Updated on 2021/07/24

写真a

 
KINEFUCHI Kiyoshi
 
Organization
Graduate School of Engineering Department of Aerospace Engineering Aerodynamics and Propulsion Associate professor
Graduate School
Graduate School of Engineering
Undergraduate School
School of Engineering Mechanical and Aerospace Engineering
Title
Associate professor
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Degree 1

  1. 博士(工学) ( 東京大学 ) 

Research Interests 5

  1. 航空宇宙工学

  2. 宇宙推進

  3. 衝撃波

  4. 極低温推進

  5. 電気推進

Research Areas 3

  1. Frontier Technology (Aerospace Engineering, Marine and Maritime Engineering) / Aerospace engineering

  2. Manufacturing Technology (Mechanical Engineering, Electrical and Electronic Engineering, Chemical Engineering) / Fluid engineering

  3. Energy Engineering / Fundamental plasma

Research History 3

  1. Nagoya University   Department of Aerospace Engineering, Graduate School of Engineering

    2019.9

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  2. Princeton University   Department of Mechanical and Aerospace Engineering

    2015.2 - 2016.1

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    Country:Japan

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  3. Japan Aerospace Exploration Agency   Space Transportation Mission Directorate

    2003.4 - 2019.8

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    Country:Japan

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Education 3

  1. The University of Tokyo

    2006.10 - 2009.9

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  2. The University of Tokyo

    2001.4 - 2003.3

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    Country: Japan

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  3. Tokyo Institute of Technology   School of Engineering   Dept. of Mechano-Aerospace Engineering

    1997.4 - 2001.3

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    Country: Japan

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Professional Memberships 4

  1. 日本航空宇宙学会

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  2. 日本機械学会

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  3. 日本衝撃波研究会

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  4. American Institute of Aeronautics and Astronautics (AIAA)

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Committee Memberships 7

  1. 日本流体力学会   中部支部運営委員  

    2021.4   

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    Committee type:Academic society

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  2. 日本衝撃波研究会   幹事  

    2020.4   

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    Committee type:Academic society

  3. 日本機械学会   2020年度年次大会実行委員  

    2020   

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    Committee type:Academic society

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  4. Asian Joint Conference on Propulsion and Power (AJCPP)   Organizing Committee Member  

    2010   

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    Committee type:Other

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  5. 日本航空宇宙学会   原動機・推進部門委員  

    2008.4 - 2012.3   

  6. 日本航空宇宙学会   会誌編集委員  

    2008.4 - 2010.3   

  7. 航空原動機・宇宙推進講演会   実行委員  

    2008 - 2012   

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Awards 6

  1. 優秀学生賞

    2021.1   令和2年度宇宙輸送シンポジウム   ドライアイスを推進剤とした電気推進システムの提案と供給実験

    眞木達朗, 杵淵紀世志, 張科寅, 渡邊 裕樹

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  2. Scilight

    2018.10   American Institute of Physics   Numerical supersonic flow simulations help improve high-speed engine design

    Kiyoshi Kinefuchi, Andrey Y, Starikovskiy, and Richard, B. Miles

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    Award type:Honored in official journal of a scientific society, scientific journal  Country:United States

  3. Featured Article, Physics of Fluids

    2018.10   American Institute of Physics   Numerical Investigation of Nanosecond Pulsed Plasma Actuators for Control of Shock-wave/Boundary-layer Separation

    Kiyoshi Kinefuchi, Andrey Y, Starikovskiy, and Richard, B. Miles

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    Award type:Honored in official journal of a scientific society, scientific journal  Country:United States

  4. 宇宙航空研究開発機構理事長賞

    2016   宇宙航空研究開発機構   基幹ロケット高度化の開発完了と飛行実証成功

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  5. Best Paper Award

    2013.10   International Electric Propulsion Conference   Application of the Hollow Cathode to DC Arcjet

    Masahiro Kinoshita, Daisuke Nakata, Kiyoshi Kinefuchi, Hitoshi Kuninaka

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    Award type:Award from international society, conference, symposium, etc.  Country:United States

  6. Best Paper Award

    2011.8   AIAA Liquid Propulsion Technical Committee   Heat Exchange and Pressure Drop Enhanced by Sloshing

    Takehiro Himeno, Daizo Sugimori, Katsutoshi Ishikawa, Yutaka Umemura, Seiji Uzawa, Chihiro Inoue, Toshinori Watanabe, Satoshi Nonaka, Yoshihiro Naruo, Yoshifumi Inatani, Kiyoshi Kinefuchi, Ryoma Yamashiro, Toshiki Morito, Koichi Okita

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    Award type:Award from international society, conference, symposium, etc.  Country:United States

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Papers 123

  1. Neutral atom density measurements of xenon plasma inside a mu 10 microwave ion thruster using two-photon laser-induced fluorescence spectroscopy Reviewed

    Ryudo Tsukizaki, Yusuke Yamashita, Kiyoshi Kinefuchi, Kazutaka Nishiyama

    VACUUM   Vol. 190   2021.8

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:PERGAMON-ELSEVIER SCIENCE LTD  

    This paper reports measurements of the xenon ground state and excited state densities inside a mu 10 microwave ion thruster. This thruster exhibits a 40% thrust enhancement upon changing from the waveguide to the discharge chamber propellant injection mode. In the present work, the associated mechanism was quantitatively evaluated using two-photon laser induced fluorescence (TALIF) spectroscopy to monitor the thruster waveguide. The 834.7 nm emission from excited state xenon was investigated with a 224.3 nm dye laser to excite the Xe I 5p61 S0 6pMODIFIER LETTER PRIME [3/2]2 state, compared with the emission without the laser. The resulting data confirm that the neutral density exhibits a linear relationship with the propellant flow rate in the cold gas and ionized state, while the ion acceleration decreases the neutral density by the same order of magnitude as the propellant utilization efficiency is changed. As the propellant flow rate increases, the collisions of neutrals that generate excited states occur in the waveguide and, when this process plateaus, the ground state emission suddenly increases. Propellant injection from the discharge chamber is evidently effective at suppressing collisions with electrons in the waveguide that generate excited states and that potentially interfere with microwave propagation.

    DOI: 10.1016/j.vacuum.2021.110269

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  2. Performance of a miniature Hall thruster and an in-house PPU Reviewed

    Naoji Yamamoto, Masatoshi CHONO, Ryudo TSUKIZAKI, Takato MORISHITA, Kenichi KUBOTA, Shinatra CHO, Kiyoshi KINEFUCHI, Toru TAKAHASHI

    Transactions of the Japan Society for Aeronautical and Space Sciences   Vol. 64 ( 3 )   2021.5

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  3. Design and testing of additively manufactured high-efficiency resistojet on hydrogen propellant Reviewed

    Giulio Coral, Kiyoshi Kinefuchi, Daisuke Nakata, Ryudo Tsukizaki, Kazutaka Nishiyama, Hitoshi Kuninaka

    Acta Astronautica   Vol. 181   page: 14 - 27   2021.4

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    Language:Japanese   Publishing type:Research paper (scientific journal)  

    This paper presents the design methodology and performance testing of an additively manufactured resistojet operating on hydrogen as propellant. Additive manufacturing allows to produce complex monolithic resistors, resulting in reliable high efficiency thrusters. The concept, to be used in combination with advanced cryogenic storage technologies, is proposed for short time and high specific impulse orbit transfers. The simplified two-dimensional thermal design approach adopted is discussed, and its application to the engineering of the resistor is shown for both Inconel 718 and tungsten. The paper reports the performance testing of the proof-of-concept version of the thruster, manufactured in Inconel 718. Experiments on hydrogen show near ideal performance, demonstrating a peak thermal efficiency of 96%. The thruster proves the validity of the design methodology proposed, and the feasibility of the approach to develop monolithic additively manufactured hydrogen resistojets as main propulsion units.

    DOI: 10.1016/j.actaastro.2020.12.047

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  4. Speckle beam-oriented schlieren technique Reviewed

    Yusuke Nakamura, Takumi Suzuki, Kiyoshi Kinefuchi, Akihiro Sasoh

    Experiments in Fluids   Vol. 62 ( 1 )   2021.1

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    Language:Japanese   Publishing type:Research paper (scientific journal)  

    Abstract: An advanced background-oriented schlieren (BOS) method, named as the speckle beam-oriented schlieren technique, was newly developed to measure the distribution of refraction angles in transparent media. A speckle pattern is generated by passing a coherent laser beam through a holographic diffuser, a pinhole, and a lens, generating a collimated background image that is projected directly onto the image sensors. Since the intensity of the background image is maintained at a high level, this method is, in principle, useful for diagnosing fast and/or low signal-to-noise phenomena, such as high-temperature gasses with radiation emission. Moreover, by splitting the background beam into two imaging paths with different focal lengths, the refraction angles can be measured for a schlieren object with uncertain location, and the depth position of the refraction angles can be resolved. This technique was demonstrated by measuring the refraction angle and the depth position distribution in a sonic jet with different injected locations. Graphic abstract: [Figure not available: see fulltext.]

    DOI: 10.1007/s00348-020-03113-3

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  5. Evaluation of heat removal performance of structure by boil-off gas to reduce propellant evaporation

    FUKUZAKI Toshiya, KINEFUCHI Kiyoshi, UMEMURA Yutaka, OKITA Koichi, SAKAI Hitoshi

        2021.1

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    Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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  6. High Temperature Test of an Electro-thermal Thruster with Multi-wall Tungsten Heater

    KINEFUCHI Kiyoshi, CORAL Giulio, NAKATA Daisuke, SUYA LATU, SAKAI Hitoshi, TSUKIZAKI Ryudo, MARU Yusuke, KOBAYASHI Hiroaki, NISHIYAMA Kazutaka

        2021.1

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    Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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  7. Electric Propulsion System with Dry Ice Propellant

    MAKI Tatsuro, KINEFUCHI Kiyoshi, CHO Shinatora, WATANABE Hiroki

        2021.1

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    Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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  8. Internal neutral density measurement for microwave discharge ion thruster by using two-photon absorption laser-induced fluorescence spectroscopy

    YAMASHITA Yusuke, TSUKIZAKI Ryudo, KINEFUCHI Kiyoshi, NISHIYAMA Kazutaka

        2021.1

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    Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    Space Transportation Symposium FY2020 (January 14-15, 2021. Online Meeting)

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  9. Keeper ignition and discharge characteristics of hollow cathode center-mounted on Hall thruster Reviewed

    Kiyoshi Kinefuchi, Shinatora Cho, Tsutomu Fukatsu, Ryudo Tsukizaki, Ikkoh Funaki, Yuya Hirano, Yosuke Tashiro, Taizo Shiiki

    Journal of Propulsion and Power   Vol. 37 ( 2 ) page: 223 - 230   2021

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    Authorship:Lead author   Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:American Institute of Aeronautics and Astronautics (AIAA)  

    This Paper explores keeper ignition and discharge characteristic of a hollow cathode center mounted on a Hall thruster to discuss the optimum startup sequence. A strong magnetic field along the hollow cathode in the centermount configuration is considered to play an important role in the ignition and discharge; however, the characteristics in this configuration has never been reported. The effect of cathode mass flow rate on the ignition and discharge with 3Akeeperignitionis examined both with and without magnetic field. The influence of cathodeheater power and flow from the anode is also investigated. The lower flow rate causes strong ionization instability, while the higher flow rate seems to induce ion acoustic turbulence. Applied magnetic field improves the ionization instability due to the decrease in electron temperature through plasma production. These results indicate ignition sequences over 1.5 mg=s flow rate with magnetic field appear promising. On the other hand, the confined electrons around the magnetic field line prevent the initiation of the discharge. The ignition and instability are improved with high heat-up power because more thermionic emission is obtained. The anode flow provides slight improvement to the characteristics. The Fourier transform also demonstrates the presence of ionization instability and ion acoustic turbulence. The behavior of peak frequencies of ionization instability can be explained based on the predator-prey theory.

