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 物理科学研究科 宇宙科学専攻 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
4研究分野
3主要な経歴
18-
2020年 - 現在
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2019年 - 現在
学歴
2-
- 1995年
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- 1990年
委員歴
6-
2020年 - 現在
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2014年 - 現在
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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
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Joint Symposium: 32nd ISTS & 9th NSAT ISTS-2019-b-006 2019年6月
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Joint Symposium: 32nd ISTS & 9th NSAT ISTS-2019-b-003 2019年6月
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Transactions of the Japan Society for Aeronautical and Space Sciences, Aerospace Technology Japan 17(4) 455-460 2019年3月 査読有り
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Journal of Propulsion and Power 35(2) 2019年3月 査読有り
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Journal of Propulsion and Power 36 2019年 査読有り
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AIAA Scitech 2019 Forum 2019年© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. The relationship between the thrust performance and the cathode temperature in a megawatt-class, self-field magnetoplasmadynamic (MPD) thruster was investigated experimentally and numerically. For various propellants, i.e. argon, hydrogen, nitrogen, and helium, the thrust performance and cathode temperature were measured at discharge currents ranging from 5 to 12 kA. Measured thrust and thrust efficiency increased with the discharge current. For hydrogen propellant, the highest thrust and thrust efficiency of 28 N and 30%, respectively, were attained at a mass flow rate of 0.4 g/s and a discharge current of 12 kA. Cathode surface temperature also increased with the discharge current. For the hydrogen propellant, the tip of the cathode was particularly heated and the temperature exceeded 4000 K. On the other hand, for the argon and helium propellants, the cathode was heated relatively entirely. Numerical results showed that the current density at the cathode tip increased significantly at high discharge currents because of high hall parameter. This can be a main reason why the cathode surface was heated particularly near the tip for the hydrogen propellant.
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AIAA Scitech 2019 Forum 2019年© 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Thrust measurements of rotating detonation engine of (1) ethylene / gas-oxygen and (2) methane / gas-oxygen with various throat geometries in a vacuum chamber to simulate different back-pressure conditions ranging from 1.1-104 kPa were conducted. For throatless rotating detonation engine, we defined equivalent throat area as the detonation channel area, and then tested four nozzle contraction ratios of 1, 1.5, 2.5, and 8. Engines could be successfully ignited by electric ignitors when initial pressure was high enough to have, at least, one detonation cell in RDE channel. We measured the combustor pressure and reveled that it was almost proportional to the throat mass flux regardless of contraction ratios and the propellant combinations. The specific impulse of methane / gas-oxygen case could achieve 84 ± 1% of ideal specific impulse at the optimum expansion for each back pressure.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 62(1) 11-19 2019年 査読有りNumerical simulation of plasma flow and the self-heating characteristics of a LaB6 hollow cathode were performed using a hybrid-PIC model. For a discharge current of 30 A and mass flow rate of 3 mg/s, the influences of an emitter temperature profile and model parameter included in an anomalous resistivity model on the plasma flow and energy flux were investigated. In the simulation, the discharge voltage was fixed at a predetermined value and the maximum emitter temperature was periodically adjusted to keep the discharge current constant. The results show that the present model predicts the keeper floating voltage within an accuracy of 20%. It is found that the main reason for the emitter temperature to rise is due to ion bombardment and accompanying recombination energy, and that the maximum emitter temperature can be kept lower as the emitter temperature profile becomes uniform. It is also shown that thermal input into the emitter is decreased when anomalous resistivity increases.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 17(2) 203-210 2019年 査読有り<p>A lanthanum hexaboride (LaB6) hollow cathode with radiative heater has been developed for high power Hall thruster neutralizer. The influences of orifice and keeper shapes on the performance of the hollow cathode and on its operating characteristics are investigated by measuring the discharge voltage and current. Six different orifice shapes are used in this study: straight, long tapered, and short tapered, each with an orifice diameter of 2 or 3 mm. The straight orifice with the 2-mm diameter has the widest range of spot-mode operation, although no effect of orifice shape is observed for the 3-mm diameter. This is because the high neutral-gas density around the orifice fosters transition to the spot mode from the plume mode with a straight Φ2-mm orifice. Regarding the effect of the keeper, having a large exit diameter and the shortest distance possible between the cathode tube and the keeper gives the widest range of spot-mode operation.</p>
