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資料番号: AA0063997007, レポート番号: JAXA-SP-07-030

There is high possibility that space debris impacts to a solar array paddle in spacecraft parts, because the solar array paddle has large area. Space debris impact to the solar array causes not only mechanical damage but also electrical damage such as arcing on the solar array through local high density plasma created by hypervelocity impact. In the worst case, Joule heating of this arcing carbonizes insulation layer and permanent short-circuit path is created. This is permanent sustained arc. However, no permanent sustained arc caused by space debris impact in orbit has been reported. Purpose of this study is evaluation of possibility of permanent sustained arc through the plasma created by debris impact. Hypervelocity impact tests to solar array coupons in the condition of pseudo power generation were conducted. We ascertained that space debris impact can lead to permanent sustained arc on the solar array due to plasma created by hypervelocity impact., 資料番号: AA0063997018, レポート番号: JAXA-SP-07-030

近年、衛星の帯電・放電に関連すると考えられる事故が国内外で多数、報告されることから、設計段階から帯電・放電の観点に立った検討を正しく行う必要性が認識されている。JAXAでは静止衛星についてはこれまで、米国で70年代に開発され世界のデファクトスタンダードとしての地位を確立していたNASCAP/GEO(NASA Charging Analyzer Program for GEosynchronous Orbit)と呼ばれる解析ツールを利用してきた。しかしその後、機能拡張、改良され、低軌道衛星、極軌道衛星にも対応可能となったツールは、輸出規制のために日本で使用することはできない。そこで2004年11月から衛星帯電解析ソフトウェアMUSCAT(Multi-Utility Spacecraft Charging Analysis Tool)の開発に着手し、2007年3月に完成した。本発表は、このMUSCATによる解析結果を実際の衛星(ETS-5: Engineering Test Satellite-5:きく5号)で取得したデータと比較し報告する。ETS-5は、静止軌道上で10年間の帯電データを取得しており、ほぼ太陽活動周期1回分のデータが得られている。このデータをMUSCATの解析結果と比較することにより、衛星電位と帯電サンプルとの電位差について検証することが可能であり、成果を帯電設計基準などに反映することができる。また、MUSCATにユーザ要求による機能の追加を行ったため、これについても報告する。, 資料番号: AA0063997026, レポート番号: JAXA-SP-07-030

A new method to simulate power generating solar arrays of spacecraft is presented. By combining a multijunction solar array, InGaP-GaAs-Ge, and a Metal Halide Lamp (MHL) secondary arcs were generated. The aim was that the arcs should be comparable to those generated by the currently used methods, which are the Solar Array Simulator (SAS) or a power supply combined with a Current Regulating Diode (CRD) circuit. The results of the experiment showed a successful generation of secondary arcs, especially of temporary sustained arcs. The arc duration is similar for all three simulation methods. The conclusion therefore is that the MHL method can be used in the same way the SAS or CRD method can be used, as it generates the same results., 資料番号: AA0063997012, レポート番号: JAXA-SP-07-030

The purpose of this research work is to develop a small size measurement system to observe the surface charge distribution on insulator in vacuum chamber for space environmental test. Recently, a lot of satellites many purposes like a telecommunication, a weather report are being launched. These satellites are flying in a very severe space environment where the high energy radiations are scattered from the sun. In this environment, electrical charging phenomenon occurs on the insulation of spacecraft surface and electrical discharge or dielectric breakdown occurs by the high electric field. Therefore devices in spacecraft have a risk of failure. So we need to find out a mechanism for the occurrence of surface discharge on insulation and a best insulating materials for space conditions. Until now, my senior at our laboratory has made a measuring system of surface charging on insulation using the BSO (Bismuth Silicon Oxide) crystal has Pockels effect. The measuring system has been used the observation of surface discharging on insulation under conditions of non-destructive and non-contact. However this measuring system is large size and the techniques for setting an optical axis is hard. So it was difficult to apply the measuring system to the space environment simulator of small vacuum chamber. So we need to make a small size measurement system for surface discharging observation. By using a plate type light source of LED (Light Emitting Diode) and changing a optical type, the system area can be reduced 200 x 150 mm from 1,200 x 700 mm. Then we researched surface charging on BSO crystal and Kapton film, by using the small system. Then we researched measuring accuracy of the small system., 資料番号: AA0063997024, レポート番号: JAXA-SP-07-030

In this paper, we report about the test results on coupons made for Working Group 1 (WG1), which was established to make a Japanese guideline for spacecraft charging design. Laboratory tests were carried out with an external circuit simulating a spacecraft power system. The solar array coupon panels simulating the hot and return ends of a string circuit were tested under various combinations of string voltage and string current. We revealed that the threshold conditions for sustained arc were different in string voltage., 資料番号: AA0063997011, レポート番号: JAXA-SP-07-030

