中野 正勝

Miniature water ion thrusters provide CubeSats with orbit transfer capability and enable missions that require high [Formula: see text]. Erosion on the grids is critical for the thruster’s lifetime and propellant utilization efficiency. It was shown numerically that the nonuniformity of the plasma in the discharge chamber causes inhomogeneous erosion of the grid. However, deviation from experiments was still a problem. In this study, we evaluate the performance and the change in the grid erosion of the miniature water ion thruster during 1000 h of operation. The propellant utilization efficiency decreases from 16% to 12% due to grid erosion. The contamination effect of the antenna and magnets on the performance of the thruster was less than that due to the change of grid geometry. In terms of grid erosion, in the center and the periphery of the grid where the ion beam density is low, the direct impingement of the ion beam is dominant, and the diameter of the accelerator grid eroded from 0.31 to 0.52 mm. On the other hand, in the middle part of the grid, where the ion beam is dense, charge exchange ions dominate the erosion, and the diameter was eroded from 0.31 to 0.39–0.45 mm in hexagonal.

<p>In order to improve ion engine's performance, the neutralization performance with two field emission cathodes in an ion engine is investigated. The neutralization performance is evaluated by the potential difference between cathode and ground, using a 100μN class ion thruster developed at Kyushu University and two 50×50mm² field emission cathodes with carbon nanotube emitter. The potential difference between cathode and ground is not only determined by cathode electron supply capacity and the position of the cathodes but also foot print of the neutralizers, it would be due to the space charge limitation. That is, the potential difference between cathode and ground is improved with increase in total emission current, and that with a single field emission cathode at emission current of 6mA is -20V, on the contrary, that with two field emission cathodes is -12V. </p>

<p>The grid design of ion thrusters is currently based on trial and error development and durability tests. These methods make it difficult to rapidly develop optimal thrusters for diverse missions. As a way to address this problem, this paper provides a method of designing grids using numerical simulation with a genetic algorithm. Screen grid thickness, screen grid hole diameter, accel grid thickness, accel grid hole diameter, accel grid potential, and grid gap are optimized for long-life grids. The JIEDI tool is a powerful tool to design grid systems; it is used here as a fitness function to evaluate the lifetime of the grid in an ion thruster. As a result of the optimization, the lifetime of the model developed using the genetic algorithm is about 6.5 times longer than that of the experimental model. The optimization of grid parameters using the Genetic Algorithm and the JIEDI tool shows promise for future design tasks.</p>

<p> 電気推進を用いた軌道間輸送機（OTV）の開発における技術的挑戦と克服に関するパネルディスカッションが行われた．太陽発電衛星の輸送では，ペイロードの太陽電池をOTVで利用できる場合，電気推進は特に有効な推進システムとなる．電気推進の推進剤として有力な候補はアルゴンであり，大出力のアルゴン推進機を開発する上での技術的課題が議論され共有された．また太陽発電衛星の実現には打ち上げ機との連携や，OTV全体でのシステム最適化が必要不可欠であり，分野を横断した研究者間の協調と，共通で意識すべき事業戦略（ロードマップ）の策定が重要であると再認識された.</p>

<p>The two-grid ion optics of a water propellant miniature ion propulsion system (MIPS) was numerically studied by full-aperture ion-optics simulation considering an inhomogeneous plasma source in the discharge chamber. Three dominant ion species (H₂O⁺, OH⁺, and H⁺) of water vapor plasma were included in the model. Their trajectories were tracked considering charge-exchange and elastic collisions between ions and neutral water molecules. The model was validated by experiments using the MIPS with two grids having 295 apertures each. The calculated accelerator impingement current agreed reasonably well with those obtained from the experiment. To increase its low propellant utilization efficiency, a new grid design method was introduced. This successfully improved neutral confinement by decreasing the diameter of the accelerator grid apertures without causing direct impingement of ions by optimizing the grid thickness distribution using the Particle In Cell-Monte Carlo Collision (PIC-MCC) calculation results.</p>

