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  • 1
    Electronic Resource
    Electronic Resource
    Electron-beam plasma enhancement of oxidation in thin aluminum samples (1990)
    [S.l.] : American Institute of Physics (AIP)
    Journal of Applied Physics 67 (1990), S. 1601-1602 
    Details
    ISSN: 1089-7550
    Source: AIP Digital Archive
    Topics: Physics
    Notes: In order to gain an insight into the effects of a space environment on materials, thin aluminum samples were exposed to an oxygen plasma produced by an electron beam. Using Rutherford backscattering spectrometry the samples exposed to the oxygen plasma were compared to samples exposed to ordinary oxygen gas. The comparison revealed a considerable increase in oxygen diffusion in the samples exposed to the charged particle environment. The amount of oxygen diffusion depends on the duration of the plasma exposure.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1063/1.345645
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    Paper (German National Licenses)
    Fulltext
  • 2
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    Electron‐beam plasma enhancement of oxidation in thin aluminum samples (1990)
    American Institute of Physics
    Details
    Publication Date: 1990-02-01
    Print ISSN: 0021-8979
    Electronic ISSN: 1089-7550
    Topics: Physics
    Published by American Institute of Physics
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    PAPER CURRENT
  • 3
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    Measurements of Drift Surfaces in Compact Auburn Torsatron (1995)
    Institute of Physics
    Details
    Publication Date: 1995-02-15
    Print ISSN: 0021-4922
    Electronic ISSN: 1347-4065
    Topics: Physics
    Published by Institute of Physics
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    PAPER CURRENT
  • 4
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    Ground-Based Simulation of Low-Earth Orbit Plasma Conditions: Plasma Generation and Characterization (2004)
    In:  CASI
    Details
    Publication Date: 2018-06-12
    Description: A 16-cm diameter plasma source operated on argon is described that is capable of producing a plasma environment that closely simulates the low Earth orbit (LEO) conditions experienced by satellites in the altitude range between 300 to 500 km. The plasma source uses a transverse-field magnetic filter, and has been successful in producing low electron temperature plasmas that contain streaming ion populations. Both of these characteristics are important because the plasma in LEO is relatively cold (e.g., Te approx. 0.1 eV) and the ram energy of the ions due to the motion of the satellite relative to the LEO plasma is high (e.g., 7,800 m/s which corresponds to approx. 5 eV for O+ ions). Plasma source operational conditions of flow rate and discharge power are presented that allow the electron temperature to be adjusted over a range from 0.14 to 0.4 eV. The expanding plasma flow field downstream of the source contains both low-energy, charge-exchange ions and streaming ions with energies that are adjustable over a range from 4 eV to 6 eV. At low flow rates and low facility pressures, the streaming ion component of the ion population comprises over 90% of the total plasma density. In the work described herein, a large area retarding potential analyzer was used to measure both electron and ion energy distribution functions in the low density, expanding plasma produced downstream of the plasma source. The benefits of using this type of plasma diagnostic tool in easily perturbed, low-density plasma are identified, and techniques are also discussed that can be used to perform real-time measurements of electron temperature. Finally, recommendations are made that may enable lower electron temperatures to be produced while simultaneously decreasing the plasma source flow rate below 1 to 2 sccm.
    Keywords: Plasma Physics
    Type: 8th Spacecraft Charging Technology Conference; NASA/CP-2004-213091
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    High Voltage Solar Array Arc Testing for a Direct Drive Hall Effect Thruster System (2004)
    In:  CASI
    Details
    Publication Date: 2018-06-12
    Description: The deleterious effects of spacecraft charging are well known, particularly when the charging leads to arc events. The damage that results from arcing can severely reduce system lifetime and even cause critical system failures. On a primary spacecraft system such as a solar array, there is very little tolerance for arcing. Motivated by these concerns, an experimental investigation was undertaken to determine arc thresholds for a high voltage (200-500 V) solar array in a plasma environment. The investigation was in support of a NASA program to develop a Direct Drive Hall-Effect Thruster (D2HET) system. By directly coupling the solar array to a Hall-effect thruster, the D2HET program seeks to reduce mass, cost and complexity commonly associated with the power processing in conventional power systems. In the investigation, multiple solar array technologies and configurations were tested. The cell samples were biased to a negative voltage, with an applied potential difference between them, to imitate possible scenarios in solar array strings that could lead to damaging arcs. The samples were tested in an environment that emulated a low-energy, HET-induced plasma. Short duration trigger arcs as well as long duration sustained arcs were generated. Typical current and voltage waveforms associated with the arc events are presented. Arc thresholds are also defined in terms of voltage, current and power. The data will be used to propose a new, high-voltage (greater than 300 V) solar array design for which the likelihood of damage from arcing is minimal.