    DOI: 10.2514/1.B37814

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  10. Performance of a miniature hall thruster and an in-house PPU<sup>*</sup>

    Chono M., Yamamoto N., Tsukizaki R., Morishita T., Kubota K., Cho S., Kinefuchi K., Takahashi T.

    Transactions of the Japan Society for Aeronautical and Space Sciences   Vol. 64 ( 3 ) page: 189 - 192   2021

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    Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:Transactions of the Japan Society for Aeronautical and Space Sciences  

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    DOI: 10.2322/tjsass.64.189

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  11. Experimental analysis of thermal behavior in cryogenic propellant tank with different pressurants Reviewed

    Kiyoshi Kinefuchi, Hideto Kawashima, Daizo Sugimori, Yutaka Umemura, Koichi Okita, Hiroaki Kobayashi, Takehiro Himeno

    Cryogenics   Vol. 112   2020.12

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    The thermal behavior of cryogenic propellant tanks is crucial issue in the operation of cryogenic propulsion systems. Herein, ground experiments were conducted in a 600-mm-diameter cryogenic tank filled with liquid nitrogen. As pre-pressurants, gaseous helium and gaseous nitrogen (of the same species as the liquid), were used to investigate its effect in accordance with actual propulsion systems. The tank was sealed after pre-pressurization to observe the self-pressurization. The evaporation rate and heat flow in the tank were estimated based on pressure and temperature measurements. In addition, the axial liquid temperature distribution was obtained through the liquid draining from the tank bottom, and a thermal stratification model was developed. These results demonstrated that the type of pre-pressurant significantly affected the thermal behavior in the tank. A higher evaporation rate and higher liquid internal energy rise rate with a thicker thermal layer were observed in the helium pre-pressurization case. The lower nitrogen partial pressure in the helium case enhanced the vaporization and growth of the thermal layer. Estimation of the power balance in the tank demonstrated that not only the ullage but also the heat mass of the tank provided heat for the evaporation and thermal layer. The evaporation occurs mainly at the contact point between the tank skin and liquid surface. The nitrogen vapor rises in a thin layer along the tank skin because of buoyancy under nitrogen pre-pressurization; however, buoyancy is lower under helium pre-pressurization, and a radial vapor flow is probably produced from the contact point instead, leading to a higher heat flux to the liquid.

    DOI: 10.1016/j.cryogenics.2020.103196

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  12. Application of a microwave cathode to a 200-W Hall thruster with comparison to a hollow cathode Reviewed

    Takato Morishita, Ryudo Tsukizaki, Naoji Yamamoto, Kiyoshi Kinefuchi, Kazutaka Nishiyama

    Acta Astronautica   Vol. 176   page: 413 - 423   2020.11

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    A hollow cathode is an efficient electron source in the self-heating mode utilizing the discharge power. However, in sub-ampere currents, it needs keeper power to maintain the thermionic electron discharge, which could decrease the thrust efficiency. To address this problem, we propose using a microwave cathode, which is based on the flight model of a microwave ion thruster neutralizer cathode, as an alternative to a hollow cathode. First, we redesigned the magnetic field of a microwave cathode discharge chamber and tested it in the diode mode configuration. The electron emission current is doubled compared to the original performance. Next, we coupled the improved microwave cathode with a 200-W class Hall thruster and compared the characteristics and performance with a hollow cathode. We confirmed that the magnetic field polarity affects the ignition characteristics. We measured the thrust by an inverted thrust stand, the ion energy distribution functions by a retarding potential analyzer, and the beam profiles by an ion collector. The thrust and thrust efficiency are equivalent for both types of cathode. The specific impulse is 10% higher in the case of the microwave cathode. Since the potential difference between the microwave cathode and ground rapidly increased at currents above 600 mA, this could be taken to be the trade-off point against the hollow cathode.

    DOI: 10.1016/j.actaastro.2020.06.049

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  13. Cryogenic propellant recirculation for orbital propulsion systems Reviewed

    Kinefuchi Kiyoshi, Kawashima Hideto, Sugimori Daizo, Okita Koichi, Kobayashi Hiroaki

    CRYOGENICS   Vol. 105   2020.1

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    DOI: 10.1016/j.cryogenics.2019.102996

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  14. Characterization of a capillary flow controller for electric propulsion Reviewed

    Kiyoshi Kinefuchi, Shinatora Cho, Ryudo Tsukizaki

    Journal of Propulsion and Power   Vol. 36 ( 4 ) page: 586 - 592   2020

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    A low-cost and reliable propellant flow controller is necessary to offer competitive electric propulsion systems. This paper first reviews xenon flow control methods for Hall thrusters and ion engines. Through a trade study, a simple and low-cost xenon flow controller applying a heated capillary is proposed. The test article is fabricated based on a developed theory and tested with xenon. The estimation based on laminar flow shows slightly different characteristics from those of the experiment, although the Reynolds number indicates a laminar condition. The turbulent flow assumption is in good agreement with the experiment at a high mass flow rate while the flow characteristics are between laminar and turbulent at a low mass flow rate. That implies that the surface roughness and tube curvature induce the laminar–turbulent transition even in the low mass flow rate case. To obtain a wider flow rate range, maintaining laminar flow in the capillary is desirable because of the high sensitivity of the friction coefficient with temperature in the laminar flow. Component improvements achievable through miniaturization and reduction of the inner wall surface roughness and capillary curvature are considered to be necessary to prevent the turbulent transition.

    DOI: 10.2514/1.B37726

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  15. Impacts of Shock and Turbulent Mach Number and Interaction Length on Planar Shock-grid Turbulence Interactions

    福嶋岳, 小川真吾, WEI Jiaxi, 中村友祐, 杵淵紀世志, 佐宗章弘

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2019   2020

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  16. Application of Two-photon Laser-induced Fluorescence Spectroscopy to Microwave Cathode Reviewed

    Ryudo Tsukizaki, Yusuke Yamashita, Kiyoshi Kinefuchi, Kazutaka Nishiyama

    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES   Vol. 63 ( 6 ) page: 281 - 283   2020

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:JAPAN SOC AERONAUT SPACE SCI  

    To investigate the neutral xenon density distribution of electric thrusters such as ion and Hall thrusters, two-photon absorption laser-induced fluorescence (TALIF) spectroscopy was applied to a microwave cathode. First, the background pressure of the vacuum chamber was measured by TALIF. In the present measurements, the ground state was excited by a 224.29 nm laser, and 834.68 nm fluorescence was detected. The first measurement confirmed that the fluorescence intensity linearly increases with respect to the ground state number density. Based on this result, the density of neutral ground-state xenon was measured at the exit of the nozzle of the microwave cathode. The variation in the density with the microwave power was successfully measured at xenon flow rates of 0.029 and 0.098 mg/s. The measured densities varied from 2.3 x 10(19) to 8.4 x 10(19) m(3) with a maximum error of +/- 20% due to the plasma fluorescence.

    DOI: 10.2322/tjsass.63.281

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  17. Cfd modeling of phase change and pressure drop during violent sloshing of cryogenic fluid in a small-scale tank

    O. Kartuzova, M. Kassemi, Y. Umemura, K. Kinefuchi, T. Himeno

    AIAA Propulsion and Energy 2020 Forum     page: 1 - 20   2020

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    This paper presents the development of a two-phase CFD model that is used to study sloshing of a cryogenic fluid (LN2) undergoing phase change generated by lateral acceleration and its effects on the ensuing heat and mass transfer and pressure drop in a small-scale tank. In this model the interface is captured using the volume-of-fluid (VOF) method [1]. Kinetic-theory based Schrage relation [2] is used for calculating phase change mass transfer at the liquid-vapor interface. Computational results are compared to the data provided by a non-isothermal sloshing experiment with phase change conducted by the University of Tokyo and JAXA in 2018 using liquid Nitrogen in a transparent tank. The effect of different thicknesses of a liquid temperature stratification layer created at the interface on the pressure drop during sloshing was studied experimentally. CFD results for interface movement and tank pressure are presented and compared in this paper to experimental data for the cases with the smallest and largest liquid temperature stratification thickness.

    DOI: 10.2514/6.2020-3794

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  18. Cfd modeling of cryogenic chilldown in a complex channel under normal and low gravity conditions

    Justin Pesich, Daniel Hauser, Jason Hartwig, Mohammad Kassemi, Kiyoshi Kinefuchi, Yutaka Umemura, Takehiro Himeno

    AIAA Propulsion and Energy 2020 Forum     page: 1 - 24   2020

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    Future NASA architectures have baselined cryogenic propulsion systems as well as cryogenic fluid management to support lunar missions and ultimately to support future missions to Mars. These missions will require chilling hardware down prior to engine restart as well as chilling lines and tanks prior to transferring and refueling these propulsion elements in orbit. In lieu of expensive tests conducted on-orbit, accurate predictive computational models of these chilldown processes can be used to reduce system and propellant mass as well as mission risk. To gain confidence in these computational models, appropriate anchoring and validation to experimental data in a relevant environment needs to be performed. Recent ground and sub-orbital flight experiments conducted by the Japan Aerospace Exploration Agency (JAXA) investigated chilldown of a complex channel resembling a turbopump bearing cavity at low flow rates. This work presents Computational Fluid Dynamics (CFD) model development of the chilldown experiment employing two-phase flow boiling models available in commercial CFD software STAR-CCM+ using Volume of Fluid (VOF) and the traditional Euler-Euler multiphase flow solvers. Comparisons of the numerical and experimental results under normal and low-gravity conditions are presented. An assessment of solid wall temperatures and phase distribution yielded important insights into multiphase solver choice, dependence on gravity environment, and challenges associated with cryogenic flow boiling prediction and validation.