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JOURNAL OF PROPULSION AND POWER 35(1) 213-223 2019年1月
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Journal of Propulsion and Power 35(1) 2019年1月 査読有り
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69th International Astronautical Congress (IAC2018) 2018-October IAC-18-C4.7-C3.5.11 2018年10月Copyright © 2018 by the International Astronautical Federation (IAF). All rights reserved. Magnetic Sail and MagnetoPlasma Sail(=MPS) are new generation space propulsion systems that produce the thrust utilizing the interaction between magnetic field formed by an on-board superconducting coil and the solar wind. MPS is a system that generates larger thrust than Magnetic Sail by magnetic field inflated by plasma injection from the spacecraft. These systems are expected to realize a long-time space transfer with extremely smaller quantity of propellant than other plasma propulsion systems, because these systems convert the momentum of solar wind into thrust. It is expected that the magnetic Reynolds number(=Rm) that indicates the strength of the diffusion effect of the magnetic field is low around the spacecraft of MPS. In addition, Rm depends on the local plasma parameters. Objectives of this study is to investigate the influence of the mass flow rate of plasma on the Rm distribution and thrust characteristics when it is considered that neutral particle effect in injected plasma. Our results indicate that Rm is less than 10 in the large part of magnetosphere when the neutral particle diffusing distribution is assumed. The peak of the thrust gain which is about 2.1 in the ideal condition decreased to about 1.8 by the influence of Rm. In addition, our results also indicate that there is an optimum mass flow rate with the maximum thrust in each condition. In considering neutral particle effect, this optimum mass flow rate is higher than that of the ideal condition.
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Proceedings of the Combustion Institute 2018年7月17日 査読有り
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AIAA Propulsion and Energy Forum AIAA-2018-4510 2018年7月
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AIAA Propulsion and Energy Forum AIAA-2018-4815 2018年7月
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AIAA Propulsion and Energy Forum AIAA-2018-4513 2018年7月© 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Hybrid-PIC simulation of a hollow cathode, in which heavy species and electrons are treated as particles and fluid, respectively, is coupled with growth model of ion acoustic turbulence(IAT), and the result is compared with the conventional model which assumes that the IAT reaches saturated stage. The result shows that the growth of IAT requires several millimeters in the orifice region, which indicates the importance of using growth model. It is also found that obvious oscillation of the space potential in the plume region emerges by using the growth model. We suggested an ion heating model applicable to the Hybrid-PIC model, and shows that the high energy ions can be reproduced by using the suggested model.
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Journal of Applied Physics 123(15) 2018年4月21日 査読有りIt is quite a challenge to design low power Hall thrusters with a long lifetime and high efficiency because of the large surface area to volume ratio and physical limits to the magnetic circuit miniaturization. As a potential solution to this problem, we experimentally investigated the external discharge plasma thruster (XPT). The XPT produces and sustains a plasma discharge completely in the open space outside of the thruster structure through a magnetic mirror configuration. It eliminates the very fundamental component of Hall thrusters, discharge channel side walls, and its magnetic circuit consists solely of a pair of hollow cylindrical permanent magnets. Thrust, low frequency discharge current oscillation, ion beam current, and plasma property measurements were conducted to characterize the manufactured prototype thruster for the proof of concept. The thrust performance, propellant ionization, and thruster erosion were discussed. Thrust generated by the XPT was on par with conventional Hall thrusters [stationary plasma thruster (SPT) or thruster with anode layer] at the same power level (∼11 mN at 250 W with 25% anode efficiency without any optimization), and discharge current had SPT-level stability (Δ < 0.2). Faraday probe measurements revealed that ion beams are finely collimated, and plumes have Gaussian distributions. Mass utilization efficiencies, beam utilization efficiencies, and plume divergence efficiencies ranged from 28 to 62%, 78 to 99%, and 40 to 48%, respectively. Electron densities and electron temperatures were found to reach 4 × 1018 m-3 (- n e / n e = ±52%) and 15 eV (- T e / T e = ±10%-30%), respectively, at 10 mm axial distance from the anode centerline. An ionization mean free path analysis revealed that electron density in the ionization region is substantially higher than the conventional Hall thrusters, which explain why the XPT is as efficient as conventional ones even without a physical ionization chamber. Our findings propose an alternative approach for low power Hall thruster design and provide a successful proof of concept experiment of the XPT.
