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第12回 宇宙環境シンポジウム（2015年11月16日-18日. 北九州国際会議場 国際会議室）, 北九州市, 福岡県, The 12th Spacecraft Environment Symposium (November 16-18, 2015. International Conference Room, Kitakyushu International Conference Center), Kitakyushu, Fukuoka, Japan, 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA1630004001, レポート番号: JAXA-SP-15-012

Primary discharge occurs on solar arrays due to their interaction with the space plasma. A solar cell may suffer degradation of electrical performance if the primary discharge occurs at the cell edge. To estimate the power generated at the end of life, it is necessary to study the details of solar cell degradation. However, throughout the world, primary discharge has not been recognized as a cause of solar cell degradation. There is now an international collaboration among institutions in Japan, France, and the United States toward a common international standardization of solar array electrostatic discharge test methods. Round-robin tests were carried out as part of this collaborative research. Laboratory experiments were performed at the same time in three institutions using the same test method and identical solar cells. Solar cell degradation was confirmed at all three institutions. It was found that a multijunction solar cell is more susceptible to damage from primary discharge than a crystalline silicon solar cell. Throughout the round-robin tests, discharge has been shown to be a significant cause of solar cell degradation.

第6回宇宙環境シンポジウム (2009年2月29日-30日. 北九州国際会議場), 6th Spacecraft Enivironment Symposium (February 29-30, 2009. Kitakyushu International Conference Center), 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0064542009, レポート番号: JAXA-SP-09-006

第6回宇宙環境シンポジウム (2009年2月29日-30日. 北九州国際会議場), 6th Spacecraft Enivironment Symposium (February 29-30, 2009. Kitakyushu International Conference Center), Most calculations of internal dielectric charging on spacecraft use tabulated values of material surface and bulk conductivities, dielectric constants, and dielectric breakdown strengths. Many of these properties are functions of temperature, and the temperature dependences are not well known. At cryogenic temperatures, where it is well known that material conductivities decrease dramatically, it is an open question as to the timescales over which buried charge will dissipate and prevent the eventual potentially disastrous discharges of dielectrics. In this paper, measurements of dielectric charging and discharging for cable insulation materials at cryogenic temperatures (ca. 90 K) are presented using a broad spectrum electron source at the NASA Marshall Space Flight Center. The measurements were performed for the James Webb Space Telescope (JWST), which will orbit at the Earth-Sun L2 point, and parts of which will be perennially at temperatures as low as 40 K. Results of these measurements seem to show that Radiation Induced Conductivity (RIC) under cryogenic conditions at L2 will not be sufficient to allow charges to bleed off of some typical cable insulation materials even over the projected JWST lifetime of a dozen years or more. After the charging and discharging measurements are presented, comparisons are made between the material conductivities that can be inferred from the measured discharges and conductivities calculated from widely used formulae. Furthermore, the measurement-inferred conductivities are compared with extrapolations of recent measurements of materials RIC and dark conductivities performed with the charge-storage method at Utah State University. Implications of the present measurements are also given for other spacecraft that may operate at cryogenic temperatures, such as probes of the outer planets or the permanently dark cratered areas on the moon. The present results will also be of interest to those who must design or operate spacecraft in more moderate cold conditions. Finally, techniques involving shielding and/or selective use of somewhat conductive insulators are presented to prevent arc-inducing charge buildup even under cryogenic conditions., 形態: カラー図版あり, Physical characteristics: Original contains color illustrations, 資料番号: AA0064542005, レポート番号: JAXA-SP-09-006

The conventional wisdom about solar array arcing in LEO is that only the parts of the solar array that are swept over by the arc-generated plasma front are discharged in the initial arc. This limits the amount of energy that can be discharged. Recent work done at the NASA Glenn Research Center has shown that this idea is mistaken. In fact the capacitance of the entire solar array may be discharged, which for large arrays leads to very large and possibly debilitating arcs, even if no sustained arc occurs. We present the laboratory work that conclusively demonstrates this fact by using a grounded plate that prevents the arc-plasma front from reaching certain array strings. Finally, we discuss the dependence of arc strength and arc pulse width on the capacitance that is discharged, and provide a physical mechanism for discharge of the entire array, even when parts of the array are not accessible to the arc plasma front. Mitigation techniques are also presented., 資料番号: AA0049206006, レポート番号: JAXA-SP-05-001E

Tests of arcing and current collection in simulated space plasma conditions have been performed at the NASA Glenn Research Center (GRC) in Cleveland, Ohio, for over 30 years and at the Marshall Space Flight Center (MSFC) in Huntsville, Alabama, for almost as long. During this period, proper test conditions for accurate and meaningful space simulation have been worked out, comparisons with actual space performance in space flight tests and with real operational satellites have been made, and NASA has achieved our own internal standards for test protocols. It is the purpose of this paper to communicate the test conditions, test procedures, and types of analysis used at NASA GRC and MSFC to the space environmental testing community at large, to help with international space-plasma arcing-testing standardization. To be discussed are: 1. Neutral pressures, neutral gases, and vacuum chamber sizes. 2. Electron and ion densities, plasma uniformity, sample sizes, and Debye lengths. 3. Biasing samples versus self-generated voltages. Floating samples versus grounded. 4. Power supplies and current limits. Isolation of samples from power supplies during arcs. 5. Arc circuits. Capacitance during biased arc-threshold tests. Capacitance during sustained arcing and damage tests. Arc detection. Preventing sustained discharges during testing. 6. Real array or structure samples versus idealized samples. 7. Validity of LEO tests for GEO samples. 8. Extracting arc threshold information from arc rate versus voltage tests. 9. Snapover and current collection at positive sample bias. Glows at positive bias. Kapton pyrolysis. 10. Trigger arc thresholds. Sustained arc thresholds. Paschen discharge during sustained arcing. 11. Testing for Paschen discharge thresholds. Testing for dielectric breakdown thresholds. Testing for tether arcing. 12. Testing in very dense plasmas (i.e., thruster plumes). 13. Arc mitigation strategies. Charging mitigation strategies. Models 14. Analysis of test results. Finally, the necessity of testing will be emphasized, not to the exclusion of modeling, but as part of a complete strategy for determining when and if arcs will occur, and preventing them from occurring in space., 資料番号: AA0049206044, レポート番号: JAXA-SP-05-001E

This paper describes the use of the Floating Potential Probe (FPP) as an “Attached Payload” on ISS. Background and motivation for building the FPP and well as detailed descriptions of its subsystems are described in another paper published in these proceedings (ref # SSC01-V-4b). With it’s solar arrays, primary/secondary power system, control/data processor unit, RF command/data link, thermal protection system, and two science instruments, the FPP displays most of the characteristics of a small spacecraft— with the exception of attitude control and propulsion subsystems. The FPP was attached to the top of the P6 truss during one of several Flight 4A EVAs. It uses an RF link to communicate with an antenna (deployed at the same time as the probe) which feeds though the module and into a transmitter/receiver and portable computer inside the habitable volume. Real time data on the ISS potential is displayed on the laptop and downlinked through the ISS server when requested. This paper will provide an overview of the major subsystems, discuss how such small satellites could be made to work within the ISS system, and the possibilities of using small satellites as attached payloads for short term science or technology experiments. We will provide insight into deployment and operational considerations, show examples of the use of such a low cost system, and discuss briefly the data and science impact of this small $1M class probe.