研究者業績
基本情報
- 所属
- 国立研究開発法人宇宙航空研究開発機構 宇宙科学研究所 宇宙飛翔工学研究系 教授総合研究大学院大学 先端学術院 宇宙科学コース 教授
- 学位
- 博士（工学）(1995年3月 東京大学)
- J-GLOBAL ID
- 200901056190267532
- researchmap会員ID
- 1000253787
- 外部リンク
研究キーワード
5研究分野
3主要な経歴
15-
2019年 - 現在
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2020年 - 2024年6月
学歴
2-
- 1995年
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- 1990年
委員歴
6-
2020年 - 2023年3月
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2014年 - 2022年3月
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2014年 - 2015年
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2014年 - 2015年
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2011年 - 2015年
受賞
3-
2014年
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2012年
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1999年
論文
290-
Journal of Propulsion and Power 1-10 2024年6月20日A coupled cylindrical rotating detonation engine (RDE) with two cylindrical RDEs (both combustors had a combustor inner diameter of 23 mm and an axial length of 30 mm) placed next to each other was tested for rocket clustering application. The objective of the experiment was to achieve two-engine synchronized initiation with a single igniter. Experiments were conducted on the inner wall of the combustors with different connecting-hole diameters and wall heights to evaluate the ignition delay time, combustion mode, and propulsion performance. The propellants were gaseous ethylene and oxygen, and experiments were conducted under constant conditions of mass flow rate ([Formula: see text]), equivalence ratio ([Formula: see text]), and backpressure (approximately 10 kPa). When the two combustion chambers were completely separated by a wall, ignition occurred with a time delay of 260 ms in the chamber without an igniter. However, when a large hole ([Formula: see text] diameter) was placed in the wall separating the two combustion chambers, synchronous initiation was successful. Synchronous initiation was also successful when the wall height was lowered (7-mm height). Under both conditions, the same level of specific impulse was achieved as for RDEs operating at the same mass flux.
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Combustion and Flame 264 113443-113443 2024年6月
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Journal of Spacecraft and Rockets 1-11 2024年5月28日There are few experimental studies on rotating detonation engines (RDEs) with liquid propellants. This study reveals the static thrust performance of a cylindrical RDE with ethanol and liquid nitrous oxide as propellants under atmospheric pressure. This RDE had an inner diameter of 40 mm, a maximum combustor length of 230 mm, a nozzle contraction ratio of 1.7, and a nozzle expansion ratio of 9.1. Nineteen experiments were conducted at total mass flow rates of [Formula: see text], mixture ratios of 3.6–5.9, and combustion pressures of 0.35–0.46 MPa, resulting in a maximum detonation velocity of [Formula: see text] (approximately 80% of the theoretical detonation velocity, [Formula: see text]), maximum thrust at sea level of 294 N, and maximum specific impulse at sea level of 148 s. In addition, the maximum characteristic exhaust velocity, [Formula: see text], was [Formula: see text], which was 99% of the theoretical value. The characteristic length of the combustion chamber at this time was 0.15 m. Since conventional rocket combustion requires 1.57 m to achieve the same [Formula: see text] efficiency, this study shows that detonation combustion can reduce the combustor size by 88%.
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Journal of Evolving Space Activities 71(2) 67-77 2024年3月 査読有り
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Proceedings of the Combustion Institute 40(1-4) 105735-105735 2024年
MISC
206-
49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference 2013年 査読有りA magnetoplasma sail (MPS) spacecraft produces an artificial magnetosphere to reflect the solar wind particles approaching the coil, and the corresponding repulsive force exerts on the coil to accelerate the spacecraft in the solar wind direction. In this paper, numerical study of plasma equilibrium in an artificial magnetosphere in interplanetary space is updated to check if the idea of plasma equilibrium is applicable to for MPS or not. It is numerically shown that releasing a low-velocity plasma from an MPS spacecraft excites an equatorial ring-current, which makes a larger magnetosphere and correspondingly a larger thrust level becomes possible. Thrust gain, which is defined as a thrust ratio between MPS and pure magnetic sail without releasing plasma, was found to be as much as 40; this thrust gain is predicted from a limited model describing the interaction between a dipole magnetic field and ions. In addition to the limited simulation, some full numerical simulations of MPS, including a solar wind to magnetosphere interaction as well as plasma equilibrium in a magnetosphere, were conducted to indicate a thrust gain as much as 3.77 is possible in an MHD regime.