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Journal of Applied Physics 123(15) 153302 2018年4月21日 査読有りIt is quite a challenge to design low power Hall thrusters with a long lifetime and high efficiency because of the large surface area to volume ratio and physical limits to the magnetic circuit miniaturization. As a potential solution to this problem, we experimentally investigated the external discharge plasma thruster (XPT). The XPT produces and sustains a plasma discharge completely in the open space outside of the thruster structure through a magnetic mirror configuration. It eliminates the very fundamental component of Hall thrusters, discharge channel side walls, and its magnetic circuit consists solely of a pair of hollow cylindrical permanent magnets. Thrust, low frequency discharge current oscillation, ion beam current, and plasma property measurements were conducted to characterize the manufactured prototype thruster for the proof of concept. The thrust performance, propellant ionization, and thruster erosion were discussed. Thrust generated by the XPT was on par with conventional Hall thrusters [stationary plasma thruster (SPT) or thruster with anode layer] at the same power level (∼11 mN at 250 W with 25% anode efficiency without any optimization), and discharge current had SPT-level stability (Δ < 0.2). Faraday probe measurements revealed that ion beams are finely collimated, and plumes have Gaussian distributions. Mass utilization efficiencies, beam utilization efficiencies, and plume divergence efficiencies ranged from 28 to 62%, 78 to 99%, and 40 to 48%, respectively. Electron densities and electron temperatures were found to reach 4 × 1018 m-3 (- n e / n e = ±52%) and 15 eV (- T e / T e = ±10%-30%), respectively, at 10 mm axial distance from the anode centerline. An ionization mean free path analysis revealed that electron density in the ionization region is substantially higher than the conventional Hall thrusters, which explain why the XPT is as efficient as conventional ones even without a physical ionization chamber. Our findings propose an alternative approach for low power Hall thruster design and provide a successful proof of concept experiment of the XPT.
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Asian Joint Conference on Propulsion and Power AJCPP2018-027 2018年3月
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2018 AIAA AEROSPACE SCIENCES MEETING (SciTech 2018) AIAA-2018-0157 2018年1月
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 16 274-279 2018年 査読有り
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Journal of Propulsion and Power 34(4) 1093-1096 2018年 査読有りAn external discharge plasma thruster was proposed, and design details of the first prototype were described to address efficient scaling-down and lifetime problems of low-power Hall thrusters. A conductive material deposition was observed on both the front wall and the anode. The deposition on the front wall was much denser near the inner and outer edges of the anode. The bright plasma was located downstream of the anode front surface by approximately 1–3 mm, depending on the location. This anode dark space and the material deposition provide support for erosion-free operation. Preliminary experiments results suggest that the external discharge plasma thruster (XPT) has discharge characteristics similar to an Stationary Plasma Thruster (SPT). Performance of the XPT is comparable to an SPT at the same power level.
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International Journal of Modern Physics D (ja) 2018年 査読有り
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Journal of Propulsion and Power 34(4) 1093-1096 2018年 査読有りAn external discharge plasma thruster was proposed, and design details of the first prototype were described to address efficient scaling-down and lifetime problems of low-power Hall thrusters. A conductive material deposition was observed on both the front wall and the anode. The deposition on the front wall was much denser near the inner and outer edges of the anode. The bright plasma was located downstream of the anode front surface by approximately 1–3 mm, depending on the location. This anode dark space and the material deposition provide support for erosion-free operation. Preliminary experiments results suggest that the external discharge plasma thruster (XPT) has discharge characteristics similar to an Stationary Plasma Thruster (SPT). Performance of the XPT is comparable to an SPT at the same power level.
