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本収録ではSuzaku(すざく: ASTRO-E2: X-ray Astronomy Satellite)によるRCW86北東shellの観測結果を報告する。イメージングおよびスペクトル解析によって、この領域からのX線を低温(〜0.3keV)および高温(〜1.8keV)のthin-thermal Plasmaと、シンクロトロン放射成分に分離、それぞれ分布が異なることを明らかにした。高温成分は低階電離の鉄を多量に含み、熱的filament(低温成分)よりSNR(SuperNova Remnant)の内側に強く分布する。また、シンクロトロンX線とは全く相関がなかった。これらの事実は高温成分の起源が逆行衝撃波によって加熱されたFe-richなイジェクタであることを強く示唆する。RCW86の北東領域では1部が比較的最近になって分子雲などの高密な領域に衝突し、順行衝撃波が一気に減速したと考えられる。その結果、宇宙線加速効率が急激に低下し、同時に逆行衝撃波が減速もしくはSNRの中心に向かって後退し始めたのであろう。これに対し、シンクロトロン放射が強い領域は未だ密度の小さい空間中を高速で膨張しており、効率的に宇宙線を加速し続けていると解釈できる。, 資料番号: AA0063722003

Puppis A超新星残骸ほぼ全体をカバーするXMM(X-ray Multi-Mirror)-Newtonによる5点の観測結果について報告する。我々は、O、Ne、Mg、Si、S、Fe元素について、等価幅分布図を作成した。その結果、Fe以外の元素については爆発噴出物の痕跡を残す領域を見つけた。一方で、Feについては至るところ、等価幅が小さいことが判った。O、Neの全体的な等価幅分布の構造は似通っており、X線放射強度の弱い南部で等価幅が大きいことが判った。これは、SNR(SuperNova Remnant)南部では星間物質の密度が低く爆発噴出物の影響が露になっているためと考えられる。SiとSの等価幅分布図はお互いに似通っており、北東部分でのみ爆発噴出物の痕跡を発見した。SNR全体的な構造に加えて、北東領域に爆発噴出物の塊を見付けた。更にその輝線がドップラーシフトしており、高速でこちらに向かって飛んでいることを発見した。, 資料番号: AA0063722002

資料番号: AA0063351017

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

2005年度にJAXA情報・計算工学センター(JEDI; JAXA's Engineering Digital Innovation Center)が発足し、これまで行われた宇宙プラズマ物理解析のプラズマシミュレーションを衛星機器の技術開発、衛星環境アセスメントなど工学的な分野に応用することが求められている。この要求に対し、JEDIでは、宇宙機と周辺宇宙プラズマの相互作用を模擬するJAXA衛星環境プラズマ数値シミュレータの基礎開発を2006年度から開始する。この衛星環境プラズマシミュレータは、これまで宇宙プラズマ現象の解析に用いられてきたParticle-In-Cell(PIC)プラズマシミュレーションをコアエンジンとし、これに、現在JAXA/九州工業大学で開発が進められている衛星帯電解析ツール(MUSCAT; Multi-Utility Spacecraft Charging Analysis Tool)で開発された高分解能衛星モデリング機能を付加することにより、宇宙プラズマによる衛星環境への影響について定量解析を行う数値シミュレーションツールである。特に、MUSCATでは解析することが困難なプラズマ非定常過程や電子ダイナミクスが関与する現象について、JAXAスーパーコンピュータを最大限に駆使して大規模数値シミュレーションを行うことによって解析を進める。具体的には、衛星からの能動的プラズマ放出による衛星環境への影響に着目し、将来的には、イオンエンジン、プラズマコンタクター、などプラズマをアクティブに利用した推進系、帯電緩和技術などの技術開発、磁気プラズマセイルなど新規航行技術開発に貢献することをめざしている。2006、2007年度は、シミュレータの基礎開発として、PICシミュレータのJAXAスーパーコンピュータへの移植、MUSCAT衛星モデリングとのリンク、衛星からの能動的プラズマ放出を模擬できるルーチンの開発とその動作テストを行い、2008年度以降のシミュレータの本格的開発への準備を行う。, 資料番号: AA0063157002, レポート番号: JAXA-SP-06-014