    DOI: 10.2514/6.2020-3818

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  19. Effects of Applied Electric Field on Supersonic Flow with Laser Energy Deposition

    久保田祥矢, 浅井宏樹, 前田和宏, 市原大輔, 杵淵紀世志, 佐宗章弘

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2019   2020

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  20. Activities in Universities for the Air Turbo Rocket Development

    佐藤哲也, 内海政春, 中田大将, 船崎健一, 武田洋一, MATTHEW Richardson, 真子弘泰, 吹場活佳, 杵淵紀世志

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 64th   2020

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  21. Effect of Pressurant Gas Species on Thermal Behavior of a Cryogenic Tank

    KINEFUCHI Kiyoshi, KAWASHIMA Hideto, SUGIMORI Daizo, KOBAYASHI Hiroaki, UMEMURA Yutaka, OKITA Koichi, HIMENO Takehiro

    The Proceedings of Mechanical Engineering Congress, Japan   Vol. 2020 ( 0 ) page: J19101   2020

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    Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:The Japan Society of Mechanical Engineers  

    DOI: 10.1299/jsmemecj.2020.J19101

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  22. In-flight S-band telemetry attenuation by ionized solid rocket motor plumes at high altitude Reviewed

    Kinefuchi Kiyoshi, Yamaguchi Hiroyuki, Minami Mineko, Okita Koichi, Abe Takashi

    ACTA ASTRONAUTICA   Vol. 165   page: 373-381   2019.12

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    DOI: 10.1016/j.actaastro.2019.09.025

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  23. Neutral ground state particle density measurement of xenon plasma in microwave cathode by two-photon laser-induced fluorescence spectroscopy Reviewed

    Yusuke Yamashita, R. Tsukizaki, Kiyoshi Kinefuchi, D. Koda, Yoshitaka Tani, Kazutaka Nishiyama

    Vacuum   Vol. 168   2019.10

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    Language:Japanese   Publishing type:Research paper (scientific journal)  

    This paper reports the first study to measure xenon neutral ground state particle density of microwave cathode by two-photon laser induced fluorescence spectroscopy (TALIF). Xenon is commonly used as a propellant in electric propulsion like Hall thrusters, ion thrusters, and their cathodes. For electric propulsion, information about neutral particles is important such as the ionization degree and the charged exchange collisions (CEX). The measurement target is XeI 5p S 6p[3/2] , which absorbs at a wavelength of 224.29 nm and emits fluorescence of 834.7 nm. The measurement system was demonstrated for three cases: cold gas, without electron extraction, with electron extraction. From three cases, the measurement system can detect a neutral ground state particle density of 10 m order without and with a plasma. In a cold gas, the neutral ground state particle density is (8.4±0.4)×10 m at 0.098 mg/s. Without electron extraction, the neutral ground density decreases by ionization and excitation With electron extraction, the density varied from 0.6 to 2.3 times compared to without electron extraction depending on anode voltage. 61 19 −3 19 −3 0 2

    DOI: 10.1016/j.vacuum.2019.108846

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  24. Two-photon absorption laser induced fluorescence with various laser intensities for density measurement of ground state neutral xenon Reviewed

    Kiyoshi Kinefuchi, Yoshio Nunome, Shinatora Cho, Ryudo Tsukizaki, Tat Loon Chng

    Acta Astronautica   Vol. 161   page: 382 - 388   2019.8

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    Electrostatic plasma thrusters such as ion engines and Hall thrusters commonly use xenon as a propellant and several measurement techniques for xenon ions and metastable neutrals have been applied to evaluate the characteristics of the thrusters. Although density measurements of ground state neutral xenon can provide crucial information on the ionization characteristics and help explain charge exchange phenomenon, much less research is available due to its technical difficulty. Two-photon absorption laser induced fluorescence is promising because it allows access to ground state xenon atoms by using around 220–260 nm wavelength lasers which have become more readily available lately. In this study, observation of the fluorescence following two-photon excitation from a room temperature (cold) xenon gas cell is conducted with 249 and 252 nm wavelength excitation at xenon pressures of 0.1 and 10 Torr. The fluorescence signals are obtained against a wide range of laser intensities, and the resulting fluorescence response comprises of a few regimes – weak-excitation, saturation, and an intermediate regime. The natural lifetime and quenching rate are evaluated by analyzing the fluorescence decay, and the result is consistent with published literature. Finally, actual application to ground tests of Hall thrusters is discussed based on the experimental results, especially with respect to their fluorescence responses.

    DOI: 10.1016/j.actaastro.2019.03.018

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  25. Development of a new MLI for orbital cryogenic propulsion systems-Thermal performance under one atmosphere to a vacuum

    Takeshi Miyakita, Kazuya Kitamoto, Kiyoshi Kinefuchi, Masanori Saitoh, Tomoyuki Hirai, Hiroyuki Sugita

    IOP Conference Series: Materials Science and Engineering   Vol. 502 ( 1 )   2019.6

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    The efficient storage of cryogenic propellants is among the key technologies for long-duration space exploration missions. For the orbital transfer vehicle, the required thermal insulation performance is more than ten times higher than that of conventional spray-on foam insulation. Conventional multi-layer insulation blankets are used as excellent insulation for spacecraft in the vacuum environment, but are not usable in the atmospheric environment. A new type of insulation-A load-bearing, non-interlayer-contact spacer MLI (LB-NICS MLI)-has been developed. The insulation performance in both air and vacuum environments is measured with a boil-off calorimeter. According to the test results, LB-NICS MLI is about 3 times superior under 1 atmosphere and 13 times superior under a vacuum, compared to foam insulation. &copy; Published under licence by IOP Publishing Ltd.

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  26. LPG Hybrid RCS for 1st Stage Reusable Rocket - Saturated Gas Thruster Test

    Okuda Kazuyoshi, Mizutani Kouichiro, Mimura Takefumi, Okumura Shunsuke, Hattori Daisuke, Kinefuchi Kiyoshi, Saito Yasuhiro, Tahara Hirokazu

        2019.1

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    Space Transportation Symposium FY2018 (January 17-18, 2019. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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  27. Investigation of cryogenic chilldown in a complex channel under low gravity using a sounding rocket Reviewed

    Kiyoshi Kinefuchi, Wataru Sarae, Yutaka Umemura, Takeshi Fujita, Koichi Okita, Hiroaki Kobayashi, Satoshi Nonaka, Takehiro Himeno, Tetsuya Sato

    Journal of Spacecraft and Rockets   Vol. 56 ( 1 ) page: 91 - 103   2019

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    Torealize high-performance cryogenic propulsion systems, the chilldown sequence has to be improved. Because the chilldown is carried out under low gravity, the effect of gravity on the two-phase flow, especially at low flow rate, should be investigated. To understand the physics under low gravity, an experiment was conducted using a sounding rocket. Two identical test sections with different mass flow rates simulated part of a turbopump, each of which has a complex flowpath including slits and a dead end. Using liquid nitrogen, the flight experiment obtained data of temperatures, pressures, void fractions, and video frames of liquid motion. Then, the flight experiment data were compared to the ground data taken under normal gravity, revealing that the slits played an important role in the chilldown process and that the test sections were quickly chilled down under low gravity. The slits of the test sections formed liquid jets, and their behaviors were different from those in the ground experiment. In the flight experiment, the jets easily reached the dead end of the test sections and cooled down the whole walls due to the increase in inertia and wettability; however, such behaviors were hardly observed in the ground experiment. The difference between the ground and flight is significant at lower flow rate.

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  28. 全電化衛星 Reviewed

    杵淵紀世志

    日本航空宇宙学会誌   Vol. 67 ( 6 ) page: 227(J‐STAGE) - 227   2019

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    DOI: 10.14822/kjsass.67.6_227

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  29. デブリ除去衛星への搭載を目指したホールスラスタの開発

    山本直嗣, 長野公勇, 森下貴都, 月崎竜童, 窪田健一, 杵淵紀世志

    宇宙航空研究開発機構特別資料 JAXA-SP-(Web)   ( 18-011 ) page: 503‐509 (WEB ONLY)   2019

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  30. Thermal performance and flow visualization of a planar heat-pipe leading edge Reviewed

    Kiyoshi Kinefuchi, Jun Tamba, Ikuhiko Saito

    Journal of Spacecraft and Rockets   Vol. 56 ( 3 ) page: 771 - 779   2019

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    The sharp leading edges of a hypersonic vehicle are promising for the enhancement of its aerodynamic performance. However, they need a cooling mechanism because they are exposed to extremely high heat flux during reentry or hypersonic cruise, which easily heats up the surface materials beyond their temperature limits. This study explores the concept of a leading edge with planar heat-pipe cooling for which the whole inner surfaces are directly covered with a capillary wick. Its isothermal characteristic and simple configuration can provide robustness; however, due to the uniqueness of the planar type, the thermofluid characteristics should be investigated. Therefore, two kinds of heating tests (the inside visualization test and high-temperature test) were conducted using water as a working fluid. The gravity effects were surveyed, changing the test article attitude. The results showed that the planar concept worked as expected; no large temperature distribution was observed along both the chord and span directions, even in a nose-up attitude or a top heat configuration; the heat transfer performance was improved in nose-down attitude. During the heat-pipe operation, a heat transfer coefficient approximately 100 times higher than that of an evacuated operation was obtained. Visualizing inside the test article revealed that the screen mesh wick was properly covered with condensed water and that typical bubble formations of nucleate boiling were observed around the tip in some cases.