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Publications of the Astronomical Society of Australia 1-17 2018年 査読有り
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JOURNAL OF PROPULSION AND POWER 34(1) 116-124 2018年1月 査読有りA pulse-detonation thruster can generate a high-repeatability small impulse at a high operating frequency. To operate a pulse-detonation cycle in a vacuum environment without a purging material, a liquid-purge method proposed by Matsuoka et al. ("Development of a Liquid-Purge Method for High-Frequency Operation of Pulse Detonation Combustor," Combustion Science and Technology, Vol. 187, No. 5, 2015, pp. 747-764) and throat at the exit of the combustor were introduced. In the demonstration experiment, gaseous ethylene-liquid nitrous oxide, as detonable mixture, and a throat having an inner diameter of 3.6 mm (blockage ratio = 87%) were used. The measured cyclic flame-propagation speeds were 2005 +/- 90 m/s and 92 +/- 4% of the estimated Chapman-Jouguet detonation speed. Consequently, a 50 Hz pulse-detonation operation without a purging material in the ambient-pressure range of 0.035-1.5 kPa was confirmed. A quasi-steady model, in which gases in the combustor are in the stationary state, was newly developed to investigate the operating characteristic of the pulse-detonation thruster. The experimental pressure history in the combustor during the burned-gas blowdown process was in good agreement with that of the model. Moreover, using the model, the thrust performance of the pulse-detonation thruster with a converging-diverging nozzle was investigated. It was found that the estimated specific impulse was comparable with that of the theoretical steady-state rocket engine.
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Aerospace Technology Japan 16 98-104 2018年 査読有り
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Journal of Physics: Conference Series 840(1) 2017年6月1日 査読有り© Published under licence by IOP Publishing Ltd. DECIGO (DECi-hertz Interferometer Gravitational wave Observatory) is the planned Japanese space gravitational wave antenna, aiming to detect gravitational waves from astrophysically and cosmologically significant sources mainly between 0.1 Hz and 10 Hz and thus to open a new window for gravitational wave astronomy and for the universe. DECIGO will consists of three drag-free spacecraft arranged in an equilateral triangle with 1000 km arm lengths whose relative displacements are measured by a differential Fabry-Perot interferometer, and four units of triangular Fabry-Perot interferometers are arranged on heliocentric orbit around the sun. DECIGO is vary ambitious mission, we plan to launch DECIGO in era of 2030s after precursor satellite mission, B-DECIGO. B-DECIGO is essentially smaller version of DECIGO: B-DECIGO consists of three spacecraft arranged in an triangle with 100 km arm lengths orbiting 2000 km above the surface of the earth. It is hoped that the launch date will be late 2020s for the present..
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AIAA SciTech Forum - 55th AIAA Aerospace Sciences Meeting 2017年Study of a Rotating Detonation Engine (RDE) has been carried out in many research institutions. The RDE has the advantage of high efficiency, simple structure and short combustor length. Toward the practical use for rocket engines, the evaluation of thrust efficiency at vacuum condition is necessary. In addition, it’s necessary to find a cooling method to endure the high heat load by detonation combustion. In this study, ethylene - oxygen mixtures are used as propellant, and we applied C/C composite, which is a heat resistant material, to a RDE and carried out a long-duration combustion test at sea level. As a result, we succeeded in the long-duration rotating detonation engine combustion demonstration of 6- 10 s at sea level for the first time in the world as rocket engine use, and the maximum thrust and the maximum specific impulse were achieved 301 N and 144 s.
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AIAA SciTech Forum - 55th AIAA Aerospace Sciences Meeting 2017年A rotating detonation turbine engine has simpler structure and higher thermal efficiency than a conventional gas turbine engine. In this study, we designed a rotating detonation turbine engine with single stage centrifugal compressor, combustion chamber, and single stage radial flow turbine which are placed on one side of rotor disk. We performed cold flow experiments and combustion experiments. As results of cold flow test, compressor can supply air 52 g/s. In combustion experiments, we supply pressured oxidizer and fuel from outside of engine. As a results, combustion propagation speed was 600 - 1300 m/s. this is about 25 – 45 % of Chapman – Jouget detonation velocity. In addition, rotational speed rises by 160 rpm during combustion.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 60(5) 320-326 2017年 査読有りThree different types of high power Hall thrusters-anode layer type, magnetic layer type with high specific impulse, and magnetic layer type with dual mode operation (high thrust mode and high specific impulse mode)-have been developed, and the thrust performance of each thruster has been evaluated. The thrust of the anode layer type thruster is in the range of 19-219 mN, with power in the range of 325-4500 W. The thrust of the high specific impulse magnetic layer type thruster was 102 mN, with specific impulse of 3300 s. The thrust of the bimodal operation magnetic layer thruster was 385mN with specific impulse of 1200 s, and 300mN with specific impulse of 2330 s. The performance of these thrusters demonstrates that the Japanese electric propulsion community has the capability to develop a thruster for commercial use.