    Keywords: Spacecraft Design, Testing and Performance
    Type: 8th Spacecraft Charging Technology Conference; NASA/CP-2004-213091
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Minimum Arc Threshold Voltage Experiments on Extravehicular Mobility Unit Samples (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: The International Space Station (ISS) is now under construction in Low Earth Orbit (LEO). The process of building the ISS requires that astronauts carry out many Extravehicular Activities. To protect the astronauts form the hazardous space environment, they are required to wear a suit known as the Extravehicular Mobility Unit (EMU). For most Extra-Vehicular Activities (EVAs) the EMU is tethered to ISS via a steel safety tether. During the course of an EVA it is common for the safety tether to contact exposed metal on both the ISS and the EMU. In this case, the single point ground of the EMU would be at the same potential as the ISS with respect to the LEO Plasma. In the event that the metal structure of the ISS begins to charge negative of the plasma potential as a result of electron collection by the ISS photovoltaic arrays, then the EMU would also be driven to a negative potential. Anodized aluminum components on the EMU would then begin to develop a charge across their amortization layer as ions from the plasma are collected. In the case where large negative potentials are applied to the EMU, dielectric breakdown may occur as a large voltage difference is developed across the thin amortization layer (oxide). The resulting arc plasma may in turn couple to the charge accumulated on the nearby ISS anodized debris shields and thereby generate a large current flow through the metal EMU structure. Current flow through the EMU could result in an electrocution hazard for the Crew Member inside the EMU - and therefore represents an important safety concern. To address this concern, a series of experiments have been undertaken. In each experiment specially prepared anodized aluminum samples were placed in a LEO representative plasma and charged until dielectric breakdown occurred in the form of an arc. This process was repeated a number of times for three sets of samples. During each test the arc voltage and current were monitored. A statistical treatment of the arc voltage threshold will be presented. In addition, safe operating voltages for the EMU are suggested.
    Keywords: Spacecraft Design, Testing and Performance
    Type: 40th AIAA Aerospace Sciences Meeting and Exhibit; Jan 13, 2002 - Jan 17, 2002; Reno, NV; United States
    Format: text
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    NASA TECHNICAL REPORTS
  • 7
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    Life Testing of the Hollow Cathode Plasma Contactor for the ProSEDS Mission (2001)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: The Propulsive Small Expendable Deployer System (ProSEDS) mission is designed to provide an on-orbit demonstration of the electrodynamic propulsion capabilities of tethers in space. The ProSEDS experiment will be a secondary payload on a Delta 11 unmanned expendable booster. A 5-km conductive tether is attached to the Delta 11 second stage and collects current from the low Earth orbit (LEO) plasma. A hollow cathode plasma contactor emits the collected electrons from the Delta II, completing the electrical circuit with the ambient plasma. The current flowing through the tether generates thrust based on the Lorentz Force Law. The thrust will be generated opposite to the velocity vector, slowing down the spacecraft and causing it to de-orbit in approximately 14 days compared to the normal 6 months. A 10-km non-conductive tether is between the conductive tether and an endmass containing several scientific instruments. The ProSEDS mission lifetime was set at I day because most of the primary objectives can be met in that time. The extended ProSEDS mission will be for as many days as possible, until the Delta 11 second stage burns up or the tether is severed by a micrometeoroid or space debris particle. The Hollow Cathode Plasma Contactor (HCPC) unit has been designed for a 12-day mission. Because of the science requirements to measure the background ambient plasma, the HCPC must operate on a duty cycle. Later in the ProSEDS mission, the HCPC is operated in a manner to allow charging of the secondary battery. Due to the unusual operating requirements by the ProSEDS mission, a development unit of the HCPC was built for thorough testing. This developmental unit was tested for a simulated ProSEDS mission, with measurements of the ability to start and stop during the duty cycle. These tests also provided valuable data for the ProSEDS software requirements. Qualification tests of the HCPC flight hardware are also discussed.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Spacecraft Charging Technology Conference; Apr 23, 2001 - Apr 27, 2001; Noordwijk; Netherlands