The spacecraft like a communication or a broadcasting satellite flying in GEO (Geostationary Earth Orbit) is always exposed to plasma and/or radioactive-rays such as alpha-, beta- and gamma-rays. When the dielectric materials, in which the spacecraft is wrapped to keep temperature in it stable, are irradiated by the high energy electron beam, sometimes an unexpected accident due to an electrostatic discharge happens. The accident sometimes causes a serious mission error of the spacecraft. However, there remain many unknown factor about a charge accumulation in dielectric materials by irradiation of electron beam. Therefore, we need to measure the charge distribution in the bulk of dielectric materials. We have been developing a system for measuring such a charge distribution in dielectric materials using pressure wave propagation method. Using this system, we tried to measure the charge distribution under electron beam irradiated Low Density Polyethylene, Polyethylene Terephthalate, Polycarbonate, Polyethylene Naphthalate in addition to Polyimide and PTFE (PolyTetraFluoroEthylene) films which are used as a thermal control layer or a optical solar reflector of spacecrafts. We observed accumulation behavior of charge in each sample, survey relationship with each sample's molecular structure. As a result, It's believed that charging characteristics of insulating material is linked to benzene ring in each sample's molecular structure., 資料番号: AA0063997025, レポート番号: JAXA-SP-07-030

Recently many spacecraft use Triple-Junction (TJ) solar cells as their primary electrical power source because of their excellent efficiency. However it is also known that triple-junction solar cells are easy to be broken by a low reverse bias voltage. Therefore a discrete by-pass diode should be connected to every solar cell in parallel for the shadow protection. Under these circumstances, TJ solar cells with integrate Monolithic Diode (MD) have been introduced to market recently. In the CICing (Coverglass Interconnect Cell) of TJ solar cell with MD, cell-to-cell interconnector is connected on MD pad. The interconnector region forms triple-junction in orbit, making primary arc inception easy. Therefore, it is necessary to study the behavior of arcs on MD solar cell array. The result of the ESD (ElectroStatic Discharge) test for MD solar array revealed that the degradation of MD solar cell is caused by not only large current but also total energy of the discharge. The waveform seems to be affected by the impedance of the solar array circuit. This paper presents the recent results of ESD test for MD solar array and proposes further investigation based on the test results., 資料番号: AA0063997013, レポート番号: JAXA-SP-07-030

High-energy electron flux at geosynchronous orbit significantly increases associated with high speed solar wind. The high-energy electron cause internal charging of satellites. It is important for satellite operations to forecast increase of the flux. Here, we report a forecast model using neural network with inputs of solar wind data., 資料番号: AA0063997010, レポート番号: JAXA-SP-07-030

The current status of measuring radiation using JAXA satellites is reviewed. Starting with Engineering Test Satellite-5 (ETS-5; KIKU-5 in Japanese) in 1987, efforts to conduct radiation measurements in space have continued using almost all Japan Aerospace Exploration Agency (JAXA formerly NASDA) satellites (ETS-6 (Engineering Test Satellite-6: KIKU-6), ADEOS (ADvanced Earth Observing Satellite: Midori), ADEOS-2 (ADvanced Earth Observing Satellite-2: Midori-2), MDS-1 (Mission Demonstration test Satellite-1: Tsubasa), DRTS (Data Relay Test Satellite: KODAMA)(ongoing) and ETS-8 (Engineering Test Satellite-8: KIKU No.8)(ongoing), and ALOS (Advanced Land Observing Satellite: Daichi)(ongoing)), in Geostationary Orbit (GEO), Geostationary-Transfer Orbit (GTO), and Low-Earth Orbit (LEO). Electrons, protons, alpha particles, and heavy ions have been the main objects of study. Future plans for radiation monitoring in JAXA, including GOSAT (Greenhouse gases Observing SATellite), Jason-2 (in collaboration with CNES), and ISS (International Space Station)/JEM (Japanese Experiment Module)/Exposure Facility/SEDA-AP (Space Environment Data Acquisition equipment-Attached Payload), are presented., 資料番号: AA0063997004, レポート番号: JAXA-SP-07-030

Preliminary ESD (ElectroStatic Discharge) ground tests with two meter-class solar panels were conducted in a simulated GEO (Geostationary Earth Orbit) environment. The results indicated that the plasma propagation velocity was approximately 10(exp 4) m/s, and the plasma propagation length was greater than 2.8 m. The neutralization current decreased with an increase in the distance from the arc spot., 資料番号: AA0063997019, レポート番号: JAXA-SP-07-030

資料番号: AA0063997028, レポート番号: JAXA-SP-07-030

As the power level of geostationary satellites increases, discharge phenomena on solar array are becoming serious threat to safe operation. Arcs on solar array can short-circuit the satellite circuit, decrease the satellite power, and then cause the satellite permanent failure. To prevent the failure caused by charging and arcing, it is necessary to investigate the mechanism of satellite charging and arcing phenomenon. The purpose of this paper is to investigate the occurrence condition of a secondary arc by measuring arc plasma characteristics in ground test. We measured the arc plasma temperature and identified the materials emitted using spectrometer at arbitrary time during arc occurring. We investigated the difference of secondary arcs occurrence condition during secondary arcs. From the spectroscopic measurement results, we found that it was necessary for shifting to the secondary arc that the metallic vapor same as the cathode material was emitted. In case of PA (Permanent Arc) dimension changes, the probability of secondary arc and TSA (Temporary Sustained Arc) occurrence became high. And plasma temperature was not affected by PA dimension. However, the metallic vapor emission of silver was greatly affected. Thus, secondary arc occurrence greatly depends on metallic vapor emission from cathode., 資料番号: AA0063997017, レポート番号: JAXA-SP-07-030