磁気プラズマセイル(Magneto Plasma Sail; MPS)とは、探査機まわりに磁気圏を形成し、高速のプラズマ流である太陽風を「磁場の帆」で受け止め、太陽風の動圧を推力として利用する推進システムである。本推進システムにおける推力は「帆」に相当する磁気圏の大きさに比例するため、磁気圏を拡大するために、探査機周りに形成された磁場にプラズマを噴射することにより、効率よく磁気圏を広げる方法が検討されている。本報告では、探査機周りの流れ場がどのように形成されるかを調べるために、電磁流体方程式による数値シミュレーションを行い、噴射プラズマによる磁気圏の変化や太陽風と磁場の干渉の様子を調べた。その解析結果について報告する。, 資料番号: AA0063157010, レポート番号: JAXA-SP-06-014

現在のプラズマの主要な計算機実験法であるPIC(Particle In Cell)法では、デバイ長よりも小さいスケールでの実験はその原理上困難となっています。私達は、今までの手法では計算機実験を行うことのできなかった微小スケールでのプラズマの計算機実験の手法を確立し、プラズマのデバイ長が宇宙機と比べて大きくなる深宇宙においての宇宙機の性能評価法を確立することを目標として研究・開発を行っています。具体的には、天文の分野で使われている多体問題専用計算機'GRAPE-6'を用いて、プラズマ内の粒子にかかるクーロン力を直接計算することにより、プラズマ内の粒子の動きを直接追う純粋な粒子法の確立を目指しています。プラズマ内の粒子の動きを直接に追う場合、プラズマに含まれる粒子の数が膨大であるため、計算時間も膨大になってしまいます。そこで、'GRAPE-6'の粒子間の相互作用を高速に計算する能力に注目し、現実的な時間内での粒子法によるプラズマの計算機実験を行いたいと考えております。現在、その第1段階として、'GRAPE-6'を用いた粒子法のラングミュアプローブモデルによる検証を行っています。この研究はJAXAオープンラボによる支援を受けて行われています。, 資料番号: AA0063157007, レポート番号: JAXA-SP-06-014

A quick review of plasma sail propulsion is provided to close up the technical and physical issues of plasma sail. A spacecraft propulsion system utilizing the energy of the solar wind was firstly proposed by Prof. Winglee in 2000. Although Winglee proposed an M2P2 (Mini-Magnetospheric Plasma-Propulsion) design with a small (20-cm-diamter) coil and a small helicon plasma source, it was criticized by Dr. Khazanov in 2003. He insisted that: (1) MHD (Magnetohydrodynamic) is not an appropriate approximation to describe M2P2 design by Winglee, and with ion kinetic simulation, it was shown that the M2P2 design could provide only negligible thrust; (2) considerably larger sails (than that Winglee proposed) would be required to tap the energy of the solar wind. We started our study in 2003, and it is shown that moderately sized magnetic sails can produce sub-Newton-class thrust in the ion inertial scale (approximately 70 km). We continue our efforts to make a feasibly sized plasma sail (Magnetoplasma Sail) by optimizing the magnetic field inflation process Winglee proposed. However, several physical issues (action and reaction force of the MPS, finite Larmor radius effect in the far region of the MPS, and so on) were shown to remain to be solved before we proceed to a system design of a spacecraft propelled by MPS., 資料番号: AA0063157008, レポート番号: JAXA-SP-06-014

資料番号: AA0063354007

第10回宇宙科学シンポジウム (2010年1月7日-8日. 宇宙航空研究開発機構宇宙科学研究本部), 10th Space Science Symposium (January 7-8, 2010. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)), Sagamihara, Kanagawa Japan, 形態: カラー図版あり, 形態: CD-ROM1枚, 著者人数: 19人, Physical characteristics: Original contains color illustrations, Note: One CD-ROM, Number of authors: 19, 資料番号: AA0064742075