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  31. Capacitive Void Fraction Sensor for Ground Test of LE-5B-3 Reviewed

    坂本勇樹, 小林弘明, 東和弘, 長尾直樹, 杉森大造, 杵淵紀世志, 佐藤哲也

    航空宇宙技術(Web)   Vol. 18 ( 0 ) page: 19‐28(J‐STAGE) - 28   2019

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    &lt;p&gt;Many space vehicles are powered by liquid hydrogen and liquid oxygen. Such fuel are cryogenic fluids, so they are easy to boil and become gas-liquid two phase flow. The LE-5B-3 engine has the capability of the idle mode firing same as the LE-5B-2 engine. Assessment of flow condition at the inlet of fuel turbo pump is important to operate the engine, because the fuel may flow in saturated condition under the idle mode in principle. In a two-phase flow state, void fraction is one of the most important parameters to assess the flow. Although many types of void fraction sensors were proposed, the capacitive technique has advantages to mount on the engine from the viewpoint of size, weight, toughness. In this study, plural circular electrodes capacitive void fraction sensor is developed for LE-5B-3 engines&#039; ground firing test. The sensor was designed based on electric field analysis, and the specification was assessed prior to the ground test. The sensor was used in qualification test, and it was succeeded in achieving stable measurement and it helped to understand the fluid state during the engine operation. The sensor design technique, the assessment results and the ground test results are discussed in this paper.&lt;/p&gt;

    DOI: 10.2322/astj.JSASS-D-18-00017

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  32. Numerical simulation on liquid hydrogen chill-down process of vertical pipeline

    Yutaka Umemura, Takehiro Himeno, Kiyoshi Kinefuchi, Yasuhiro Saito, Jason W. Hartwig, Daniel M. Hauser, Barbara Sakowski, Wesley L. Johnson, Andre C. Leclair, Osamu Fukasawa

    AIAA Propulsion and Energy Forum and Exposition, 2019     2019

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    In order to improve the cryogenic propellant management technologies for a liquid hydrogen rocket with high specific impulse, JAXA, the University of Tokyo, and the NASA Glenn Research Center have jointly organized a multi-agency model validation collaboration project. As part of this project, JAXA's boiling simulation was validated with NASA's experimental data on vertical pipeline chill-down. Simulation results were in good agreement with the experimental data obtained using an improved boiling model to reproduce the spray flow. This activity achieved liquid hydrogen turbo-pump simulation at JAXA for grasping the boiling flow phenomenon from engine cut-off to re-ignition. This joint research resulted in an international cooperative relationship for discussing the cryogenic propellant management technologies necessary to develop next-generation liquid rockets.

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  33. Development and testing of a high-performance 3D printed inconel resistojet

    Giulio Coral, Kiyoshi Kinefuchi, Daisuke Nakata, Kazutaka Nishiyama, Hitoshi Kuninaka

    Proceedings of the International Astronautical Congress, IAC   Vol. 2019-October   2019

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    A 3D printed Inconel resistojet is proposed as an option for short time and high fuel efficiency orbit transfers. The current thruster is presented as a proof-of-concept for high performance high temperature variants. Experiments on N2 propellant have been conducted, and the measured performance parameters are presented. Finally, the extra application of the 3D printed resistojet as part of a hybrid electro-chemical thruster is presented.

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  34. International Space Exploration Scenario (Transportation Architecture) of JAXA

    成田伸一郎, 池永敏憲, 杵淵紀世志, 張科寅, 出原寿紘, 森戸俊樹, 阪口剛史, 宮北健, 梅村悠, 降籏弘城, 池田直史, 和田勝, 関谷優太, 星野健, 古賀勝, 中島潤, 佐藤直樹

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 63rd   2019

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  35. Numerical investigation of nanosecond pulsed plasma actuators for control of shock-wave/boundary-layer separation Reviewed

    Kiyoshi Kinefuchi, Andrey Y. Starikovskiy, Richard B. Miles

    Physics of Fluids   Vol. 30 ( 10 )   2018.10

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    This study numerically explores the flow physics associated with nanosecond pulsed plasma actuators that are designed to control shock-wave induced boundary-layer separation in a Mach 2.8 supersonic flow. By using two dielectric barrier surface discharge actuator configurations, parallel and canted with respect to the flow velocity vector, a previous experiment suggested that the actuator worked in two ways to influence the interaction: boundary layer heating and vorticity production. The heating effect was enhanced with the parallel electrode and made the boundary-layer separation stronger, while the canted electrode produced vorticity and suppressed the boundary-layer separation due to the momentum transfer from the core flow. Because the detailed physical processes are still unclear, in this paper a numerical investigation is undertaken with a large eddy simulation and an energy deposition model for the plasma actuation, in which the dielectric barrier discharge produced plasma is approximated as a high temperature region. The flow characteristics without the plasma actuation correspond to the experimental observation, indicating that the numerical method successfully resolves the shock-wave/boundary-layer interaction. With the plasma actuation, complete agreement between the experiment and calculation has not been obtained in the size of the shock-wave/boundary-layer interaction region. Nevertheless, as with the experiment, the calculation successfully demonstrates definite difference between the parallel and canted electrodes: the parallel electrode causes excess heating and increases the strength of the interaction, while the canted electrode leads to a reduction of the interaction strength, with a corresponding thinning of the boundary layer due to the momentum transfer. The counter flow created by the canted actuator plays an important role in the vortex generation, transferring momentum to the boundary layer and, consequently, mitigating the shock induced boundary layer separation.

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  36. Theoretical and experimental study of the active control of bubble point pressure using a magnetic field and its applications Reviewed

    K. Kinefuchi, H. Kobayashi

    Physics of Fluids   Vol. 30 ( 6 )   2018.6

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    Space propulsion systems use screen mesh devices as filters to block contaminants and as propellant management devices to settle the propellants. The bubble point pressure indicates the basic capillary performance for liquid acquisition of screen meshes. Actively controlling the bubble point pressure can result in flexible and efficient operation of the propulsion systems. High-performance cryogenic propellants, such as liquid hydrogen and oxygen, exhibit magnetic properties. Therefore, a method to actively control the bubble point pressure of cryogenic propellants by applying a magnetic field is proposed in this study. The magnetic pressures affect the pressure balance around the gas-liquid interface separated by the screen mesh, which can thereby control the bubble point pressure. To demonstrate the concept and theoretical basis, a bubble point experiment is conducted using a ferrofluid and solenoid. This experiment proves that the magnetic field actively controls the bubble point pressure and performs both suppression and enhancement of the liquid acquisition performance of the screen mesh. The theory related to magnetic pressures is observed to successfully predict the experimental results. The feasibility of the active control of the bubble point pressure of liquid oxygen is discussed based on the validated theory, and two applications of this technique in cryogenic propulsion systems are depicted.

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  37. Current and Future Missions Using Electric Rocket Propulsion Reviewed

    西山 和孝, 杵淵 紀世志

    プラズマ・核融合学会誌 = Journal of plasma and fusion research   Vol. 94 ( 2 ) page: 60 - 65   2018.2

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  38. Anode shape dependence on thrust performance in a 200 W Class Hall thruster

    長野公勇, 山本直嗣, 窪田健一, 杵淵紀世志

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 62nd   page: ROMBUNNO.1E15   2018

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  39. Analytical Design of a Hybrid Electro‐Chemical Thruster

    CORAL Giulio, KINEFUCHI Kiyoshi, NAKATA Daisuke, SHIMADA Toru, NISHIYAMA Kazutaka, KUNINAKA Hitoshi

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 62nd   page: ROMBUNNO.1N14   2018

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  40. Void Fraction Measurement of Cryogenic Multiphase Flow in Microgravity Reviewed

    小林弘明, 小林弘明, 坂本勇樹, 杵淵紀世志, 佐藤哲也

    日本航空宇宙学会論文集   Vol. 66 ( 6 ) page: 147‐152(J‐STAGE) - 152   2018

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    &lt;p&gt;Reducing the amount of propellant for re-cooling is an important issue for the rocket propulsion system using cryogenic fuel. Immediately after the start of the engine, the liquid fuel boils and becomes two-phase flow. In the state of two-phase flow, the void fraction, which is the gas-liquid ratio, is one of the important value for flow control. For above problem, we are developing void fraction measurement system for the cryogenic fluid. These devices were attached to the S310-43 sounding rocket for the purpose of &quot;measuring two-phase flow behavior and heat transfer characteristics during coasting flight.&quot; These devices withstood the vibration shock test of 40G and succeeded to measure the void fraction of liquid/gas nitrogen two phase flow under vacuumed and microgravity circumstance. This report explains development and experiment results of the void fraction sensor and a capacitance amplifier. &lt;/p&gt;

    DOI: 10.2322/jjsass.66.147

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  41. Thermal design and experimental verification of the 3D-printed resistojet

    Daisuke Nakata, Kiyoshi Kinefuchi

    2018 Joint Propulsion Conference     2018

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    A 3D-printed resistojet, wherein the heater and flow path were united, was produced. The heater consists of multi-layer shells and is difficult to break down. In this paper, the thermal design and the results of the thrust measurement are reported. The measured electrical resistance matched with the predicted one. Both the heater and the thrust reached 70 percent efficiency when paired with a nitrogen propellant at a mass flow rate of 0.2 g/s.

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  42. Facility effect characterization of 6-kw class hall thruster in newly developed high power ep test facility

    Kiyoshi Kinefuchi, Shinatora Cho, Yoshiki Matsunaga, Daisuke Goto, Hiroki Watanabe, Takahiro Yabe, Tadahiko Sano, Tsutomu Fukatsu, Ikkoh Funaki

    2018 Joint Propulsion Conference     2018

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    Japan Aerospace Exploration Agency, JAXA, has been developing an all-electric geostationary satellite named Engineering Test Satellite 9 (ETS-9). It has domestic 6-kW class Hall thruster for both orbit raising and station keeping. Both the 6 kW thruster input power and xenon mass flow rate, 20 mg/s, are the largest ever in Japanese electric propulsion development, therefore, a new test facility with high pumping speed and appropriate cooling system is required. The construction of the new test facility, HTDT (Hall Thruster Development Test) facility, has been completed in Nov 2017. The 3 m in diameter and 8 m long vacuum chamber with 21 cryopumps has successfully achieved 220 kL/s pumping speed with xenon. To remove the 6 kW heat load from the Hall thruster, the facility uses a closed-loop water cooling system and no liquid nitrogen and Freon, hence, the operation cost can be reduced compared with conventional facilities for high power electric propulsion. The requirement to test the 6-kW Hall thruster has been achieved through a 30-hour long qualification test. The facility effect – differences in thruster performance and discharge current oscillation characteristics between in ground facilities and in orbit – is regarded as an important risk in the Hall thruster development. To characterize the facility effect, the 6-kW Hall thruster has been tested in both HTDT facility and JAXA’s another test facility. Preliminary results of the facility effect investigation, comparison of thruster performance and current oscillation against background pressure, are discussed.