Possibility of arcing in high voltage arrays have necessitated the theoretical and experimental study of charging and arcing by Indian Space Research Organisation (ISRO), which is planning to increase the satellite bus voltage above the current level of 42 volt. The study, named Spacecraft Plasma Interaction eXperiment (SPIX) was undertaken, together with Institute for Plasma Research (IPR), India to (1) develop software, which is able to predict the floating potential of an object of simplified geometry in space plasma typically of geosynchronous bimaxwellian type, which is also able to calculate the differential potential between two such objects, (2) to develop software to solve the differential equations describing primary arcing phenomena developed by Cho, a pioneer in the field, (3) to study the primary arcing threshold and frequency and to identify the arcing sites of a solar array coupon immersed in a LEO (Low Earth Orbit) like laboratory plasma and to (4) find out the arcing threshold of sustained arcs by introducing a solar array simulator voltage between adjacent strings in a solar array coupon immersed in a LEO-like plasma environment. The experiments have largely yielded results akin to that of other investigators, although some differences have been observed in some experiments. A short term future aim would be to increase the sample space in the last type of experiments mentioned above., 資料番号: AA0063997001, レポート番号: JAXA-SP-07-030

Polar Earth Orbit (PEO) is a peculiar orbit where energetic auroral electrons and low-temperature ionospheric plasma coexist. There is a risk of charging and subsequent arcing in PEO, which was demonstrated unfortunately by the total loss of ADEOS-2 (ADvanced Earth Observing Satellite-2: Midori-2) satellite in 2003. Charging in PEO must be properly assessed in the early satellite design phase by a spacecraft charging analysis tool. The plasma environment, namely its density and energy, is the crucial parameter for the spacecraft charging analysis. The balance between the auroral electrons and the low temperature ionospheric ions determine the charging potential of a PEO satellite. We statistically analyzed the environmental condition in PEO using the data of auroral electrons and thermal ions measured by DMSP (Defense Meteorological Satellite Program) satellites. Probability of a given combination of current densities of auroral electrons and thermal ions has been derived. The energy spectrum of the auroral electrons are also classified into several types and correlation with the thermal ions is analyzed. Combinations of the plasma parameters used by a spacecraft charging simulation software have been identified and database regarding the probability of occurrence of each combination has been formulated., 資料番号: AA0063997006, レポート番号: JAXA-SP-07-030

資料番号: AA0063997029, レポート番号: JAXA-SP-07-030

小型科学観測衛星「れいめい」(INDEX: INnovative technology Demonstration EXperiment)は、2005年8月に打ち上げられ順調にオーロラ微細構造の観測を続けている。我々は、れいめいに搭載された3対のラングミュアプローブ(CRM: plasma CurRent Monitor)を用いたプラズマパラメータの高分解能観測を行っている。CRMのデータはオーロラ帯プラズマの密度、温度の計測に有効であるだけでなく、飛翔体環境の理解を深めるためにも有効で、シミュレーションによる解析と比較することによって、オーロラ粒子による衛星帯電の機構解明に役立つものと考えられる。一般に、極域においては降り込みオーロラ粒子の影響によって衛星は負に帯電し、ラングミュアプローブによるプラズマパラメータの計測は困難であるといわれている。CRMでは、シングルプローブ1対とダブルプローブ2対による複数センサーによる観測を行い、衛星帯電の影響や太陽光による光電子の影響、ウェイクの影響を考慮した観測を試みた。これらの影響は非常に複雑で、計算機シミュレーションとの比較を行うことによって初めて、複雑な飛翔体環境の効果を取り除いたプラズマパラメータの推定が可能になる。れいめい衛星が打ち上げられてから現在までの観測データとNuSPACEによるシミュレーションと比較することによって、CRMによる観測によってプラズマパラメータ以外に衛星の帯電状態、降り込み粒子との因果関係などが明らかになった。, 資料番号: AA0063585014, レポート番号: JAXA-SP-06-035

