown to be a weak function of discharge voltage with most erosion occurring at the lowest value, 300 V. The Technology Demonstration Unit (TDU) 3 keeper electrode eroded with each operating condition, with high magnetic field yielding the greatest erosion rate. The TDU-1 keeper electrode exhibited net deposition suggesting its configuration is more consistent with meeting overall HERMeS service life requirements. Ratios of molybdenum to graphite erosion rates suggests, with high uncertainty, that the sputtering ions are originating downstream of the thruster exit plane, striking the surface with small angles of incidence.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2019-219731 , IEPC?2017?207 , E-19456 , GRC-E-DAA-TN48801 , International Electric Propulsion Conference; Oct 08, 2017 - Oct 12, 2017; Atlanta, GA; United States
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    NASA TECHNICAL REPORTS
  • 4
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    End-of-Test Performance and Wear Characterization of NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test (2014)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The NASA's Evolutionary Xenon Thruster (NEXT) program is developing the next-generation solar electric ion propulsion system with significant enhancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to quantify the thruster propellant throughput capability. Testing was recently completed in February 2014, with the thruster accumulating 51,184 hours of operation, processing 918 kg of xenon propellant, and delivering 35.5 MN-s of total impulse.As part of the test termination procedure, a comprehensive performance characterization was performed across the entire NEXT throttle table. This was performed prior to planned repairs of numerous diagnostics that had become inoperable over the course of the test. After completion of these diagnostic repairs in November 2013, a comprehensive end-of-test performance and wear characterization was performed on the test article prior to exposure to atmosphere. These data have confirmed steady thruster performance with minimal degradation as well as mitigation of numerous life limiting mechanisms encountered in the NSTAR design. Component erosion rates compare favorably to pretest predictions based on semi-empirical models used for the thruster service life assessment. Additional data relating to ion beam density profiles, facility backsputter rates, facility backpressure effects on thruster telemetry, and modulation of the neutralizer keeper current are presented as part of the end-of-test characterization. Presently the test article for the NEXT LDT has been exposed to atmosphere and placed within a clean room environment, with post-test disassembly and inspection underway.
    Keywords: Spacecraft Propulsion and Power
    Type: GRC-E-DAA-TN16314 , AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 28, 2014 - Jul 30, 2014; Cleveland, OH; United States
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    NASA TECHNICAL REPORTS
  • 5
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    A New Method for Analyzing Near-Field Faraday Probe Data in Hall Thrusters (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents a new method for analyzing near-field Faraday probe data obtained from Hall thrusters. Traditional methods spawned from far-field Faraday probe analysis rely on assumptions that are not applicable to near-field Faraday probe data. In particular, arbitrary choices for the point of origin and limits of integration have made interpretation of the results difficult. The new method, called iterative pathfinding, uses the evolution of the near-field plume with distance to provide feedback for determining the location of the point of origin. Although still susceptible to the choice of integration limits, this method presents a systematic approach to determining the origin point for calculating the divergence angle. The iterative pathfinding method is applied to near-field Faraday probe data taken in a previous study from the NASA-300M and NASA-457Mv2 Hall thrusters. Since these two thrusters use centrally mounted cathodes the current density associated with the cathode plume is removed before applying iterative pathfinding. A procedure is presented for removing the cathode plume. The results of the analysis are compared to far-field probe analysis results. This paper ends with checks on the validity of the new method and discussions on the implications of the results.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2013-216567 , AIAA Paper-2013-4118 , E-18755 , NASA-GRC-TN10191 , IECEC; Jul 14, 2013 - Jul 17, 2013; San Jose, CA; United States|AIAA/ASME/ASEE JPC; Jul 14, 2013 - Jul 17, 2013; San Jose, CA; United States
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    NASA TECHNICAL REPORTS
  • 6
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    A New Method for Analyzing Near-Field Faraday Probe Data in Hall Thrusters (2013)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents a new method for analyzing near-field Faraday probe data obtained from Hall thrusters. Traditional methods spawned from far-field Faraday probe analysis rely on assumptions that are not applicable to near-field Faraday probe data. In particular, arbitrary choices for the point of origin and limits of integration have made interpretation of the results difficult. The new method, called iterative pathfinding, uses the evolution of the near-field plume with distance to provide feedback for determining the location of the point of origin. Although still susceptible to the choice of integration limits, this method presents a systematic approach to determining the origin point for calculating the divergence angle. The iterative pathfinding method is applied to near-field Faraday probe data taken in a previous study from the NASA-300M and NASA-457Mv2 Hall thrusters. Since these two thrusters use centrally mounted cathodes, the current density associated with the cathode plume is removed before applying iterative pathfinding. A procedure is presented for removing the cathode plume. The results of the analysis are compared to far-field probe analysis results. This paper ends with checks on the validity of the new method and discussions on the implications of the results.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2013-216567 , E-18755 , GRC-E-DAA-TN10191 , AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 15, 2013 - Jul 17, 2013; San Jose, CA; United States|International Energy Conversion Engineering Conference; Jul 15, 2013 - Jul 17, 2013; San Jose, CA; United States
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    NASA TECHNICAL REPORTS
  • 7
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    The Ion Propulsion System for the Solar Electric Propulsion Technology Demonstration Mission (2015)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Asteroid Redirect Robotic Mission is a candidate Solar Electric Propulsion Technology Demonstration Mission whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a subsequent human-crewed mission. The ion propulsion subsystem must be capable of operating over an 8-year time period and processing up to 10,000 kg of xenon propellant. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as an enabling element of an affordable beyond low-earth orbit human-crewed exploration architecture. Under the NASA Space Technology Mission Directorate the critical electric propulsion and solar array technologies are being developed. The ion propulsion system for the Asteroid Redirect Vehicle is based on the NASA-developed 12.5 kW Hall Effect Rocket with Magnetic Shielding thruster and power processing technologies. This paper presents the conceptual design for the ion propulsion system, a status on the NASA in-house thruster and power processing is provided, and an update on acquisition for flight provided.