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流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM) 44th-2012 ROMBUNNO.1D01 2012年
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Proceedings of the International Astronautical Congress, IAC 10 7979-7985 2012年 査読有り
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48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2012 2012年 査読有り
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日本物理学会講演概要集 67 126-126 2012年
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Proceedings of the Plasma Conference 2011 24E12 2011年9月 査読有り
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低温工学・超電導学会講演概要集 = Meetings of Cryogenics and Superconductivity,84, p. 183 2011年5月
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流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム講演集(CD-ROM) 43rd-2011 ROMBUNNO.2C02 2011年
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AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit 47th Vol.6 4857-4863 2011年
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2011 30th URSI General Assembly and Scientific Symposium, URSIGASS 2011 2011年In order to develop completely electrodeless next generation plasma thrusters for deep space missions, we have initiated the HEAT (Helicon Electrodeless Advanced Thruster) project. In our scheme, source plasmas are generated by means of the highly efficient helicon-wave discharge they are then electromagnetically accelerated using external antennas to yield a thrust. The entire process can be achieved without using any eroding electrodes, leading to plasma thrusters of a limitless lifetime. © 2011 IEEE.
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低温工学・超電導学会講演概要集 = Meetings of Cryogenics and Superconductivity,83, p. 248 2010年12月
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VACUUM 85(5) 574-578 2010年11月Thrust performance tests were conducted for an arrayed multi-plasmajet structure consisting of rectangular nozzle elements each with an exit height of 05 mm and length of 05 mm To evaluate the thrust characteristics of the arrayed micro-multi-plasmajet the thrust was measured using a calibrated cantilever-type thrust stand in vacuum Using 3 x 3 nozzle elements the micro-plasmajet array showed stable DC operation Compared to conditions with cold-gas flow the application of a DC discharge showed a significant improvement in thrust performance of at least 20% for thrust and specific impulse Typical values achieved for thrust specific impulse and thrust efficiency at an Input power of 6 3 W were 8 5 mN 77 s and 021 respectively (C) 2010 Elsevier Ltd All rights reserved
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IEEE TRANSACTIONS ON ANTENNAS AND PROPAGATION 58(10) 3282-3288 2010年10月The ionized exhaust plumes of solid rocket motors may interfere with RF transmission under certain flight conditions. To understand the important physical processes involved, we measured microwave attenuation and phase delay due to the exhaust plume during sea-level static firing tests for a full-scale solid propellant rocket motor. The measured data were compared with the results of a detailed simulation performed using the frequency-dependent finite-difference time-domain (( FD)(2)TD) method. The numerically derived microwave attenuation was in good agreement with experimental data. The results revealed that either the line-of-sight microwave transmission through ionized plumes or the diffracted path around the exhaust plume mainly affects the received RF level, which depends on the magnitude of the plasma-RF interaction.
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CLASSICAL AND QUANTUM GRAVITY 27(8) 084010 2010年4月A space gravitational-wave antenna, DECIGO (DECI-hertz interferometer Gravitational wave Observatory), will provide fruitful insights into the universe, particularly on the formation mechanism of supermassive black holes, dark energy and the inflation of the universe. In the current pre-conceptual design, DECIGO will be comprising four interferometer units; each interferometer unit will be formed by three drag-free spacecraft with 1000 km separation. Since DECIGO will be an extremely challenging mission with high-precision formation flight with long baseline, it is important to increase the technical feasibility before its planned launch in 2027. Thus, we are planning to launch two milestone missions. DECIGO pathfinder (DPF) is the first milestone mission, and key components for DPF are being tested on ground and in orbit. In this paper, we review the conceptual design and current status of DECIGO and DPF.
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宇宙航空研究開発機構研究開発報告 9 1-9 2010年2月Improvements have been made to the grid erosion evaluation code (JIEDI-1) to achieve robust convergence and to enhance the accuracies of grid life estimations. The major improved features of the code are 1) the use of Newton-Raphson method and the ICCG solver, 2) modifications of the inlet boundary conditions, 3) incorporation of the differential sputtering yield and redeposition effect of the grid material, and 4) solution-adaptive mesh generation. The last feature is realized by the commercial mesh generation software "pro-STAR". Numerical code prediction of grid mass loss showed good agreement with that of μ10 ion engine life test, with robust convergence and improved simulation time by a factor of 3〜4 in comparison with the previous code.
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2007/2008年度 宇宙関連プラズマ研究会論文集 58-66 2010年2月 査読有り
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航空原動機・宇宙推進講演会講演論文集(CD-ROM) 50th ROMBUNNO.AJCPP2010-135 2010年
主要な書籍等出版物
6講演・口頭発表等
561-
第63回航空原動機・宇宙推進講演会／北部支部2024年講演会ならびに第5回再使用型宇宙輸送系シンポジウム 日本航空宇宙学会
共同研究・競争的資金等の研究課題
28-
日本学術振興会 科学研究費助成事業 基盤研究(S) 2020年8月 - 2025年3月
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日本学術振興会 科学研究費助成事業 特別推進研究 2019年4月 - 2024年3月
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日本学術振興会 科学研究費助成事業 基盤研究(B) 2018年4月 - 2021年3月
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日本学術振興会 科学研究費助成事業 基盤研究(B) 2017年4月 - 2021年3月
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日本学術振興会 科学研究費助成事業 挑戦的研究(萌芽) 2019年6月 - 2021年3月