資料番号: AA0063585002, レポート番号: JAXA-SP-06-035

資料番号: AA0063585027, レポート番号: JAXA-SP-06-035

MUSCAT(Multi-Utility Spacecraft Charging Analysis Tool)は地球低軌道(LEO: Low Earth Orbit)、静止軌道(GEO: Geosynchronous Earth Orbit)並びに極軌道(PEO: Polar Earth Orbit)それぞれの環境に相当する、帯電状態を含む宇宙機-宇宙プラズマ相互作用を解析する計算ツールである。本開発プロジェクトは2004年11月に開始され、2006年3月にβ版がユーザーテストのために公開された。ベータ版では宇宙環境を模擬するために必要な要素の大部分を既に含んでいる。この要素とは、(1)高機能なグラフィカル・ユーザー・インターフェース(GUI: Graphical User Interface)、(2)各種物理現象を模擬するに必要な数値解析機能、(3)高速計算機能、(4)検証実験結果およびスーパーコンピューターによる大規模高精度計算結果との相互比較による妥当性の検証である。(1)として、MUSCATユーザーはGUIを用いて衛星形状と表面材料物性を容易に入力し、解析モデルを作成することが出来る。(2)として、MUSCATは光電子放出、2次電子放出、誘電体表面の導電電流による効果に加え、GEOプラズマのダブル・マクスウェル分布、PEOにおけるオーロラ電子電流の効果を解析モデルに含んでいる。(3)としてMUSCATは並列ワークステーション上での並列計算機能を実装している。さらに4)としてMUSCATの解析結果は検証実験結果と非常に良好な一致を示した。また、MUSCATの試験的な運用をGreenhouse gases Observing SATellite(GOSAT)に対して行った。この目的の1つは地上試験と平行して現実の詳細な衛星形状を用いてGOSATの帯電状態を推定することであるが、もう1つの目的はMUSCATの性能検証である。MUSCATは2007年3月に完成予定である。, 資料番号: AA0063585026, レポート番号: JAXA-SP-06-035

資料番号: AA0063585013, レポート番号: JAXA-SP-06-035

資料番号: AA0063585019, レポート番号: JAXA-SP-06-035

Frequent satellite failures due to charging and arcing on solar array surface, there are strong demands for more thorough ElectroStatic Discharge (ESD) ground test before the launch. It is necessary to validate that a given solar array design can withstand repeated arcing in space before the launch. Test conditions, however, differ among countries, as there is no international standard to define the conditions. An international research project started in 2005 to establish International Standard Organization (ISO) standard on ESD test. In the first year, test coupons of the international round robin tests were produced and the preliminary experiments were carried out. In September 2006, the team members gathered at ONERA, Toulouse France. Demonstration experiments were carried out and in-depth discussion was done regarding the preliminary test results of KIT (Kyushu Institute of Technology). In November 2006, the First International Workshop on ISO Standardization of Electrostatic Discharge Test of Satellite Solar Array was held in Kitakyushu. The first version of the working draft of ISO document was discussed in the workshop and the participants agreed to seek the ISO standard by as early as Fall 2008., 資料番号: AA0063585003, レポート番号: JAXA-SP-06-035

Spacecrafts flying in GEO (Geo-synchronous Earth Orbit) are always exposed to plasma and radioactive-rays such as alpha-, beta- or gamma-rays. Especially in GEO, dielectric materials are expected to be irradiated by high energy electron-beam. In the case of high energy electron beam irradiation, the electrons are injected into the bulk of dielectric materials and they accumulate in them. Because there are no basic data based on practical experiments, it is difficult to expect when and how an accident of ESD (Erector-Static Discharge) caused by accumulated charge will happen on spacecraft. Therefore, we need to measure the charge distribution in the bulk of dielectric materials. We have been developing a system for measuring such a charge distribution in dielectric materials using, so-called, PIPWP (Piezo-Induced Pressure Wave Propagation) method. However, the spatial resolution was low and it was difficult that the space charge distribution measurement in the material used with actual spacecrafts. To obtain a good experiment result, we must use the PEA (Pulsed Electro-Acoustic) method. It is necessary to develop a new PEA measurement system, for a real-time, continuous measurement to secure reliability while irradiating the sample. In this report, It introduces a small, real-time Space Charge Measurement System using PEA method., 資料番号: AA0063585030, レポート番号: JAXA-SP-06-035

資料番号: AA0063585018, レポート番号: JAXA-SP-06-035

資料番号: AA0063585017, レポート番号: JAXA-SP-06-035

In this paper, we report about the test result for Working Group 1 (WG1), which was established to make a Japanese guideline for spacecraft charging design. We investigated threshold conditions of sustained arc inception. The TJ (Triple Junction) GaAs solar cells were tested. Laboratory tests were carried out with an external circuit simulating a spacecraft power system. Solar array coupon panels simulating the hot and return ends of a string circuit were tested under various combinations of string voltage and string current. We revealed that the threshold conditions of sustained arc formation were different in string voltage even if the string current are same., 資料番号: AA0063585032, レポート番号: JAXA-SP-06-035