資料番号: AA0063354032

1924年I.Langmuirによって理論的に確立されたDCプローブ(Langmuir and Mott-Smith、1924)は地球を取り巻く電離圏における基本的な物理量である電子温度、密度を測定するプローブとして、第二次大戦直後米国がドイツから持ち帰ったV-2 rocketに搭載されて以来(Reifman and Dow, 1949)、多くの観測ロケット、科学衛星に使用されてきており、電離圏研究においてDCプローブは必要不可欠の観測器として、極めて重要な役割を果たしてきた。本稿はまず、地球電離圏E、F領域、電離圏上部および内部プラズマ圏における電子温度に関して世界の最先端を走ってきた日本が得た最近の成果を中心に概説する、本稿の最後に電子温度、密度に関する未解決の物理現象、および今後の測定に関する課題について記述する。, DC Langmuir probe which was established by I. Langmuir in 1924 has been used to measure Te (electron temperature) and Ne (electron density) including plasma irregularities in the ionosphere since 1947 and its contribution to the ionospheric Physics is beyond our description. We review the measurement of electron temperature/density in space, which have been conducted by means of Langmuir probe. We review the recent findings which have been obtained in Japan with Langmuir probes in Japan and list up the problems, which still remain unsolved., 資料番号: AA0049213001, レポート番号: JAXA-SP-05-020

We are studying the problems associated with high voltage power systems in space. Especially we are interested in the potential distribution of the solar array that is resistant to the electrical discharge. We have carried out experiment on the interaction between the space plasma and the high voltage solar array. An array of electrodes distributed on a dielectric material was used to simulate the inter-connectors of the solar array panel in space environment. One of major concerns in the usage of the high voltage solar array in space is the arc discharge on the array. Based on the plasma sheath theories, there is a possibility to control or to prevent the discharge by selecting a potential distribution of the electrode array. As the first step to find the potential distribution that is tolerate to the discharge, we measured the current to the electrodes, changing the spacing of the electrodes. This paper presents the experimental results suggesting that we can control the discharge by selecting a proper potential distribution of the high voltage solar array., 資料番号: AA0049213013, レポート番号: JAXA-SP-05-020

ISAS/JAXAスペース・プラズマ共同研究設備において、種々のスペース・プラズマ関連実験に対応できるプラズマを容易に生成可能な大容量(内径73.8cm、長さ486cm)ヘリコンプラズマ装置を開発したので、装置の概要、この装置で生成されるプラズマの特性について報告する。本ヘリコン装置では、励起アンテナとして真空容器端のクウォーツ・ガラス窓外大気中に設置されたフラット・スパイラル・アンテナを使用している。ヘリコン放電の場合、n(sub e)≧10(exp12) /立法センチメートルという高密度プラズマを1kW以下の入力パワー(周波数7MHz)で生成可能という高効率rf放電を達成している。因みに、現在までに使用したガス種はアルゴン、ヘリウム、水素であり、上記高密度モードはアルゴン、ヘリウムで達成された。また、(1)励起アンテナ近傍の磁場配位を調節する、(2)励起アンテナからの電磁放射パターンを変化させる、という2通りの方法でプラズマの径方向密度プロフィルを容易に制御できることを見出した。, A large-volume (73.8 cm in diameter and 486 cm in axial length) helicon-plasma device has been developed at ISAS/JAXA in order to perform various space-plasma related experiments. With the use of a large (43 cm in diameter) flat-spiral antenna located just outside a quartz-glass window at one end of the vacuum vessel, the device is capable of producing high-density plasma (gas species used so far are argon, helium, and hydrogen) of n(sub e) that is greater than or equal to 10(exp 12) /cu cm with moderate rf input power of less than 1 kW (at 7 MHz), showing excellent discharge efficiency. Innovative methods of controlling the radial profile of the plasma density have also been demonstrated (1) by changing the magnetic field configuration near the spiral antenna and (2) by changing the antenna radiation-field pattern., 資料番号: AA0049213008, レポート番号: JAXA-SP-05-020

深宇宙探査に適した推進システムとして、太陽風のエネルギーを利用した推進システムがいくつか提案され研究されているが、これらの推進システムの多くについては推力の定量的な評価がなされておらず、発生メカニズムの理解が不十分である。本稿では、これら推進システムのうち最もシンプルなシステムである磁気セイルの推力特性について数値解析を行ない、推力発生メカニズムを明らかにした。この磁気セイルおよび磁気セイルと類似した新しい推進システムの研究状況を合わせて報告し、これら太陽風を利用した推進システムの将来展望を探る。, For the last few years, several deep space propulsion systems making use of the energy of the solar wind have been proposed and researched. Thrust production mechanisms of these propulsion systems, however, have not been understood well, hence they are now intensively studied in the United State and in Japan. In this paper, the thrust characteristics of the Magnetic Sail, which is considered as one of the simplest system in these propulsion systems, are quantitatively obtained by numerical simulations, which successfully revealed the momentum transferring process from the solar wind to the coil of the Magnetic Sail. Also, other propulsion systems making use of the solar wind are introduced in comparison with the Magnetic Sail, so as to reveal the unknown physics and the applicability to deep space missions., 資料番号: AA0049193001, レポート番号: JAXA-RR-05-014