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  43. Control of shock-wave/boundary-layer interaction using nanosecond-pulsed plasma actuators Reviewed

    Kiyoshi Kinefuchi, Andrey Y. Starikovskiy, Richard B. Miles

    Journal of Propulsion and Power   Vol. 34 ( 4 ) page: 909 - 919   2018

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    Nanosecond-pulsed surface dielectric barrier discharge plasma actuators are used to control shock-wave/ boundary-layer interactions. Characterization of the separation region without plasma actuation is conducted to understand the interaction based on measurements of pressure distribution, schlieren imaging, and velocimetry by the femtosecond laser electronic excitation tagging technique. The results show a weak separation due to the interaction. Three types of plasma actuators are applied to control the separation. Schlieren images are taken by a high-speed camera to evaluate the magnitude of the interaction. It is shown that the plasma actuators affect the flow in two different ways: heat generation in the boundary layer, and generation of vorticity near the surface.When the first effect is dominant, the shock-wave/boundary-layer interaction becomes stronger and the size of the separation bubble increases. If the vorticity generation prevails, it suppresses the separation due to the momentum transfer from the main flow to the boundary layer. The experimental results of the three actuator's geometries provide the design guidelines for nanosecond-pulse-driven electrodes to control the interaction.An optimal actuation frequency is found through an investigation of the frequency response of the flow.

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  44. 観測ロケットによる低重力環境における極低温沸騰二相流観察実験

    姫野武洋, 幅大地, 更江渉, 杵淵紀世志, 梅村悠, 薮崎大輔, 杉森大造, 小林弘明, 野中聡, 佐藤哲也

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   page: ROMBUNNO.2A2‐2   2018

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  45. 高出力レーザを用いたXeプラズマの分光測定

    山下裕介, 月崎竜童, 杵淵紀世志, 神田大樹, 谷義隆, 西山和孝, 國中均

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 62nd   page: ROMBUNNO.P31   2018

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  46. High Performance and Low Cost Resistojet with 3D Additive Manufacturing

    Kinefuchi Kiyoshi, Matsunaga Yoshiki, Fujii Go, Ikeda Hirohide, Nakata Daisuke, Sakai Hitoshi, Llanillo Rodel

        2017.1

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    Space Transportation Symposium FY2016 (January 19-20, 2017. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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  47. Development Plan of Hall Thruster Sub-system for Next Generation Engineering Test Satellite

    Kinefuchi Kiyoshi, JAXA Hall Thruster R&D Tea

        2017.1

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    Space Transportation Symposium FY2016 (January 19-20, 2017. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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  48. Development of 200 W Class Hall thruster system

    長野公勇, 山本直嗣, 窪田健一, 藤井剛, 杵淵紀世志

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 61st   page: ROMBUNNO.3E02   2017

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  49. Thermal Design and Thrust Measurement of the 3D-Printed Registojet

    中田大将, 杵淵紀世志

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 61st   page: ROMBUNNO.2E03   2017

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  50. 多層断熱材に適用する樹脂スペーサの構造検討

    北本和也, 宮北健, 杵淵紀世志

    構造強度に関する講演会講演集   Vol. 59th   page: 88‐90   2017

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  51. World Trends in All-electric Satellites Reviewed

    杵淵紀世志, 杵淵紀世志

    日本航空宇宙学会誌   Vol. 65 ( 9 ) page: 274‐279(J‐STAGE) - 279   2017

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    &lt;p&gt;全電化静止衛星では推進系としてイオンエンジン,ホールスラスタ等の電気推進のみを搭載し,軌道保持のみならず軌道上昇も電気推進にて行う.これにより従来の化学推進系に比して,長い遷移期間を要するものの大幅に搭載推薬量を削減することができる.2015年,米にて世界初の全電化静止衛星が打上げられ,イオンエンジンによる約半年間の軌道上昇後,静止軌道に到達した.これに追従すべく,各国で全電化静止衛星およびそれに搭載される電気推進の研究開発が活発化している.ホールスラスタはその高推力電力比により3~4カ月程度の短期間遷移を実現でき,さらに低コスト,ロバスト等の特徴も有し,全電化衛星向け電気推進の最有力候補とされる.本稿では全電化静止衛星の世界動向を俯瞰するとともに,全電化衛星への搭載が計画されているホールスラスタの開発動向についても概説する.&lt;/p&gt;

    DOI: 10.14822/kjsass.65.9_274

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  52. Density Measurement of Ground State Xe Atom by Two-photon LIF

    杵淵紀世志, 布目佳央, 張科寅, 月崎竜童, CHNG Tat Loon

    航空原動機・宇宙推進講演会講演論文集(CD-ROM)   Vol. 57th   page: ROMBUNNO.1A17   2017

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  53. Chill-down Process Simulation for Upper Cryogenic Propellant System

    梅村悠, 姫野武洋, 大平勝秀, 河南治, 杵淵紀世志, 小林弘明

    航空原動機・宇宙推進講演会講演論文集(CD-ROM)   Vol. 57th   page: ROMBUNNO.2A01   2017

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  54. Liquid nitrogen chill-down process prediction by direct interface tracking approach

    Yutaka Umemura, Takehiro Himeno, Osamu Kawanami, Wataru Sarae, Kiyoshi Kinefuchi, Hiroaki Kobayashi, Osamu Fukasawa

    53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017     page: .   2017

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    The payload capacity of launch vehicles must be increased in order to extend space exploration and development beyond low-Earth orbit into the solar system. A propellant system using a cryogenic fluid such as liquid oxygen or liquid hydrogen must reduce the amount of unusable propellant due to evaporation and boiling. However, in space exploration and development, where the safety and reliability of missions are critical, predictions of boiling heat transfer using existing technology are not sufficiently reliable for thermal management design, given the lack of pertinent knowledge and relevant research. Therefore, the objective of this research is to understand and accurately predict boiling heat transfer by developing numerical simulation tools for two-phase flows that consider phase change. This paper presents a recent research activity toward the development of chill-down process simulation technology. The cryogenic chilldown experiment conducted in a vertical pipeline and in a complicated channel was verified via simulation to show the effectiveness of the simulation tool under development.

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  55. Thrust Performance of a 200 W Class Hall Thruster Reviewed

    Applied Plasma Science   Vol. 25 ( 2 ) page: 85-88   2017

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    A 200 W class Hall thruster has been developed for the small satellites main propulsion. In order to overcome the degradation owing to the miniaturization, it has a unique magnetic field configuration, similar to "magnetic shielding". The thrust performance was measured using a pendulum thrust stand. The thrust, thrust to power ratio, thrust efficiency is achieved to 13.1 mN,56 mN/kW and 0.36, respectively at discharge voltage of 300 V, xenon mass flow rate of 1.0 mg/s. The obtained thrust performance is comparable to the other 200 W class thruster.

    DOI: 10.34377/aps.25.2_85

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  56. Status of Large Test Facility of Hall Thruster for Next Generation Engineering Test Satellite

    杵淵紀世志, 張科寅, 松永芳樹, 船木一幸, 渡邊裕樹

    航空原動機・宇宙推進講演会講演論文集(CD-ROM)   Vol. 57th   page: ROMBUNNO.1A05   2017

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  57. Study on a new type of multilayer insulation for cryogenic propellant systems of orbital transfer vehicles.

    宮北健, 北本和也, 杵淵紀世志, 斎藤雅規, 平井智行, 杉田寛之

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 61st   page: ROMBUNNO.3F13   2017

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  58. Liquid nitrogen chill-down process prediction by direct interface tracking approach Reviewed

    Umemura Y.

    53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017     2017

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    DOI: 10.2514/6.2017-4761

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  59. Control of shock wave - Boundary layer interaction using nanosecond dielectric barrier discharge plasma actuators

    Kiyoshi Kinefuchi, Audrey Y. Starikovskiy, Richard B. Miles

    52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016     page: .   2016

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    In this study, nanosecond pulse surface dielectric barrier discharge plasma actuators are used to control shock wave boundary layer interaction in Mach 2-8 supersonic air flow. An oblique shock wave is generated by a 14 degree wedge-shaped shock generator, and interacts with the boundary layer. First, characterization of the separation region without plasma actuation is conducted to understand the interaction region based on measurement of pressure distribution, schlieren imaging and velocimetry by the femtosecond laser electronic excitation tagging technique. The results show a weak separation due to the interaction occurs in this configuration. Then, three types of plasma actuator are fabricated and applied to control the separation. Schlieren images are taken with several us camera speeds and the size of interaction region or location of reflected shock wave are measured to evaluate the magnitude of the interaction. It was shown that the nanosecond surface dielectric barrier discharge plasma actuators work in two different ways: the heat generation in the boundary layer, and generation of the vorticity near the surface. In the first case the shock wave - boundary layer interaction becomes stronger and the size of separation bubble increases. In the second case the vorticity production successfully suppresses the boundary layer separation due to momentum transfer from the main flow to the boundary layer. The experimental results of the three actuator configurations provide design guidelines for nanosecond pulse driven electrodes to control the shock wave - boundary layer interaction. An optimal nanosecond pulse frequency was found through an investigation of the frequency response of the flow.