資料番号: AA0063585009, レポート番号: JAXA-SP-06-035

資料番号: AA0063585031, レポート番号: JAXA-SP-06-035

Electrical charge accumulation characteristics in dielectric materials for a spacecraft panel have studied for understanding of spacecraft charging problem in high energy radiation environment. This study especially has focused on a spacecraft charging accident caused by internal charge accumulation in dielectric materials irradiated by proton beam. The internal charge accumulation of those bulks measurement was carried out using PEA (Pulsed Electro-Acoustic) method. As the results, we obtained a positive charge accumulation in the bulk of samples after proton beam irradiation. However, it was not observed a positive charge accumulation that a quartz glass has an Al evaporated layer on the surface of proton beam irradiation side. Therefore, in this paper, we compare between the PEA's internal charge measurement results and energy band model because of we discus the charge formation mechanism in the bulk irradiated by proton beam. Furthermore, we studied how to measure the secondary electron emission (SEE) of metal and insulating materials used for satellite thermal insulation or other such purposes. SEE yield measurement is very important for analyzing charge accumulation on the satellite surfaces due to the space environment because electron emission. Therefore, we tried to measure the SEE yield. In this report, we introduce the SEE yield measurement results of reference materials (Au, Ag and Quartz glass) and insulating materials. From those results, we discuss the characteristics of SEE that depend on each material., 資料番号: AA0049500035, レポート番号: JAXA-SP-06-006

The key parameters in modeling spacecraft charging are the secondary electron emission (SEE) coefficient and photoelectron emission (PE) coefficient of metals and insulating materials. This work is focused on the analysis and measurement of SEE and PE. A review is presented of methods to measure SEE and PE including the using SEM with short-pulsed electron for accelerating voltage of 600 V-5 kV, a vacuum chamber with a low energy electron beam for accelerating voltage 100 V-2 kV, a measurement system of photo-electron emission, and a photoelectron emission image by vacuum ultraviolet light (VUV). This review is based on joint studies conducted for Japan Aerospace Exploration Agency (JAXA) at High Energy Accelerator Research Organization (KEK), at Musashi Institute of Technology and at Saitama University to investigate the relation to spacecraft charging. This report summarizes some typical measurement results after briefly explaining the measurement project., 資料番号: AA0049500058, レポート番号: JAXA-SP-06-006

Space Solar Power System (SSPS) transfers enormous amount of electrical energy of microwave. As the microwave may diffract to solar array surface, there is concern of multifactor discharge caused by the microwave. There has been no experiment on interaction between high intensity microwave and solar array. The verification experiment is essential for Space Solar Power System to become a reality. We examined the interaction between solar array surface and microwave in a vacuum chamber. We investigated a discharge phenomenon on the solar array surface that receives RF irradiation with 400 W maximum power and 5.8 GHz frequency generated by a magnetron. We observed discharge on solar array depending on the microwave strength and vacuum condition. The discharge occurred mostly at N-electrodes of solar cells that acted as an antenna as good as a dipole antenna. We calculated the distribution of electric field inside the chamber by Finite Difference Time Domain (FDTD) method. The result of numerical simulation shows that macroscopic electric field is not high enough to cause multifactor discharges. Future works involve more precise estimation of local electric field and gas pressure near the N-electrodes of solar cells., 資料番号: AA0049121007

Since a failure of polar orbiting satellites caused by charging has become a serious issue, Japanese Aerospace Exploration Agency has decided to develop a computational tool that can calculate charging status of not only for a polar orbiting satellite but also for a GEO or a low inclination LEO satellite jointly with Kyushu Institute of Technology. The aim of the tool is to give satellite designers chances to identify the charging hazard in the satellite design phase with user-friendly interface. The software named, Multi-utility Spacecraft Charging Analysis Tool (MUSCAT), has been developed since November 2004, and will be released in March 2007., 資料番号: AA0049500033, レポート番号: JAXA-SP-06-006

In the future, LEO spacecraft will be larger and higher powered. Because of the balance of leakage currents through ambient space plasma, their main conductive body will have a higher negative potential without plasma contactor operation. When spacecraft operate with a higher voltage, more intensive electric breakdown, i.e. arcing, is suspected to occur on the surface. In this study, unsteady physical processes inside an arc spot, such as ablation and heating of insulator, and plasma generation and acceleration etc., were studied using Computational Fluid Dynamics (CFD). Direct-Simulation-Monte-Carlo Particle-In-Cell (DSMC-PIC) plasma simulation was also carried out to examine influences of ambient space plasma on plasma expansion processes outside the arc spot. The calculated arc current increased with time; had a peak and then decreased. Inside the arc spot, the calculated plasma resistance rapidly decreased with time; was kept low level and jumped just before extinguishment of arc. As a result the plasma resistance characteristics agreed with the arc current characteristics. The calculated ablation rate rapidly increased with time; had a peak and then gradually decreased, although the calculated arc spot diameter gradually increased with time. Furthermore, the calculated arc spot diameter gradually increased with initial stored energy, although it was smaller than experimental ones. Both the neutral particle number density and the electron number density were the highest near arc initiation and decreased with time. Both the number densities were relatively high inside the arc spot compared with those outside it. The temperature of insulator surface in contact with plasma rapidly increased up to 5,000 K near arc initiation and gradually decreased to 4,000 K. Outside the arc spot, neutral particles in addition to charged particles around spacecraft played an important role in expansion of arc plasma by intensive ionization near the arc spot. Accordingly, high voltage operation of LEO spacecraft might bring drastic degradation of insulator surface by arcing, depending on insulator material properties and ambient plasma conditions., 資料番号: AA0049206064, レポート番号: JAXA-SP-05-001E