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 60(2) 67-76 2017年 査読有りA sensitivity analysis of a fully kinetic particle code was conducted to investigate the importance of uncertainties associated to physical parameters. A 500 W-class laboratory model magnetic-layer Hall thruster was used as the testbed. The sensitivities of the physical parameters, including thermal accommodation coefficient, anode/wall temperature, Bohm diffusion coefficient, electron injection current, cathode coupling voltage, and background pressure, were quantified one-byone on a conservative possible range. The results suggest the wall erosion prediction is more sensitive to the physical parameters than the thrust or the discharge current. Among the physical parameters, sensitivity to the Bohm diffusion coefficient and parameters related to the neutral flow (i.e., thermal accommodation coefficient and anode/wall temperatures) were dominant. It was hence found that uncertainties in the physical parameters related to the neutral flow had comparable influence on the Bohm diffusion coefficient despite the low attention they attracted.
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JOURNAL OF PROPULSION AND POWER 33(1) 80-88 2017年1月 査読有りThe rotating detonation engine is a propulsion system that obtains thrust using continuously existing detonation waves. A rotating detonation combustor usually has an annular shape that allows detonation waves to propagate in the circumferential direction. In this study, we used a disk-shaped rotating detonation combustor with a combustion chamber with flat-plane glass walls to observe the structure of the phenomena. Self-luminescence, shadowgraphs, and schlieren visualization experiments were performed and compared. Results revealed that detonation waves were propagating in a mixture layer of three gases, fuel, oxidizer, and burned gas at 1600 to 900 m/s; Chapman-Jouguet velocity was 2376 m/s. Waves maintained a three-dimensional complicated wave shape in the disk-shaped combustion chamber with parallel-jet injectors.
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PROCEEDINGS OF THE COMBUSTION INSTITUTE 36(2) 2673-2680 2017年 査読有りThe structure of detonation waves in rotating detonation combustors (RDCs) and their combustion chamber pressure characteristics have not yet been fully clarified due to the complexity and shape of the RDC combustion chamber. Therefore, a disk-shaped RDC was used in this study to visualize the inside of the combustion chamber while simultaneously measuring its pressure. Forward-tilting rotating detonation waves were observed, and a schematic was proposed for them. The initial velocity of the forward-tilting rotating detonation wave was 1200 +/- 160 m/s, and it subsequently increased to 1600 +/- 160 m/s; meanwhile, the Chapman-Jouguet (CJ) velocity was 2376 m/s. There are several reasons why the velocity may have differed so widely from the CJ value, including the presence of burned gas in front of the detonation wave, the complicated wave structure due to non-uniformity of the mixture in the RDC, insufficient propellant mixing, and the difference between the true and actual wave propagation direction. The velocity and amplitude of the combustion chamber static pressure appeared to be correlated. Averaged combustion chamber static pressure reached 0.432 MPa, which was 89.0% and 92.8% of the fuel and oxidizer plenum pressure, respectively. Dynamic pressure was also estimated using an equilibrium calculation. The resulting dynamic pressure was 0.008 MPa, and estimated total pressure was 0.440 MPa; these values were 90.1% and 94.6% of the fuel and oxidizer plenum total pressure, respectively, even though pressure was lost through the small diameter injector holes. (C) 2016 by The Combustion Institute. Published by Elsevier Inc.
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JOURNAL OF PROPULSION AND POWER 32(2) 383-391 2016年3月 査読有りA rotary-valved four-cylinder pulse detonation rocket engine system, Todoroki II, was developed, in which two novel techniques, the use of an inflow-driven motor and an inverted oxidizer cylinder, were introduced. The total length of the system was 1910mm; its total weight when filled with ethylene-nitrous-oxide propellant and helium purge gas was 32.5kg; and the engine weight was 9.6kg. In a ground firing test with a duration of 1500ms, a thrust-to-engine-weight ratio of 2.7 was achieved. Thus, it was demonstrated that a multicylinder pulse detonation rocket engine system can be used as a practical thrust mechanism. Using a launch and recovery system, a flight-simulating test was conducted to evaluate the features and viability of the engine design. The launch and recovery system operated perfectly, and Todoroki II reached a height of about 9.7m. The operation of the pulse detonation rocket engine under conditions simulating real vertical flight without constraint forces with a duration of about 1200ms and a thrust-to-engine-weight ratio of 2.5 was demonstrated. No serious impact of the vibration caused by the pulse detonation rocket engine operation or the rotation of the rotary valve on the flight was observed.