<p> 宇宙太陽光発電においては大量の物資輸送が求められ，そのためには膨大な推進剤が消費される．現在の主流となっているキセノンは高価であるため，現実的ではない．そこで我々は，常温常圧で固体であり昇華性を持つ物質を推進剤として使用することを提案する．固体推進剤を使用することにより，高圧タンクやシステム全体の大幅な削減も期待でき，より低価格の軌道間輸送システムが構築可能となる．昇華性推進剤が代替燃料となり得るのか，イオンエンジンを用いて性能を評価した結果を報告する．</p>

<p> SSPSの経済的成立性を左右する要因の一つはSSPSの輸送コストである．本稿では化学推進を用いた打ち上げロケットと電気推進を用いた軌道間輸送機によりSSPSを静止軌道に輸送する場合のコストについて議論する．現状の技術レベルと航空宇宙産業における経験曲線から予測されるSSPSの輸送コストは1万トンのSSPSを10年で構築するのに約1兆円かかり，10基のSSPSを輸送することで約5000憶円となる．このSSPSの輸送にはアルゴンを推進剤とした100 kW級の電気推進機の開発が不可欠である．</p>

<p>地球近傍の軌道間輸送はこれまで化学推進が唯一の選択肢であったが，電気推進の実用化が進むとともに，大電力衛星バスの普及によって十分な電力を確保することが可能となったことから，電気推進を人工衛星の位置や姿勢制御ばかりでなく，軌道移行にも用いる全電化推進が現実のものとなってきている．本稿では，全電化推進を用いたミッションの具体例として，今後も輸送需要が見込まれる静止衛星の軌道投入をはじめ，日本独自のシステムである準天頂衛星や射点の制約からドッグレッグターンをせざるをえない太陽同期軌道への衛星投入を扱った．また，大量の物資の輸送が必要となる宇宙太陽発電衛星の構築について，全電化することによる輸送コストの低減効果の評価を行った．全電化推進は必ずしも短期間の軌道間輸送には向いていないが，適切な条件設定によって，化学推進を単独で用いるよりも大きなメリットをもたらすことが期待できる．</p>

A micro-thruster using stacked solid propellant pellets for nano-satellites has the high versatility. This thruster can generate arbitrary total impulse by changing the number of installed pellets. In order to install the thruster in a nano-satellite, it is necessary to clarify conditions of stable operation. We changed materials and manufacturing methods of the thruster, and measured the combustion chamber pressure and the burning rate. A certain condition of the thruster achieved stable operation. A difference between the burning rates of the propellant pins caused unexpected combustion chamber pressure rise. A silicone rubber contributed stable operation because it made the burning rates of the propellant pins even.

<p> SSPSの経済的成立性を左右する要因の一つは輸送コストである．本論文では，部分再使用打ち上げロケット（GLV）と電気推進を用いた完全再使用軌道往還機（OTV）によりSSPSを輸送するシナリオに基づいてそのコスト計算を行う．現状の技術レベルとその延長で実現可能なSSPSの輸送コストは，10000 tonのSSPSを10年で構築するのに約1.27兆円であり，アルゴンを推進剤とした比推力2660 sの電気推進で実現される．一方，GLVのみでは44%高いコストとなった．感度解析を行うことでSSPSの輸送コストにインパクトを与えるパラメータを抽出し，SSPSの輸送コストを3000億円にするのに必要な技術革新として，GLVおよびOTVの初期製造コストを75%削減することが必要であることが明らかとなった．</p>

The burning rates of small sized Boron Potassium Nitrate (BKNO₃) pellets (3.2-mm-dia., B:KNO₃: binder = 28:70:2) under low pressures were experimentally measured; they increased with pressure (r = 71.1p^0.589[mm], p[MPa]) at an ambient pressure higher than 5 kPa and exhibited little dependence on pressure (r = 3.0±0.5mm/s) at an ambient pressure lower than 5 kPa. The measured pressure increase after reaction becomes smaller under 1 kPa and the photographs taken by a high-speed camera showed unreacted particle emissions at 0.2kPa. These results suggest that the reaction of BKNO₃ at low pressures takes place in condensed phase. From thrust stand measurement, the pressure working on the combustion surface is estimated to be 1.2kPa, which implies that the dependence of burning rates on ambient pressure vanishes under around 1kPa.