In this study, a numerical model of the relationship between the double probe potential with respect to spacecraft potential and the plasma parameters (density, temperature...) has been developed. A fit of the data between 1 and 80/cu cm has been performed which allows to calibrate the model with suitable parameters for the photoelectron emission. The model can then be alternatively used for either density or temperature estimate. Uncertainties and range of validity are discussed., 資料番号: AA0049206034, レポート番号: JAXA-SP-05-001E

The Floating Potential Measurement Unit (FPMU) is an instrument developed to study the spacecraft surface charging of the International Space Station (ISS). It has been developed by Utah State University Space Dynamics Lab under contract to NASA. The surface charging of the ISS is an interesting problem due to its extended size and the use of high voltage solar arrays with exposed interconnects. The FPMU consists of four instruments, a floating potential probe, two Langmuir probes and a plasma impedance probe. These probes will measure the floating potential of the ISS, electron density, and electron temperature with redundancy. The FPMU is being integrated into the International Space Station at one of the existing external camera locations that places it in clear ram flow of the space plasma. Operational constraints of the ISS will result in the FPMU being used to obtain snapshots of data and not as a continuous monitor of the ISS charging and environment. Currently the FPMU is awaiting launch when the Space Shuttle returns to service and may be operational by September 2005. This paper presents an overview of the FPMU., 資料番号: AA0049206082, レポート番号: JAXA-SP-05-001E

In order to realize 400 volts operation in LEO, arcing caused by interaction between spacecraft and surrounding LEO plasma must be overcome. We have been investigating designs of high voltage solar array. This paper is the summary report for all the tests. The coupon with ion barrier film made from ETFE (called 'film coupon') never suffered any arc. We carried out the tests under realistic orbit environment for the film coupon. It was confirmed that the covered film performed successfully in all the situations. This coupon has never arced for more than 25 hours that was equivalent to 1 percent power degradation in 30 years operation. No film contamination and degradation were observed after 90 C baking. No primary arc occurred at all vacuum pressure and no frictional charging occurred by contact. With respect to debris impact, the covered film had little damage even if the film support was hit directly. Also, film attachments were well supported after the heat cycle of - 90 to + 90 C by reducing tension of film. Conductive substrate, however, suffered many arcs at - 400 V. Also, a sustained arc phenomenon between cells and substrate was induced by simulated debris impact. Therefore, use of flexible substrate is adequate for 400 V solar array in LEO environment., 資料番号: AA0049206036, レポート番号: JAXA-SP-05-001E

Electric arcs observed between solar cells have been identified as the origin of power losses on orbit. Such arcs have been identified as 'vacuum arcs' and have been shown to produce severe damage with functional destruction if the solar array provides sufficient current (1 A) under sufficient voltage (20 V) to sustain it. In the work presented here, we deal with power transfer systems encountered in equipment like SADM (Solar Array Drive Mechanism) involving many packaged circuits which are designed to carry high current (greater than 5 A) under voltages higher than 50 volts. This study is an investigation of vacuum arc propagation from one circuit to another. The experiments were performed on samples made with several metallic tracks and up to three electrical circuits (six tracks) were used at the same time. The experimental set-up consisted in triggering a first arc between two adjacent tracks while biasing two, three or four more tracks, and to monitor any kind of connection or propagation. The results obtained show that a vacuum arc can propagate from one circuit to another: arcs triggered by propagation between tracks separated by more than 10 mm were seen. This propagation is due to the expansion of the arc plasma produced by the first arc towards other tracks. From our experiments we have learned that any kind of connections between the electrical circuits are possible but a vacuum arc would more easily propagate by connecting to one of the tracks where the first arc was triggered. Arc movement on one track has been also observed along several centimeters. The damages produced by such arcs were heavy and; destructive: complete vaporization of anode tracks was observed. On cathode tracks, only partial melting and vaporization due to cathode spots of vacuum arcs were seen., 資料番号: AA0049206058, レポート番号: JAXA-SP-05-001E

Geospace Environment Simulator (GES) has started as one of the advanced computing research projects at the Earth Simulator Center in Japan Marine Science and Technology Center since 2002. By using this computing resource, a large scale simulation which reproduces a realistic physical model can be utilized not only for studying the geospace environment but also for various human activities in space. ES project aims to reproduce fully kinetic environment around a satellite by using the 3-dimensional full-particle electromagnetic simulation code which includes spacecraft model inside. Spacecraft can be modeled by the unstructured-grid 3D FPEM code. We will report current status of porting our simulation codes onto the ES and our concept of achieving the satellite environment in conjunction with the space weather., 資料番号: AA0049206066, レポート番号: JAXA-SP-05-001E