資料番号: AA0063353027

資料番号: AA0063353026

地球磁気圏において1keVから100keVの電子は中エネルギー電子と呼ばれ、電子の加速加熱メカニズムを知る上で重要である。しかし技術的に数keVから数10keVの電子はギャップ領域になっており、正確な計測を行うことは困難であった。我々はアバランシェ・フォトダイオード(APD: Avalanche Photodiode)という素子に着目して、中エネルギー電子の計測が高精度で行えることを示した。また、JAXAの観測ロケットS-310-35号機によってAPD素子の宇宙空間での実証実験を行い、3.5-65keVの電子の計測に成功した。, 資料番号: AA0063353048

資料番号: AA0047899104

資料番号: AA0049122108

フォーメーションフライト地球磁気圏観測ミッションSCOPEは、遠地点30Reの長楕円軌道上を5機の衛星が互いに5〜5,000km程度の距離で編隊飛行することにより、地球磁気圏におけるプラズマ、電磁場、プラズマ波動の多点同時観測を行う。多点同時観測によって測定対象の空間的、時間的な広がりを正確に知るには、観測時の各衛星の相対位置、相対時刻を正しく把握する必要がある。現在必要とされている相対位置、時刻決定精度は最高で50m、5μsecであり、これは従来の地上からの軌道決定では不可能な値である。本稿では衛星間相互通信を用いて、上述の精度を達成する相対軌道決定手法について、検討状況を紹介する。, 資料番号: AA0049122049

資料番号: AA0049122047

宇宙航空研究開発機構(JAXA)・宇宙科学研究本部(ISAS)では、GEOTAILにつづく地球周りのプラズマ観測ミッションとして、SCOPE(cross-Sale COupling in Plasma universE)ミッションを検討中である。SCOPEでは、5機の衛星を1〜5,000kmという距離をおいて配置することで、これまで困難とされてきた波動の空間・時間成分を分解することを目指す。衛星システムの観点から見た本ミッションの際立った特徴は、複数の衛星が編隊飛行(フォーメーションフライト)を行うことであり、このために、各衛星の相対的なダイナミクスを高精度で把握する技術が必須となる。また、編隊を構成する5機の衛星は、観測データや相対測距データのやりとりのため、地上の支援のないときにも有機的に結びついている必要があり、自律的な衛星間通信の技術も要求される。本稿は、SCOPEの概要および、2004年度のSCOPEワーキンググループでの検討結果を紹介するものである。, 資料番号: AA0049122048

Bare-tether systems are one of the greatest-efficiency electrodynamic tethered systems. The system with an uninsulated portion of the metallic tether itself to collect electrons from the space plasma is operated as a thruster or a power generator on a satellite. Ground experiments were carried out to understand phenomena of electron collection by a bare tether in space. Metallic tether samples were exposed to a simulating Low-Earth-Orbit plasma flow as varying tether sample diameter and length, and plasma velocity. A magnetic field was also applied. The normalized collection current increased with plasma velocity. The existence of magnetic field raised the normalized collection current because of the edge effect of a tether sample or the three-dimensional effect. A high collection current above the orbital-motion-limited current could be achieved with a magnetic field. The current characteristics of a bare tether strongly depended on plasma velocity and surrounding magnetic field., 資料番号: AA0049206070, レポート番号: JAXA-SP-05-001E