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    Overview of Advanced Space Propulsion Activities in the Space Environmental Effects Team at MSFC (2000)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: Exploration of our solar system, and beyond, requires spacecraft velocities beyond our current technological level. Technologies addressing this limitation are numerous. The Space Environmental Effects (SEE) Team at the Marshall Space Flight Center (MSFC) is focused on three discipline areas of advanced propulsion; Tethers, Beamed Energy, and Plasma. This presentation will give an overview of advanced propulsion related activities in the Space Environmental Effects Team at MSFC. Advancements in the application of tethers for spacecraft propulsion were made while developing the Propulsive Small Expendable Deployer System (ProSEDS). New tether materials were developed to meet the specifications of the ProSEDS mission and new techniques had to be developed to test and characterize these tethers. Plasma contactors were developed, tested and modified to meet new requirements. Follow-on activities in tether propulsion include the Air-SEDS activity. Beamed energy activities initiated with an experimental investigation to quantify the momentum transfer subsequent to high power, 5J, ablative laser interaction with materials. The next step with this experimental investigation is to quantify non-ablative photon momentum transfer. This step was started last year and will be used to characterize the efficiency of solar sail materials before and after exposure to Space Environmental Effects (SEE). Our focus with plasma, for propulsion, concentrates on optimizing energy deposition into a magnetically confined plasma and integration of measurement techniques for determining plasma parameters. Plasma confinement is accomplished with the Marshall Magnetic Mirror (M3) device. Initial energy coupling experiments will consist of injecting a 50 amp electron beam into a target plasma. Measurements of plasma temperature and density will be used to determine the effect of changes in magnetic field structure, beam current, and gas species. Experimental observations will be compared to predictions from computer modeling.
    Keywords: Spacecraft Propulsion and Power
    Type: 11th Advanced Propulsion Workshop; May 31, 2000 - Jun 02, 2000; Pasadena, CA; United States
    Format: text
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    NASA TECHNICAL REPORTS
  • 9
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    Internal Electrostatic Discharge Testing of Inert Solid Rocket Motor Materials (2019)
    In:  CASI
    Details
    Publication Date: 2019-07-19
    Description: No abstract available
    Keywords: Space Processing
    Type: M19-7359 , Applied Space Environments Conference (ASEC); May 12, 2019 - May 17, 2019; Los Angeles, CA; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    NEXIS Reservoir Cathode 2000 Hour Life Test (2004)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: The current design of the Nuclear Electric Xenon Ion System (NEXIS) employs a reservoir cathode as both the discharge and neutralizer cathode to meet the 10 yr thruster design life. The main difference between a reservoir cathode and a conventional discharge cathode is the source material (barium-containing compound) is contained within a reservoir instead of in an impregnated insert in the hollow tube. However, reservoir cathodes do not have much life test history associated with them. In order to demonstrate the feasibility of using a reservoir cathode as an integral part of the NEXIS ion thruster, a 2000 hr life test was performed. Several proof-of-concept (POC) reservoir cathodes were built early in the NEXIS program to conduct performance testing as well as life tests. One of the POC cathodes was sent to Marshall Space Flight Center (MSFC) where it was tested for 2000 hrs in a vacuum chamber. The cathode was operated at the NEXIS design point of 25 A discharge current and a xenon flow rate of 5.5 sccm during the 2000 hr test. The cathode performance parameters, including discharge current, discharge voltage, keeper current; keeper voltage, and flow rate were monitored throughout test. Also, the temperature upstream of cathode heater, the temperature downstream of the cathode heater, and the temperature of the orifice plate were monitored throughout the life of the test. The results of the 2000 hr test will be described in this paper. Included in the results will be time history of discharge current, discharge voltage, and flow rate. Also, a time history of the cathode temperature will be provided.
    Keywords: Spacecraft Propulsion and Power
    Type: 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 11, 2004 - Jul 14, 2004; Fort Lauderdale, FL; United States
    Format: text
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    NASA TECHNICAL REPORTS
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