    DOI: 10.2514/6.2016-5070

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  60. Numerical Chill-down Process Simulation for Cryogenic Fluid Propellant System

    UMEMURA Yutaka, HIMENO Takehiro, OHIRA Kastuhide, KAWANAMI Osamu, KINEFUCHI Kiyoshi, KOBAYASHI Hiroaki

    The Proceedings of the Thermal Engineering Conference   Vol. 2016 ( 0 ) page: ROMBUNNO.C113   2016

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    <p>The payload capacity of launch vehicles must be increased in order to allow the exploration and development of space to be extended from low-Earth orbit into the solar system. A propellant system using a cryogenic fluid must reduce the amount of unusable propellant due to evaporation and boiling. However, in space exploration and development, where safety and reliability of missions are critical, predictions of the boiling heat transfer of current technology are not sufficiently reliable for thermal management design due to a lack of knowledge and relevant research. Therefore, the objective of this research is to understand and accurately predict boiling heat transfer by developing numerical simulation tools for two-phase flows that consider phase change. In this paper, recent research activity toward the development of chill-down process simulation technology is presented.</p>

    DOI: 10.1299/jsmeted.2016.C113

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  61. 宇宙輸送機の無効推薬量削減に向けた自由表面流数値解析

    梅村悠, 姫野武洋, 杵淵紀世志, 杉森大造, 薮崎大輔

    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)   Vol. 48th-34th   page: ROMBUNNO.2C14   2016

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  62. 全電化衛星の世界動向

    杵淵紀世志

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 60th   page: ROMBUNNO.1I03   2016

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  63. Effects of Surrounding Temperature and Pressure on Performance of A Gas Thruster

    角銅 洋実, 尾崎 祥梧, 杵淵 紀世志, 池田 博英

    宇宙科学技術連合講演会講演集   Vol. 60th   page: ROMBUNNO.3I18   2016

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  64. 軌道間輸送機に向けた極低温新様式断熱法の検討

    宮北健, 北本和也, 斎藤靖博, 杵淵紀世志, 水谷忠均, 畠中龍太, 斎藤雅規, 平井智行, 杉田寛之

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 60th   page: ROMBUNNO.2K12   2016

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  65. Prediction of in-flight radio frequency attenuation by a rocket plume Reviewed

    Kiyoshi Kinefuchi, Koichi Okita, Ikkoh Funaki, Takashi Abe

    Journal of Spacecraft and Rockets   Vol. 52 ( 2 ) page: 340 - 349   2015.3

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    During rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmission under certain conditions. A computational fluid dynamics and finite difference time-domain method coupling approach was established for predicting interference and radio frequency attenuation levels during an actual rocket flight. The detailed plasma flowfield and radio frequency transmission characteristics were revealed in the calculations. The calculated far-field received levels were compared with the in-flight attenuation data at different look angles (angles between the vehicle axis and the line of sight of the antennas), and the calculated results showed good agreement with the flight data over a wide range of look angles. An adaptation of the model, based on the diffraction theory, proved appropriate both for rough estimation of attenuation and for conducting a preliminary analysis of signal/rocket plume interactions.

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  66. Sounding rocket experiment on chill-down process with liquid nitrogen in a complex channel

    Wataru Sarae, Kiyoshi Kinefuchi, Daisuke Yabusaki, Daizo Sugimori, Takeshi Fujita, Koichi Okita, Yutaka Umemura, Keiichiro Fujimoto, Hideyo Negishi, Hiroaki Kobayashi, Takehiro Himeno, Tetsuya Sato, Satoshi Nonaka

    51st AIAA/SAE/ASEE Joint Propulsion Conference     page: .   2015.1

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    © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In the present experiment, by using the sounding rocket’s sub-orbital ballistic flight, realized the gravitational environment similar to that of liquid-fueled rockets during its coasting flight. In the low-gravity state, the cryogenic test fluid, liquid nitrogen, was introduced into the test sections which had similar shapes to the flow channels in the cryogenic propulsion systems. The boiling of liquid nitrogen inside the test-sections and the transition of flow regimes from gas/liquid two-phase flow to liquid mono-phase flow were successfully visualized. The temperatures, pressures and void fractions in each channel were measured as well. The mechanisms enhancing heat transfer were discussed based on the visualization. In the present case, compared with the corresponding ground test, it was confirmed that the two-phase flow in the complex channel could wet the heat transfer surfaces more easily due to the absence of gravity, and that more uniform chill-down effect could been obtained.

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  67. Numerical modeling of boiling flow in a cryogenic propulsion system

    Yutaka Umemura, Takehiro Himeno, Kiyoshi Kinefuchi, Naoki Tani, Hideyo Negishi, Hiroaki Kobayashi, Katsuhide Ohira, Osamu Fukasawa

    51st AIAA/SAE/ASEE Joint Propulsion Conference     page: .   2015

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    The payload capacity of launch vehicles must be increased in order to allow the exploration and development of space to be extended from low-Earth orbit into the solar system. A propellant system using a cryogenic fluid such as liquid oxygen or liquid hydrogen must reduce the amount of unusable propellant due to evaporation and boiling. However, in the space exploration and development where safety and reliability of missions are critical, predictions of the boiling heat transfer of the present technology are not sufficiently reliable for thermal management design due to a lack of knowledge and relevant research. Therefore, the objective of this research is to understand and accurately predict boiling heat transfer by developing numerical simulation tool for two-phase flows that consider phase change. In this paper, some recent research activities toward the development of chill-down process simulation technology are presented.

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  68. Measuring Two-phase Flow Behavior and Heat Transfer Characteristics during Coasting Flight, Development of Experimental Equipment for S-310-43 Sounding Rocket Reviewed

    小林弘明, 杵淵紀世志, 更江渉, 梅村悠, 藤本圭一郎, 薮崎大輔, 杉森大造, 姫野武洋, 佐藤哲也, 北古賀智史, 角悠輝, 坂本勇樹, 野中聡, 藤田猛

    日本航空宇宙学会論文集   Vol. 63 ( 5 ) page: 188-196 (J-STAGE) - 196   2015

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    The Japan Aerospace Exploration Agency launched the S-310-43 sounding rocket from the Uchinoura Space Center on Aug.04, 2014 for the purpose of investigating such behavior as boiling and flow of cryogenic liquid rocket propellant in an environment simulating coasting flight on orbit by using the sounding rocket's sub-orbital ballistic flight. In the low-gravity state, the cryogenic fluid (liquid nitrogen) was introduced into the test sections of similar shapes to the flow channels in the cryogenic propulsion systems. The boiling of liquid nitrogen inside the test-sections and the transition of flow regimes from gas/liquid two-phase flow to liquid mono-phase flow were visualized. The temperatures, pressures and void fractions of each channels were measured as well. Development of the experimental equipment for S-310-43 sounding rocket is described in this paper.

    DOI: 10.2322/jjsass.63.188

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  69. 観測ロケット実験による極低温沸騰二相流観察実験

    更江渉, 杵淵紀世志, 小林弘明, 梅村悠, 藤本圭一郎, 薮崎大輔, 杉森大造, 姫野武洋, 野中聡, 藤田猛, 佐藤哲也

    航空原動機・宇宙推進講演会講演論文集(CD-ROM)   Vol. 55th   page: ROMBUNNO.1B14   2015

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  70. 観測ロケット S310 -43 号機試験における極低温二相流のボイド率計測

    北古賀 智史, 角 悠輝, 坂本 勇樹, 佐藤 哲也, 小林 弘明, 杵淵 紀世志, Sumi Yuki, Sato Tetsuya, Kobayashi Hiroaki, Kinebuchi Kiyoshi

    平成26年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium FY2014     2015

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    平成26年度宇宙輸送シンポジウム(2015年1月15日-16日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県資料番号: SA6000036020レポート番号: STCP-2014-020

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  71. Sounding rocket experiment on chill-down process with liquid nitrogen in a complex channel

    Sarae W.

    51st AIAA/SAE/ASEE Joint Propulsion Conference     2015

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    DOI: 10.2514/6.2015-4213

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  72. Ignition characteristics of 15 kW Arcjet with swirl injection

    Yoshida Koki, Nakata Daisuke, Kinefuchi Kiyoshi, Hosoda Satoshi, Koda Daiki, Tsukizaki Ryudo, Kuninaka Hitoshi

        2015

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    Space Transportation Symposium FY2014 (January 15-16, 2015. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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  73. High Power Hydrogen Electric Propulsion for the Space Exploration

    KINEFUCHI Kiyoshi, OKITA Koichi, KUNINAKA Hitoshi, NAKATA Daisuke, TAHARA Hirokazu

      Vol. 2014   page: ELECTRIC PROPULSION 2,K. KINEFUCHI   2014

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  74. 宇宙輸送系に関わる自由表面流の数値解析

    梅村悠, 姫野武洋, 根岸秀世, 杵淵紀世志, 大平勝秀, 井上智博, 渡辺紀徳

    数値流体力学シンポジウム講演論文集(CD-ROM)   Vol. 28th   page: ROMBUNNO.F02-4   2014

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  75. Technical Challenges for Advanced Arcjets Reviewed

    Daisuke NAKATA, Kiyoshi KINEFUCHI, Satoshi HOSODA, Masahiro KINOSHITA, Hitoshi KUNINAKA

    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN   Vol. 12 ( ists29 ) page: To_1_1 - To_1_5   2014

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    DOI: 10.2322/tastj.12.to_1_1

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  76. Space Exploration Strategy with High Power Plasma Propulsions

    KINEFUCHI Kiyoshi, SAITO Yasuhiro, NAGAO Naoki, OKITA Koichi, FUNAKI Ikkoh, KUNINAKA Hitoshi

    プラズマプロセシング研究会プロシーディングス(CD-ROM)   Vol. 32nd   page: ROMBUNNO.S2-2   2014

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  77. Preliminary study of high power hydrogen electric propulsion for the space exploration

    Kiyoshi Kinefuchi, Koichi Okita, Hitoshi Kuninaka, Daisuke Nakata, Tomoya Suzuki, Hirokazu Tahara

    50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference 2014   Vol. 50th Vol.2   page: 1243 - 1252   2014

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    High power electric propulsion system is strongly required for future orbital space transportation. MPD (Magneto-Plasma-Dynamic) thrusters and DC (Direct Current) arcjets with hydrogen as a propellant are promising candidates for the missions because of their high performance and adaptability to high power operation. However, to use hydrogen for long term orbital missions, its storage in orbit is crucial issue to be considered. Firstly, we proposed a hydrogen storage and feed system for electric thrusters by applying our technologies derived from the liquid hydrogen launch vehicles. Secondly, we present R&D activities of hydrogen MPD thruster and DC arcjet, especially focusing on the improvement of their performance and durability. Then, development strategy of hydrogen electric thrusters is also discussed. Finally, advantages of hydrogen electric thruster were shown compared with conventional xenon thrusters through mission analyses of lunar orbit insertion and GTO-GEO transportation.

    DOI: 10.2514/6.2014-3507

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  78. 大型電気推進実現に向けた試験設備構想と世界状況

    杵淵 紀世志, 長尾 直樹, 齊藤 靖博, 沖田 耕一, 國中 均, Kinebuchi Kiyoshi, Nagao Naoki, Saito Yasuhiro, Okita Koichi, Kuninaka Hitoshi

    平成25年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium FY2013     2014

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    平成25年度宇宙輸送シンポジウム(2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県資料番号: SA6000016097レポート番号: STEP-2013-024

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  79. 「きく6号」イオンエンジン用の試験装置

    梶原 堅一, 杵淵 紀世志, Kajiwara Kenichi, Kinebuchi Kiyoshi

    平成25年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium FY2013     2014

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    平成25年度宇宙輸送シンポジウム(2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県資料番号: SA6000016099レポート番号: STEP-2013-026

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  80. DCアークジェットの軽量化・長寿命化

    中田 大将, 杵淵 紀世志, 木下 昌洋, 國中 均, Nakata Daisuke, Kinebuchi Kiyoshi, Kinoshita Masahiro, Kuninaka Hitoshi

    平成25年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium FY2013     2014

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    平成25年度宇宙輸送シンポジウム(2014年1月16日-17日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県資料番号: SA6000016113レポート番号: STEP-2013-040

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  81. Prediction of in-flight radio frequency attenuation by a rocket plume by applying CFD/FDTD coupling

    Kinefuchi K.