The main objective of the SPIS project is to offer an open source, homogenous and extensible range of numerical tools to model precisely spacecraft plasma interactions. Regarding the spacecraft charging process, the scope of expected applications of SPIS is much larger with, for instance, the modeling of electrostatic sheath structure or the characterization of space plasma instruments. New technologies like the electrical propulsion or high power solar generators open new technical issues and challenges, achievable only through an accurate modeling. Today, application domains are both scientific and industrial, taking into account different approaches, constraints and cultures of users. SPIS aims to be an easy-to-use and fast engineering and scientific code, accurate and adaptable to specific and complex cases. From the User Interface (UI) point of view, to bring together both the scientific and industrial spheres represents a real challenge because their user requirements are different. The necessity to take into account a very complex modeling, that includes complex pre and post-processing phases, worsens this situation. The adaptability and extensibility, which are the key requirements, of the system prevent a classical approach based on monolithic codes. In order to follow the adaptability and extensibility of the framework, the user interface has to follow the Object Oriented Approach (OOA), and the possibilities of dynamic loading offered by modern languages, like Java and Python/Jython. Regarding the simulation kernel, SPIS-NUM, which constitutes the core of SPIS system, this analysis led to the development of a proper sub-project, called SPIS-UI, to handle the issues previously mentioned., 資料番号: AA0049206052, レポート番号: JAXA-SP-05-001E

Results from SMART-1 plasma measurements indicate that the spacecraft is in general floating 10 to 30 Volts negative with respect to the ambient plasma. An analytical model based on the interaction of the charge exchange plasma with the spacecraft and the solar panels is shown to be in good quantitative agreement with most of the observations. Progress for simulating the overall interaction between the thruster plume and the spacecraft by the use of PIC simulation codes is also presented., 資料番号: AA0049206100, レポート番号: JAXA-SP-05-001E

We summarize the characteristics of the DC electric field measurement by the double probe system, EFD-P, aboard Geotail. The accuracy and correction factors for the gain (effective length) and off-set, which depends on ambient plasma conditions, are provided., 資料番号: AA0049206028, 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

When we operate solar array at high voltage (greater than 100 V) in LEO environment, arcing occurs on the solar array surface. The solar cell can suffer degradation of electric performance due to only one arc with some probability. We performed degradation test of solar array coupons against arcing under simulated LEO environment in order to identify the threshold of arc energy to cause the solar cell degradation. The solar array coupons are made of silicon solar cells. The LCR circuit was attached to a solar array coupon through a vacuum chamber. In order to change trigger arc energy, we changed the value of capacitance of LCR circuit. The electric performance was measured inside the chamber without opening the chamber. The threshold of arc energy was identified. The arcing which has enough energy to cause the solar cell degradation may occur on solar array of 100 V satellite in Low Earth Orbit environment., 資料番号: AA0049206027, レポート番号: JAXA-SP-05-001E

The detrimental effects of spacecraft charging are well known, particularly when the charging leads to arcing on solar array and consequently to permanent power loss. Motivated by these concerns, ALCATEL SPACE has undertaken several test campaigns to qualify solar array design and to investigate margins. This paper describes the state-of-the-art of ALCATEL SPACE solar array ESD laboratory test set-up and procedure. The primary discharges are created using a plasma source and a negative polarization of the sample rear side. The required power to feed the arc can be delivered by a Solar Array Simulator with low capacitance power sources. The test sample is fully representative of a flight model., 資料番号: AA0049206045, レポート番号: JAXA-SP-05-001E

Space plasma can build up high differential potentials on satellite surfaces resulting in Electrostatic Discharges (ESD), sometimes causing anomalous behavior of satellite systems. As these phenomena give rise to possible risks affecting the performance of satellite, the charging is one of major concerns for satellite manufacturers. It has been investigated the influence of charging phenomena on satellite systems and its mitigation techniques. Our activities include not only investigation of influences on power systems but also that of other risks such as Electromagnetic Compatibility (EMC) or immediate surge to electronic device. In this paper, some of our assessments and the concerned problems are reported. It is obvious that international standardized guidelines for designing satellite systems against ESD will be established in the near future. It would be desirable if the requirements shall be appropriate and practical for satellite systems design, manufacturing and testing., 資料番号: AA0049206035, レポート番号: JAXA-SP-05-001E