資料番号: AA0063347033

(1) ONE PHENOMENON, But Two Primary Manifestations: 1. Surface Charging. 2. Internal Oharging. (2) Existing Design Guidelines for Spacecraft Charging: 1. Guidelines are far from complete at this time, yet there are many. (3) Wide Assortment of Phenomena in Differing Applications: 1. CASE STUDY 1: In-space Insulator Pulsing. 2. CASE STUDY 2: Coupling from a Breakdown into a Circuit. (4) Spacecraft Failure Studies: 1. ELECTRIC FIELD FROM CHARGE IN INSULATING MATERIAL IS A PRIMARY CAUSE OF SPACECRAFT ON-ORBIT ANOMALIES. (5) Electric Field Controls the ESD Pulse Rate and Pulse Amplitude. (6) Complete solution of the problem requires knowledge of electric field -very difficult; Ongoing Studies., 資料番号: AA0049206041, レポート番号: JAXA-SP-05-001E

A large solar array paddle with the power generation of 7 kW was developed for the Advanced Land Observing Satellite (ALOS). With the deployed dimension of 22 x 3 m in a polar orbit, this 9-panel rigid paddle has insulator Silver-Teflon thermal sheets and exposed bypath diode boards on its back face, and cover-glass integrated silicon solar cells on its front face. A charging analysis suggested that large negative potentials on the dielectric back-surfaces and at satellite ground may be induced through the ALOS's auroral passage in off-nominal conditions. The ALOS's baseline panel design was tested for the electron beam radiation and the plasma interaction, which simulated charging situations near the poles. Both the back face and the front face were tested, and arc thresholds were identified. Possibility of sustained arc and surge voltage, as well as survivability against estimated accumulation of arcs, were investigated. The back face showed small negative arc thresholds for both the insulator surface potential and the spacecraft ground voltage. Surface flashover was observed over the silver-Teflon coating. Although both the back and front faces demonstrated immunity against sustained arcs, design modifications to mitigate the back face's susceptibility for charging and arcing were experimentally examined and implemented. The conductive adhesive that surrounded the baseline Silver-Teflon sheet and covered the CFRP face-sheets eliminated arcs at the silver-Teflon edges and the face-sheet. The Kapton film shielding the diode board protected the exposed power line from arcs., 資料番号: AA0049206097, レポート番号: JAXA-SP-05-001E

We carried out electron-beam irradiation experiments concerning the arcing discharges on solar arrays for space use. It is very important to make the discharge mechanism clear and to reflect it to a guideline for designing the solar arrays in order to maintain the high reliability of the future large spacecraft systems. We used a pair of real GaAs solar cells on an aluminum substrate as a sample. The sample was biased to - 9 kV and irradiated by the electron beam of the energy of 10 keV. The voltage from DC battery was applied to the gap between the cells under that condition. The discharge characteristics were obtained as parameters of the gap distance and the applied voltage between the cells. From the waveforms of transient discharge currents, the discharge time was analyzed to evaluate the harmfulness of arcing discharges to solar cells. As a result, the detrimental arcing did not occur even under several conditions of the cell gap distance and the cell gap voltage below 120 V, although a few hundreds of ESD (Electrostatic Discharge) took place. Therefore, the high voltage use of solar arrays is thought not to cause sustained arcing discharge to destroy the satellite system if arrays are designed to lower the maximum voltage between cell strings with an adequate distance., 資料番号: AA0049206018, レポート番号: JAXA-SP-05-001E

The anomaly of Japanese satellite ADEOS-2 (Advanced Earth Observing Satellite 2) appeared in 2003 is supposed to be caused by arcing discharge initiated due to ESD (Electrostatic Discharge) in the bundled wire-harnesses of the solar paddle. In order to make the conditions of the occurrence of the discharges clear, it is necessary to evaluate the discharge characteristics in vacuum quantitatively. Therefore we tried to investigate the characteristics of arcing discharge between two electrodes consisting of a pair of thin copper plates. The discharges were triggered by ESD released from a capacitor. These characteristics were obtained as parameters of the amplitude of the ESD, the pressure in the chamber and so on. We preliminarily obtained the result that the arcing discharge time increased with the amplitude of the ESD., 資料番号: AA0049206019, レポート番号: JAXA-SP-05-001E

The Hypervelocity Impact Facility at Space Research Institute, Auburn University has recently completed a facility upgrade that permits the impact testing of space materials within the cryogenic and elevated temperature range. Sample temperatures within the range of 24-420 K have been achieved for polymer films. These wide temperature range capabilities add to the facilities current testing experience with impact initiated plasma discharge testing for solar cell arrays. The facility utilizes a plasma drag gun to accelerate a variety of simulated micrometeorite or orbital debris materials in the 50 to 150 micrometer range to velocities between 5 and 12 km/s. For each test 5 to 50 particles impact the surface of the target sample within an impact area of approximately 15 cm in diameter. The test chamber can accommodate samples approximately 1.5 x 2.0 m for ambient and heated tests, and 48 x 48 cm for cryogenic samples. The gun and test chambers are evacuated by He cryopumps and dry roughing pumps to produce a clean, oil free environment. Utilizing a streak camera and PMT detection system, the correspondence between individual particle size, speed and impact site can be determined. Standard post analysis yields: micrographs of each impact site, dimensions of the pertinent impact characteristics and individual particle velocity and size estimates., 資料番号: AA0049206020, レポート番号: JAXA-SP-05-001E