    49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference     2013.9

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  82. A Trial Manufacure and A Fundamental Study on Recirculation Pump for Rocket Engine

    WATANABE Mitsuo, HASHIMOTO Tomoyuki, KINEFUCHI Kiyoshi, SUGITA Eiichiro

      Vol. 41 ( 4 ) page: 193 - 200   2013.4

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  83. Prediction of In-Flight Radio Frequency Attenuation by Rocket Plume Applying Diffraction Theories Reviewed

    Kiyoshi Kinefuchi, Ikkoh Funaki, Takashi Abe

    JOURNAL OF SPACECRAFT AND ROCKETS   Vol. 50 ( 1 ) page: 150-158   2013.1

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    DOI: 10.2514/1.A32223

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  84. 水素系電気推進による軌道間輸送の初期検討

    杵淵 紀世志, 沖田 耕一, 國中 均, Kinebuchi Kiyoshi, Okita Koichi, Kuninaka Hitoshi

    平成24年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium: FY2012     2013.1

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    平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856104レポート番号: STEP-2012-021

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  85. 15kW級DCアークジェットのカソード性能比較

    木下 昌洋, 杵淵 紀世志, 中田 大将, 細田 聡史, 國中 均, Kinoshita Masahiro, Kinefuchi Kiyoshi, Nakata Daisuke, Hosoda Satoshi, Kuninaka Hitoshi

    平成24年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium: FY2012     2013.1

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    平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF資料番号: AA0061856113レポート番号: STEP-2012-030

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  86. ISASあきる野実験施設におけるH2Aロケット高度化ベントリテンション開発試験

    杵淵 紀世志, 沖田 耕一, 更江 渉, 藤田 猛, 小林 弘明, 八木下 剛, 小林 清和, 徳永 好志, 堀 恵一, 佐藤 哲也, 西村 真二, 北山 治, Kinefuchi Kiyoshi, Okita Koichi, Sarae Wataru, Fujita Takeshi, Kobayashi Hiroaki, Yagishita Tsuyoshi, Kobayashi Kiyokazu, Tokunaga Yoshiyuki, Hori Keiichi, Sato Tetsuya, Nishimura Shinji, Kitayama Osamu

    平成24年度宇宙輸送シンポジウム: 講演集録 = Proceedings of Space Transportation Symposium: FY2012     2013.1

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    平成24年度宇宙輸送シンポジウム (2013年1月17日-1月18日. 宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)), 相模原市, 神奈川県形態: カラー図版あり形態: PDF著者人数: 12名資料番号: AA0061856048レポート番号: STCP-2012-048

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  87. Prediction of in-flight radio frequency attenuation by a rocket plume by applying CFD/FDTD coupling

    Kiyoshi Kinefuchi, Koichi Okita, Ikkoh Funaki, Takashi Abe

    49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference   Vol. 1 PartF   2013

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    During rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency (RF) transmission under certain conditions. To clarify the physical process involved and to establish the estimation methodology, a plume–RF interference experiment during a sea-level static firing test of a full-scale solid rocket motor was conducted. The result of the ground experiment was adequately matched by a computational fluid dynamics (CFD) model of the plume flow field coupled to a finite-difference time-domain (FDTD) model of RF transmission. The CFD/FDTD coupling method was then refined for predicting interference and RF attenuation levels during an actual rocket flight. The calculated far-field received levels were compared with the in-flight attenuation data at different look angles (angles between the vehicle axis and the line-of-sight of the antennas). The calculated results showed good agreement with the flight data over a wide range of look angles. An adaptation of the model, based on the diffraction theory, proved appropriate both for rough estimation of attenuation and for conducting a preliminary analysis of signal/rocket plume interactions.

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  88. 数値シミュレーションを活用した宇宙空間におけるガスプルームコンタミ評価

    谷洋海, 磯部直樹, 中川貴雄, 杵淵紀世志, 谷直樹, 根岸秀世

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 57th   page: ROMBUNNO.1F14   2013

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  89. Ground Experiment for Development of Liquid Propellant Acquisition Devices under Microgravity Reviewed

    杵淵紀世志, 加納康仁, 齊藤靖博, 沖田耕一, 姫野武洋

    航空宇宙技術(Web)   Vol. 12 ( 0 ) page: 73-77 (J-STAGE) - 77   2013

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    For future space transportation system development, efficient liquid propellant acquisition technologies under microgravity could be required to realize long-term missions in orbit. Microgravity environment is generally established through drop tower, parabolic flight by airplane or orbital/suborbital experiment. These methods are large-scale or not flexible so that compact and simple method is needed for the efficient development. To respond to such request, we propose a static experiment on the ground to realize the similar static free surface under microgravity. The ground experimental result was compared with the results of two types of two-phase flow simulation codes for the verification of the methodology and discuss the characteristics of these numerical codes. We also simulated the free surface under the actual flight condition by using the same simulation method and result showed the validity of the experimental method.

    DOI: 10.2322/astj.12.73

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  90. 基幹ロケット上段推進系の発展と国際有人探査への貢献―極低温軌道間輸送機の研究開発―

    杵淵紀世志, 齊藤靖博, 西平慎太郎, 更江渉, 杉森大造, 沖田耕一, 谷直樹, 梅村悠, 小林弘明, 姫野武洋, 石川佳太郎, 青山太一

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 57th   page: ROMBUNNO.1J05   2013

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  91. 固体ロケット飛翔中のプルーム・電波干渉の解析予測

    杵淵紀世志, 船木一幸, 沖田耕一, 安部隆士

    航空原動機・宇宙推進講演会講演集(CD-ROM)   Vol. 53rd   page: ROMBUNNO.JSASS-2013-0001   2013

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  92. ロケットエンジン用リサーキュレーションポンプの試作と基礎的な試験について Reviewed

    渡邉光男, 橋本知之, 杵淵紀世志, 杉田栄一郎

    ターボ機械   Vol. 41 ( 4 ) page: 193 - 200   2013

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  93. 液体窒素のスロッシングに伴うタンク内圧力変化の研究

    杉森大造, 藤田真澄, 飯田久訓, 小川洋平, 杵淵紀世志, 薮崎大輔, 姫野武洋, 梅村悠, 石川勝利

    日本機械学会年次大会講演論文集(CD-ROM)   Vol. 2013   page: ROMBUNNO.G051014   2013

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  94. G051014 Investigation of the tank pressure change caused by sloshing of liquid nitrogen

    SUGIMORI Daizo, FUJITA Masumi, IIDA Hisanori, OGAWA Yohhei, KINEFUCHI Kiyoshi, YABUSAKI Daisuke, HIMENO Takehiro, UMEMURA Yutaka, ISHIKAWA Katsutoshi

    The Proceedings of Mechanical Engineering Congress, Japan   Vol. 2013 ( 0 ) page: _G051014 - 1-_G051014-5   2013

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    For the prediction of heat transfer coupled with sloshing phenomena in the propellant tanks of launch vehicle, the pressure drop induced by heat transfer and the dynamic motion of cryogenic liquid in sub-scale vessels were experimentally observed. The correlation between the pressure drop and liquid motion was confirmed in the experiment. Results of the test suggests that pressure is sensitive when gaseous species are same as liquid. In addition, even if a gaseous temperature is different (about 3 OK), the amount of pressure drop does not change.

    DOI: 10.1299/jsmemecj.2013._G051014-1

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  95. Computational fluid dynamics and frequency-dependent finite-difference time-domain method coupling for the interaction between microwaves and plasma in rocket plumes Reviewed

    K. Kinefuchi, I. Funaki, T. Shimada, T. Abe

    Physics of Plasmas   Vol. 19 ( 10 ) page: 102112-102112-7   2012.10

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    Under certain conditions during rocket flights, ionized exhaust plumes from solid rocket motors may interfere with radio frequency transmissions. To understand the relevant physical processes involved in this phenomenon and establish a prediction process for in-flight attenuation levels, we attempted to measure microwave attenuation caused by rocket exhaust plumes in a sea-level static firing test for a full-scale solid propellant rocket motor. The microwave attenuation level was calculated by a coupling simulation of the inviscid-frozen-flow computational fluid dynamics of an exhaust plume and detailed analysis of microwave transmissions by applying a frequency-dependent finite-difference time-domain method with the Drude dispersion model. The calculated microwave attenuation level agreed well with the experimental results, except in the case of interference downstream the Mach disk in the exhaust plume. It was concluded that the coupling estimation method based on the physics of the frozen plasma flow with Drude dispersion would be suitable for actual flight conditions, although the mixing and afterburning in the plume should be considered depending on the flow condition. © 2012 American Institute of Physics.

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  96. Experimental investigation on thermochemical phenomena in SiFRP Reviewed

    Kenichi Hirai, Yoshiki Matsuura, Kiyoshi Kinefuchi, Toru Kamita

    Advanced Composite Materials   Vol. 21 ( 5-6 ) page: 459 - 475   2012.10

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    This study focuses on understanding and modeling the physical phenomena that occur in degraded zones of silica-phenolic (SiFRP) materials under exposure to high-temperature gasses when applied to a liquid rocket engine (LRE) combustor. Although understanding and modeling these phenomena is considered essential in designing an LRE combustor, few studies on these fields can be found in the available literature. Basically, it is well known that when ablators are heated, a pyrolysis reaction proceeds in them, forming three distinct zones: a charred, a decomposed, and a virgin zone. The obtainable information for the thermal response of SiFRP in ground-firing tests is classified in two categories. The first category involves the equilibrium state characteristics after a long time has elapsed following burnout. This refers to the degraded thickness distribution, which reflects 3D information (the combustor's inner surface x the thickness direction) regarding the heat load distribution over the entire combustor's inner surface, owing to the highly insulating nature of SiFRP. The second category involves the transient characteristics of the propagation of the degraded zones in SiFRP, which can be detected using an ultrasonic testing (UT) method. In this paper, the progress of in-depth phenomena of SiFRP and their physical variations were intentionally studied. Our aim was to clarify and specify the quantitative threshold values of the interface points that characterize each degraded zone and the UT reflection point, and then express these values in terms of physical quantities that could appear in a numerical analysis. © 2012 Japan Society for Composite Materials, Korean Society for Composite Materials and Taylor & Francis.