A new Spacecraft Plasma Interaction Software, SPIS, has been developed in the frame of the Spacecraft Plasma Interaction Network, SPINE. This software can a priori simulate the kinetics processes of ions and electrons taking into their space charge and their interaction with spacecraft surfaces. It is open source and freely available worldwide. While the development and the qualification of the software functionalities is under the responsibility of a consortium led by ONERA under ESA contract, the test and validation of the applicability of the code to physics problem is under SPINE responsibility. The validation program includes step-by-step applications of the software to sheath modeling in simple geometry, artificial plasma injection, and spacecraft charging. We report here on the progress along this program including Langmuir probe tests with spherical and cylindrical geometry, and comparison with other software for more complex geometries., 資料番号: AA0049206053, レポート番号: JAXA-SP-05-001E

The development of a new software for spacecraft plasma interactions modeling was started in Europe at the end of 2002. This Spacecraft Plasma Interaction Software (SPIS) is developed for and by SPINE community (Spacecraft Plasma Interaction Network in Europe) on an open source basis. The ESA contractors, ONERA and Artenum, are in charge of the development of SPIS framework and main numerical modules. In the framework of this collaborative development, SPINE community has several roles. It first expressed its requirements. It is now testing and validating the code on challenging test cases, and is starting to develop extra modules. The software framework is based on the integration or interfacing with available open source tools for CAD, 2D meshing, 3D meshing, GUI, post-processing and graphical display. The numerical routines allow the modeling of the plasma dynamics (kinetic or fluid, electrostatic with possible extension to electromagnetic) and its coupling with the spacecraft (equivalent circuit approach). The modeling of all types of environments and devices is supported (LEO/GEO/PEO..., EP/solar arrays...). The emphasis was put on the modularity of the code to allow the interoperability of the modules, through an object-oriented (OO) approach throughout the code (Java language, Jython script language). Requirements and design of SPIS code were defined during the first semester of 2003. The first version of the code was released in spring 2004, and extended versions later in 2004 and 2005. This paper mostly focuses on the numerical core of the solvers and their OO architecture., 資料番号: AA0049206051, レポート番号: JAXA-SP-05-001E

Recently solar array has become higher in potential and larger in capacitance. Therefore, possibility of collision between space debris and enlarged solar array has been pointed out. Actually, a lot of debris and dust impacts were confirmed on fuselage of retrieved satellite SFU and solar array paddle of satellite Eureca. If space debris collides with solar array of an orbit satellite, it causes generation of high-density plasma by debris impact induced dielectric breakdown of satellite component and the phenomenon called discharge. This discharge short circuit and current does not flow into a load of the satellite. And the very worst event by this discharge is operational end of the satellite. However, any events of discharge phenomenon by debris impact cannot be yet confirmed. But we cannot ignore such possibility of discharge by debris impact. The purpose of the present paper is to investigate discharge condition due to debris impact which yields us reduction of electric power of solar array, and to reduce influence of the impact on satellite missions. In this study, a solar array coupon was tested under hypervelocity impact in which a projectile was launched by a two-stage light gas gun installed in KIT. As a result, we verified discharge event in the hypervelocity impact ground test., 資料番号: AA0049206078, レポート番号: JAXA-SP-05-001E

Since the 1970's, Balmain, Bogorad, and others proposed seemingly different scaling laws for using small-scale laboratory test results to estimate the peak of current transients from Electrostatic Discharges (ESD) on thermal blankets, solar panels, and other large dielectric surfaces on satellites in earth orbit. We derive a general scaling law and show that the scaling laws proposed by Balmain, Bogorad, and others are special cases. Moreover, we show that for arbitrary shape of the surface area of a dielectric, one can predict both the pulse shape, I(sub ESD)(t), and the rate of change, dI(sub ESD)(t)/dt, of an ESD-induced transient., 資料番号: AA0049206084, レポート番号: JAXA-SP-05-001E

資料番号: AA0063347020

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 the space development and utilization, it is very important to understand the interactions between spacecraft/structures and space plasma environment as well as the natural phenomena occurring in space plasma. In order to evaluate the spacecraft-plasma interactions quantitatively to contribute to the progress of space utilization and space technology, we aim to develop a proto model of Geospace Environment Simulator (GES) by making the most use of the conventional full Particle-In-Cell (PIC) plasma simulations. For the development of GES, we have been using the Earth simulator which is one of the fastest supercomputer in the world. GES can be regarded as a numerical chamber in which we can virtually perform space experiments and analyze the temporal and spatial evolution of spacecraft-plasma interactions. GES will be able to provide fundamental data regarding various engineering aspects such as the electrostatic charging and electromagnetic interference of spacecraft immersed in space plasmas, which will be useful and important information in determining designs and detailed specifications of spacecraft and space systems., 資料番号: AA0049206054, レポート番号: JAXA-SP-05-001E