    Keywords: Spacecraft Propulsion and Power
    Type: IEPC-2015-008 , ISTS-2015-b-008 , GRC-E-DAA-TN24790 , International Symposium on Space Technology and Science; Jul 04, 2015 - Jul 10, 2015; Kobe-Hyogo; Japan|Nano-Satellite Symposium; Jul 04, 2015 - Jul 10, 2015; Kobe-Hyogo; Japan|International Electric Propulsion Conference (IEPC); Jul 04, 2015 - Jul 10, 2015; Kobe-Hyogo; Japan
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    NASA TECHNICAL REPORTS
  • 8
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    The Ion Propulsion System for the Solar Electric Propulsion Technology Demonstration Mission (2015)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Asteroid Redirect Robotic Mission is a candidate Solar Electric Propulsion Technology Demonstration Mission whose main objectives are to develop and demonstrate a high-power solar electric propulsion capability for the Agency and return an asteroidal mass for rendezvous and characterization in a companion human-crewed mission. The ion propulsion system must be capable of operating over an 8-year time period and processing up to 10,000 kg of xenon propellant. This high-power solar electric propulsion capability, or an extensible derivative of it, has been identified as a critical part of an affordable, beyond-low-Earth-orbit, manned-exploration architecture. Under the NASA Space Technology Mission Directorate the critical electric propulsion and solar array technologies are being developed. The ion propulsion system being co-developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory for the Asteroid Redirect Vehicle is based on the NASA-developed 12.5 kW Hall Effect Rocket with Magnetic Shielding (HERMeS0 thruster and power processing technologies. This paper presents the conceptual design for the ion propulsion system, the status of the NASA in-house thruster and power processing activity, and an update on flight hardware.
    Keywords: Spacecraft Propulsion and Power
    Type: IEPC-2015-008 /ISTS-2015-b-008 , GRC-E-DAA-TN24654 , Nano-satellite Symposium (NSAT); Jul 04, 2015 - Jul 10, 2015; Kobe-Hyogo; Japan|International Electric Propulsion Conference (IEPC); Jul 04, 2015 - Jul 10, 2015; Kobe-Hyogo; Japan|International Symposium on Space Technology and Science (ISTS); Jul 04, 2015 - Jul 10, 2015; Kobe-Hyogo; Japan
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    NASA TECHNICAL REPORTS
  • 9
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    Plasma Potential and Langmuir Probe Measurements in the Near-field Plume of the NASA-457Mv2 Hall Thruster (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In order to further the design of future high-power Hall thrusters and provide experimental validation for ongoing modeling efforts, plasma potential and Langmuir probe measurements were performed on the 50-kW NASA-457Mv2. An electrostatic probe array comprised of a near-field Faraday probe, single Langmuir probe, and emissive probe was used to interrogate the near-field plume from approximately 0.1 - 2.0 mean thruster diameters downstream of the thruster exit plane at the following operating conditions: 300 V, 400 V and 500 V at 30 kW and 500 V at 50 kW. Results have shown that the acceleration zone is limited to within 0.4 mean thruster diameters of the exit plane while the high-temperature region is limited to 0.25 mean thruster diameters from the exit plane at all four operating conditions. Maximum plasma potentials in the near-field at 300 and 400 V were approximately 50 V with respect to cathode potential, while maximum electron temperatures varied from 24 - 32 eV, depending on operating condition. Isothermal lines at all operating conditions were found to strongly resemble the magnetic field topology in the high-temperature regions. This distribution was found to create regions of high temperature and low density near the magnetic poles, indicating strong, thick sheath formation along these surfaces. The data taken from this study are considered valuable for future design as well as modeling validation.
    Keywords: Spacecraft Propulsion and Power
    Type: E-18469-1 , 48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 30, 2012 - Aug 01, 2012; Atlanta, GA; United States
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    NASA TECHNICAL REPORTS
  • 10
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    Plasma Potential and Langmuir Probe Measurements in the Near-field Plume of the NASA-300M Hall Thruster (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In order to aid in the design of high-power Hall thrusters and provide experimental validation for existing modeling efforts, plasma potential and Langmuir probe measurements were performed in the near-field plume of the NASA-300M Hall thruster. A probe array consisting of a Faraday probe, Langmuir probe, and emissive probe was used to interrogate the plume from approximately 0.1 - 2.0 mean thruster diameters downstream of the thruster exit plane at four operating conditions: 300 V, 400 V, and 500 V at 20 kW as well as 300 V at 10 kW. Results show that the acceleration zone and high-temperature region were contained within 0.3 mean thruster diameters from the exit plane at all operating conditions. Isothermal lines were shown to strongly follow magnetic field lines in the near-field, with maximum temperatures ranging from 19 - 27 eV. The electron temperature spatial distribution created large drops in measured floating potentials in front of the magnetic pole surfaces where the plasma density was low, which suggests strong sheaths at these surfaces. The data taken have provided valuable information for future design and modeling validation, and complements ongoing internal measurement efforts on the NASA-300M.
    Keywords: Instrumentation and Photography
    Type: AIAA Paper 2012-4115 , E-18459 , 48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 29, 2012 - Aug 01, 2012; Atlanta, GA; United States
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