<p>A micro-solid rocket as the propulsion system for 1&ndash;10 kg-class micro-spacecraft is proposed here. The micro-solid rocket uses a boron/potassium nitrate pellet as propellant and its total impulse is about 1.5 Ns. Higher total impulse is needed for a propulsion system on small spacecraft to perform advanced space missions such as sample return, formation flight, and active debris removal. To increase the total impulse, it is necessary to increase the propellant mass. However, there is a difficulty in producing new sizes of solid propellant. The author designed a 20&ndash;50 Ns-class micro-solid rocket which uses a stack of existing multiple B/KNO₃ pellets. The side of the propellant pellets was sealed with epoxy resin to prevent an abnormal combustion chamber pressure rise. As a result, all the propellant was burned without an abnormal pressure rise in all combustion tests.</p>

多大な時間とコストを必要とするイオンエンジンの実時間寿命評価試験を数値解析により支援するため,宇宙航空研究開発機構(JAXA:Japan Aerospace Exploration Agency)では,国内の大学と協力し,イオン加速グリッド耐久認定用数値解析JIEDI(JAXA's Ion Engine Development Initiative)ツールの研究開発を実施してきた.本章では,小惑星探査機「はやぶさ」に搭載されたμ10イオンエンジンの地上における実時間の耐久試験結果とJIEDIツールによる損耗計算結果の比較評価について説明し,JIEDIツールが各グリッドの損耗形状の傾向を良く再現し,損耗質量に関しても,実験の計測誤差の範囲で実験結果を再現できることを示した.また,JIEDIツールによる設計支援の実例として,JAXAにおいて新規開発が行われているμ20イオンエンジンにおける,パラメトリックな損耗解析による寿命を考慮した加速グリッド電圧の設計結果について説明し,μ20イオンエンジンにおいてスクリーングリッド電圧を2300Vに設定した場合,電子逆流およびグリッドの構造破壊に陥ることなく,2万時間の要求作動時間を達成するには,加速グリッド電圧を-350Vに設定する必要があることを示した.

Numerical studies were conducted on a small ion engine system called μ1, which was developed for micro-spacecraft propulsion. For its miniaturized application, the μ1 ion engine is very small in size and its beam current and propellant utilization efficiency are lower than those of standard-sized ion engines. As it operates in a low perveance range with a molybdenum two-grid system, the acceleration grid experiences very severe erosion, and the extraction ion beam performance degrades substantially because of the scattering of highly diverted ions and neutrals and the impact of backstreaming charge exchange ions. In this study, the decrease in the propellant utilization efficiency and the lifetime of the μ1 ion engine were firstly assessed using the JIEDI tool, considering the effect of the change in extraction ion beam performance. Lifetime analysis showed that the μ1 grid system can operate over its required lifetime of 10000 h, however, the propellant utilization efficiency decreased significantly from 0.42 at the beginning of life to 0.30 at the grid structural failure of 27000 h. The agreement of grid erosion patterns after the first 150-h of operation between the simulation and the experiment supports the validity of the analysis. In addition to the lifetime assessment, the grid parameters were parametrically searched to achieve both high propulsion performance and long life. One of the strategies is to double the diameters of the screen and accelerator grid holes as well as the accelerator grid thickness, keeping the separation distance between the screen and accelerator grids unchanged.

Cost evaluation of a solar power satellite (SPS) in-space transportation using Hall thruster propulsion systems is conducted to obtain specific characteristics of Hall thrusters that yield significant transportation cost reduction. Therefore, the transportation scenario is optimized in the first place: the choice of reusing or disposing orbit transfer vehicles (OTVs), and the power of the propulsion system. The result shows that the case of reusing OTVs is superior to the case of disposing them, because the OTVs manufacturing costs are predominant in the total cost. In addition, the power has only limited influence on the cost. Further, specific characteristics of Hall thrusters are discussed, which are necessary to achieve a target cost of $3.75 billion (300 billion yen) for the sum of two costs: the in-space transportation cost from a low Earth orbit (LEO) to a geostationary Earth orbit (GEO) and the propellant launch cost from Earth to LEO.