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52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 2016年© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. A laboratory model of a hollow cathode with a radiative heater was built and tested for discharge currents of 10-50 A. Plasma ignitions and also discharge operations were successfully demonstrated even if the heater was not wrapped with an insulator, which will lead to robustness against heater breakdown. The voltage-current characteristic indicates that mode transition occurred between 30 and 40 A for a mass flow rate of 30 sccm. To investigate the flow field, a Hybrid-PIC simulation was conducted for a mass flow rate of 30 sccm and a discharge current of 30 A. The result shows acceptable distributions of the electron density, electric potential, and electron temperature inside and outside the cathode. The keeper’s floating voltage was close to the experimental data, but was slightly higher than experimental data. Changing the parameters of the anomalous resistivity can adjust the keeper’s floating voltage, but it has strong impact on the plasma properties.
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52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 2016年© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In this paper, we present external discharge plasma thruster (XPT), a prototype of erosion free low power Hall thruster which produces and sustains plasma discharge completely outside a cavity. Details of this novel Hall thruster design and preliminary test results are presented. The thrust and the anode specific impulse ranged from 0.5 to 17.4 mN, and from 108 to 1240 sec respectively at anode potentials of 100-150-200-250 V with anode mass flow rates of 0.48-0.95-1.43 mg/s. The anode efficiency ranged from 2.4 to 25.6 % at discharge powers from 11 to 412 W. Preliminary experiment results suggest that XPT has similar discharge characteristics with SPT and TAL. Performance of XPT is comparable to SPT and TAL at the same power level, and very stable operation (Δ<0.2) is possible over wide range of operational conditions.
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52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 2016年© 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Magnetoplasmadynamic (MPD) thruster is a candidate of next generation electric propulsion system for a mission that high thrust is required. We are developing the high power Self-field MPD arcjet of several hundred kW class using a numerical and thermal design tool. We are testing quasi-steady MPD thrusters as the validation tool for numerical tool. In order to validate the numerical simulation or the predict of sheath effect, the electrode temperature measurement is necessary flowing to a great impact on the numerical simulation. This paper reports the cathode temperature measurement of the quasi-steady state MPD thruster with argon propellant in 160-1940 kW range and the cathode temperature using the 2-color pyrometer that we recently developed. The cathode tip temperature is near 3000 K above 13 kA. We reveals the temperature distribution heating only the cathode tip peculiar to the quasi-steady state experiment in lower the theoretical critical current. In addition, the cathode temperature change is clarified of 1 ms quasi-steady operation.
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Aerospace Technology Japan 14 165-171 2016年 査読有り
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Aerospace Technology Japan 14 165-171 2016年 査読有り
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TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 59(2) 100-103 2016年 査読有りA one hundred hour test of a 100 mA class microwave discharge neutralizer was performed to investigate the erosion rate and deposition in the neutralizer. The mass loss of each component (antenna, discharge chamber wall, front yoke, back yoke, insulator and orifice) was measured in diode mode (contact voltage of 50 V) at a constant incident microwave power of 8W and xenon mass flow rate of 49 mu g/s. The measured erosion rates of the discharge chamber and the antenna were 180 mu g/hr and 10 mu g/hr, respectively. The element concentration of the deposited material on the back yoke and insulator were measured; the contamination on the insulator was found to contain 34% copper from the wall and 6% molybdenum from the antenna.
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Physics of Plasma 22(10) 2015年10月 査読有り
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53rd AIAA Aerospace Sciences Meeting 2015年A study on a rotating detonation engines (RDE) having the advantage of high frequency operation (1~100 kHz) and simple structure has been carried out in many research groups for the practical use toward a rocket engine. However, there are many unsolved parts about internal flow of a RDE, and the state of its channel exit has not been solved exactly. In this study, we used gas ethylene and gas oxygen as propellant. A slide mechanism for thrust measurement was developed. We will evaluate and the thrust performance and the state of channel exit of a RDE by the nozzle shape. We performed combustion test of a RDE with a conical-shape tail. As a results, thrust of a RDE with a conical-shape tail was achieved 101 ~ 308 N.
MISC
206主要な書籍等出版物
6講演・口頭発表等
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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
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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