Bulk charge accumulation in polymers under electron beam irradiation in vacuum environment was observed using newly developed measurement system. Recently, some accidents in spacecraft due to the charging up of the electric potential have been reported. Some of them seem to be caused by discharge due to the charge accumulation in bulk of insulating materials at relatively higher altitude environment. However, there had been no useful method to measure the bulk charge in materials. On the other hand, we had developed the bulk charge measurement system for the polymeric insulation materials used in insulating materials of power cable. Therefore, we have applied this system to the polymers used in spacecraft. Using the improved system, we have observed internal charge behavior in Kapton and PTFE films under electron beam irradiation in vacuum condition. It was found that an amount of accumulated charge in PTFE was much larger than that in Kapton., 資料番号: AA0049206015, レポート番号: JAXA-SP-05-001E

We proposed a power supply for Electrostatic Discharge (ESD) test on solar arrays. The Current Regulative Diode (CRD) was adopted in this power supply. The CRD can flow a constant current in spite of the applied voltage. The CRD can also reduce the surge current due to a large internal capacitance in the power supply. We measured the performance of the CRD power supply on short-circuit condition. We also tested this circuit for ESD tests. The CRD circuit showed good performance and could cause the sustained arc as well as the secondary arc. We confirmed that the CRD power supply was useful for ESD ground test., 資料番号: AA0049206016, レポート番号: JAXA-SP-05-001E

Polyimide is one of the most widely used polymers as spacecraft materials. Polyimide has also been used as a reference material to measure atomic oxygen fluence in LEO. Erosion properties of polyimide by hyperthermal impact of atomic oxygen under various synergistic exposure conditions need to be well understood for an accurate fluence measurement. In this study, the effect of surface charging on the atomic oxygen-induced erosion of polyimide was studied. The erosion rates of polyimide were not altered by bias voltages up to +/- 1,000 V under 4.8 eV atomic oxygen exposure condition. In contract, 20 percent increase in erosion rate was observed by 1.1 eV atomic oxygen beam exposure. It was confirmed that the effect of surface charging on the erosion of polyimide depends on the translational energy of impinging atomic oxygen., 資料番号: AA0049206060, レポート番号: JAXA-SP-05-001E

The space environment in Low-Earth orbit (LEO) has until recently been considered relatively benign to high levels of spacecraft charging. However, it has been found that the DMSP spacecraft at 840 km can charge to very large negative voltages (up to - 2,000 V) when encountering intense precipitating electron events (auroral arcs) while traversing the auroral zone. The occurrence frequency of charging events, defined as when the spacecraft charged to levels exceeding 100 V negative, was highly correlated with the 11-year solar cycle with the largest number of events occurring during solar minimum. This was due to the requirement that the background thermal plasma density be low, at most 10(exp 4)/sq cm. During solar maximum, the plasma density is typically well above that level due to the solar EUV ionizing radiation, and although the occurrence frequency of auroral arcs is considerably greater than at solar minimum, the occurrence of high-level charging is minimal. Indeed, of the over 1,600 events found during the most recent solar cycle, none occurred during the last solar maximum. A dependence of the magnitude of the spacecraft frame charge on density was observed with the largest negative voltages, approximately 2,000 V, only occurring at very low densities. A 'worst-case' charging electron spectrum for low-altitude was developed by applying a potential drop of 20 kV to source data that went into the MIL-STD-1809 specification for near-geosynchronous orbit and combining it with DMSP spectra consistent with a 20-kV drop. The results of this study have implications on a number of LEO satellite programs, including the International Space Station (ISS). The plasma density in the ISS orbit, at a much lower altitude than DMSP, is well above that at 840 km and rarely below 10(exp 4)/sq cm. However, in the wake of the ISS, the plasma density can be two orders of magnitude or more lower than the background density and thus conditions are ripe for significant charging effects. With an inclination of 51.6 degrees, the ISS does enter the auroral zone, particularly during geomagnetic storms and substorms when the auroral boundary can penetrate to very low latitudes. This has significant implications for EVA operations in the ISS wake., 資料番号: AA0049206094, レポート番号: JAXA-SP-05-001E

Validation test cases for the SPARCS code (S. Clerc et al., 8th SCTC, 2003) are presented. More specifically daylight charging of spacecraft is addressed. The main mechanism for charging in daylight is the formation of a potential barrier which leads to partial recollection of secondary electrons and photo-electrons. We first validate the recollection of photo-electrons using the particle tracking capabilities of SPARCS. The test cases involve a positively biased metallic surface. The spherically symmetric and the plane configurations are studied. Numerical results for the net total yield agree with the analytic expression. To validate the potential barrier formation due to surface charging, we compute the charging of a dielectric sphere in sunlight. Due to the complexity of this problem, an exact analytical solution cannot be found. Instead, a simplified monopole-dipole approximation is shown to give a reasonable agreement. Finally, the computation of daylight charging benchmark of a model telecom spacecraft is presented. Agreement with published results (Davis et al. 8th SCTC, 2003) is good., 資料番号: AA0049206069, レポート番号: JAXA-SP-05-001E