Spacecraft Charging Mitigation is an art form that dates back to the 1960's. The most common practice involves controlling the resistance per unit surface area, and/or the resistance per unit volume of dielectric materials. This is implemented with proper bonding to reference ground to provide a charge bleed-off path. We show how grounding configuration affects the effectiveness of the mitigation method, and we propose a general rule-of-thumb that guarantees optimal results., 資料番号: AA0049206092, レポート番号: JAXA-SP-05-001E

Behind a spacecraft in a supersonic ion flow, a negatively charged wake will form. In a sufficiently tenuous plasma, where the potential of a sunlit spacecraft becomes positive, the size of the ion wake can become very much larger than what may be expected from the spacecraft geometrical size. This occurs if the energy of the flowing ions is less than what is needed to overcome the positive spacecraft potential. As the ion flow must be supersonic, this means that the plasma flow has to be quite cold. It turns out that such conditions are quite common in the terrestrial magnetosphere, particularly in the polar wind region. The polar wind cold plasma flow at geocentric distances outside 10 Earth radii can have ion flow kinetic energy around 10 eV, ion temperature of a few eV and be so tenuous that the potential of a spacecraft reaches above 20 V. This plasma is rarely measured on scientific spacecraft, as the ions cannot reach an ion detector mounted on the highly positive spacecraft. Consequently, there are only a few studies of its properties and distribution. However, the cold ion density can be estimated by comparing the number of ions actually detected on the spacecraft with independent density estimates, for example spacecraft potential measurements or wave signatures. Another effect of the negatively charged wake is its impact on electric field measurements on the spacecraft, which shows a clear signature of the wake electrostatic potential. While this is a contamination to the measurement of the natural electric field, it is also potentially useful to derive flow properties like velocity and temperature. In order to understand the problems and exploit the possibilities of this effect, we have analyzed data from the Cluster double probe instrument EFW and used the PicUp3D code to perform PIC simulations. This study has given new information on the properties of the wake, as well as the photoemission from the electric field probes., 資料番号: AA0049206089, レポート番号: JAXA-SP-05-001E

Performances of ED-tethers using either spherical collectors or bare tethers for drag, thrust, or power generation, are compared. The standard Parker-Murphy model of current to a full sphere, with neither space-charge nor plasma-motion effects considered, but modified to best fit TSS1R results, is used (the Lam, Al'pert/Gurevich space-charge limited model will be used elsewhere). In the analysis, the spherical collector is assumed to collect current well beyond its random-current value (thick-sheath). Both average current in the bare-tether and current to the sphere are normalized with the short-circuit current in the absence of applied power, allowing a comparison of performances for all three applications in terms of characteristic dimensionless numbers. The sphere is always substantially outperformed by the bare-tether if ohmic effects are weak, though its performance improves as such effects increase., 資料番号: AA0049206085, 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

The classical estimations give the maximum potential of spherical satellite (spacecraft) during the electron beam injection in the ionosphere that much exceed the value obtained in real rocket experiments even those conducted at night and without artificial neutralization. The possibilities of compensation are considered on the basis of theoretical models taking into account the collective beam-plasma interactions. The examples of numerical modeling of the charge-discharge dynamics are provided for the APEX spacecraft conditions. It is shown that changing the pulse form and rise rate (dI/dt) can change the structure of space charge zone i.e. try to create the strong disturbance or resonance conditions in the time interval of (pi/omega(sub p)) less than or equal to t less than or equal to t(sub ion). From these positions we consider the APEX facilities used to control the inflow current, the spectrum of energetic particles, the high-frequency oscillations and the spatial distribution of optical emissions in the injector vicinity with high time resolution., 資料番号: AA0049206029, レポート番号: JAXA-SP-05-001E

資料番号: AA0063347011

資料番号: AA0048468007, レポート番号: JAXA-SP-04-010

資料番号: AA0048468025, レポート番号: JAXA-SP-04-010

資料番号: AA0048468014, レポート番号: JAXA-SP-04-010