    DOI: 10.1080/09243046.2012.743715

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  97. Experimental Investigations on the SiFRP Surface Thermal Phenomena

    HIRAI Kenichi, KINEFUCHI Kiyoshi, KAMITA Toru

      Vol. 38 ( 5 ) page: 200 - 206   2012.9

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  98. 15kW級DCアークジェットの作動特性

    木下昌洋, 杵淵紀世志, 中田大将, 細田聡史, 國中均

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 56th   page: ROMBUNNO.1J16   2012

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  99. Experimental Investigations on the SiFRP Surface Thermal Phenomena Reviewed

    平井研一, 杵淵紀世志, 紙田徹

    日本複合材料学会誌   Vol. 38 ( 5 ) page: 200 - 206   2012

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    This study focuses on the understanding and modeling of the physical phenomena occurring in the heated surface of Silica Phenolic (hereafter referred as SiFRP) under exposure to high temperature gases when applied to liquid rocket engine (LRE) combustor. Although the understanding and modeling of the phenomena is supposed to be essential in designing LRE combustor, the works done in this fields are seldom seen in open literatures. When we look at the inner surface of SiFRP combustor of LRE after ground firing tests, the signs of white streak, melting and recession are sometimes observed. In this research, the experimental characterization activities on the SiFRP surface phenomena are conducted, based on the various optical measurements (digital camera, SEM (scanning electron microscope), XRF (X-ray fluorescence), spectral reflectivity (wavelength: 400-700 nm)) for the heated samples, namely, the ones from arcjet heating tests under Air or N<sub>2</sub> stream conditions or the ones from SiFRP combustor after LRE.

    DOI: 10.6089/jscm.38.200

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  100. 推進薬管理技術の高度化へ向けた数値流体解析手法の研究

    姫野武洋, 杵淵紀世志, 沖田耕一, 谷直樹, 野中聡

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 56th   page: ROMBUNNO.1H14   2012

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  101. Experimental Investigations on the Thermochemical Phenomena in the SiFRP Reviewed

    平井研一, 松浦芳樹, 杵淵紀世志, 紙田徹

    日本複合材料学会誌   Vol. 38 ( 6 ) page: 228 - 235   2012

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    This study focuses on the understanding and modeling of the physical phenomena occurring in the degraded zones of Silica Phenolic (hereafter referred as SiFRP) under exposure to high temperature gases when applied to liquid rocket engine (LRE) combustor. Although the understanding and modeling of the phenomena is supposed to be essential in designing LRE combustor, a few works done in this fields appear in the open literatures. Basically, it is well known that when heated, the pyrolysis reaction proceeds in the SiFRP, forming 3 distinct zones of charred, decomposed and virgin zone, respectively. The obtainable information for the thermal response of SiFRP at ground firing tests is classified in 2 categories. The first is the equilibrium state characteristics after long time elapsed from the burnout, namely, the degraded thickness distribution, which reflects 3-D information (combustor inner surface×thickness-direction) of heat load distribution over the entire combustor inner surface thanks to the highly insulating nature of SiFRP. The second is the transient characteristics with regard to the degraded zones propagation in the SiFRP, which can be detected by application of ultrasonic testing (UT) method. In this paper, the progress of in-depth phenomena of SiFRP and the physical variation were intentionally studied. We strive to clarify and specify the quantitative threshold values of the interface points which characterizes each degraded zone and UT reflection point and eventually express the threshold values in terms of physical quantities that could appear in numerical analysis.

    DOI: 10.6089/jscm.38.228

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  102. 極低温軌道間輸送機(CPS)の実現に向けて―基幹ロケット上段推進系の発展構想―

    杵淵紀世志, 齊藤靖博, 山西伸宏, 更江渉, 沖田耕一, 谷直樹, 小林弘明, 姫野武洋, 青山太一, 北山治

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 56th   page: ROMBUNNO.1A04   2012

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  103. 軌道間輸送システムの構想と研究開発計画

    沖田耕一, 杵淵紀世志, 齊藤靖博, 長尾直樹, 山西伸宏, 國中均

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 56th   page: ROMBUNNO.3S12   2012

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  104. 軌道間輸送システムのミッション&技術ロードマップ

    齊藤靖博, 杵淵紀世志, 沖田耕一, 國中均

    宇宙科学技術連合講演会講演集(CD-ROM)   Vol. 56th   page: ROMBUNNO.1J09   2012

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  105. 基幹ロケット推進系開発計画

    杵淵紀世志, 東伸幸, 更江渉, 沖田耕一, 谷直樹, 小林弘明, 八木下剛, 北山治, 姫野武洋

    宇宙科学技術連合講演会講演集(CD−ROM)   Vol. 55th   page: ROMBUNNO.1S13   2011

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  106. Development of composite structures for launch vehicles Reviewed

    Kiyoshi Kinefuchi, Toshiyuki Uzawa, Isao Tate, Miki Nishimoto, Hirotaka Igawa, Toru Kamita

    Nihon Kikai Gakkai Ronbunshu, A Hen/Transactions of the Japan Society of Mechanical Engineers, Part A   Vol. 77 ( 773 ) page: 81 - 89   2011

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    Ultimate weight reduction and low manufacturing cost are strongly required in the development of next-generation launch vehicles. It is one of effective methods for these purposes to change structural materials from conventional aluminum alloys to composites, such as CFRP (Carbon Fiber Reinforced Plastics). Under these situations, we developed a launch vehicle structure by using CFRP/honeycomb sandwich panels in order to achieve both weight and cost reduction. FEM (Finite Element Method) simulation was utilized to effectively carry out the structural design. Metallic structures are applied for the flanges (joining interfaces between the other structures) in general; however, we used CFRP structures also for the interface flanges for the weight reduction and manufacturing simplicity. The designed CFRP structure was fabricated and a static load test was conducted considering flight loads during actual launch. The strain distribution was obtained by FBG (Fiber Bragg Grating) sensors and three-dimensional displacement was measured by laser tracker as well as normal strain gauges in the limit load test. These measured data were compared with the FEM simulation results and they show good agreement. We will apply these manufacturing, testing, simulation and sensing techniques for the future launch vehicle developments. © 2011 The Japan Society of Mechanical Engineers.

    DOI: 10.1299/kikaia.77.81

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  107. Heat exchange and pressure drop enhanced by sloshing

    Takehiro Himeno, Daizo Sugimori, Katsutoshi Ishikawa, Yutaka Umemura, Seiji Uzawa, Chihiro Inoue, Toshinori Watanabe, Satoshi Nonaka, Yoshihiro Naruo, Yoshifumi Inatani, Kiyoski Kinefuchi, Ryoma Yamashiro, Toshiki Morito, Koichi Okita

    47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011   Vol. 47th Vol.3   page: 1793-1832   2011

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    For the prediction of heat transfer coupled with sloshing phenomena in the propellant tanks of reusable launch vehicle, the pressure drop induced by heat transfer and the dynamic motion of liquid in sub-scale vessels were experimentally observed and numerically investigated. The correlation between the pressure drop and liquid motion was confirmed in the experiment. The mechanisms enhancing heat transfer were discussed based on the computation. It was suggested that splash and wavy surface induced by violent motion of liquid cause the pressure drop in the closed vessel. In addition, as the preliminary investigation, non-isothermal sloshing of liquid nitrogen and liquid hydrogen were successfully visualized and pressure drop depending on the gaseous species was discussed. © 2011 by Takehiro Himeno.

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  108. Frequency-dependent FDTD simulation of the interaction of microwaves with rocket-plume Reviewed

    Kiyoshi Kinefuchi, Ikkoh Funaki, Takashi Abe

    IEEE Transactions on Antennas and Propagation   Vol. 58 ( 10 ) page: 3282 - 3288   2010.10

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    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) 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. © 2010 IEEE. 2

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  109. Development of L-605 Nozzle Extension

    MOTOGI Takayuki, MOTEKI Manabu, YOSHIDA Taiga, TATE Isao, KINEFUCHI Kiyoshi, KAMITA Toru, TORII Yoshihiro

    航空原動機・宇宙推進講演会講演集(CD−ROM)   Vol. 50th   page: ROMBUNNO.AJCPP2010-116   2010

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  110. Thermal Issues in the Design of Ablative Combustion Chamber of Liquid Rocket Engines

    HIRAI Kenichi, MATSUURA Yoshiki, KINEFUCHI Kiyoshi, KAMITA Toru

    航空原動機・宇宙推進講演会講演集(CD−ROM)   Vol. 50th   page: ROMBUNNO.AJCPP2010-149   2010

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  111. Numerical Prediction of Microwave-Rocket Plume Interaction

    KINEFUCHI Kiyoshi, FUNAKI Ikkoh, SHIMADA Toru, ABE Takashi

    航空原動機・宇宙推進講演会講演集(CD−ROM)   Vol. 50th   page: ROMBUNNO.AJCPP2010-144   2010

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  112. Experimental investigation on microwave interference in full-scale solid rocket exhaust Reviewed

    Kiyoshi Kinefuchi, Ikkoh Funaki, Toru Shimada, Takashi Abe

    Journal of Spacecraft and Rockets   Vol. 47 ( 4 ) page: 627 - 633   2010

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    Under certain conditions during rocket flight, ionized exhaust plumes from solid rocket motors may interfere with RF transmission. To understand the relevant physical processes involved in this phenomenon, measurement of microwave attenuation and phase delay caused by rocket exhaust plumes was attempted in a sea-level static firing test for a full-scale solid propellant rocket motor. The measured data were analyzed by comparing them with simulation results for an exhaust plume flowfield. The results revealed that the change in the shock structure in the plume affects the microwave attenuation level, since it significantly affects the plasma density at the measuring location. When the plasma density in the plume is low, the microwaves can penetrate the plume. The plume plasma properties were successfully estimated for that situation in which the numerically calculated attenuation level agreed well with the experimental results. On the other hand, high-density plasma in the plume does not allow penetration. Therefore, microwaves bypass around the plume and the diffraction effect becomes dominant. Copyright © 2010 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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  113. Experimental Investigation and Numerical Approach of Liquid propellant residual in Rocket Tank Bottom

    FUJIMURA Masanobu, YAMADA Keisuke, KINEFUCHI Kiyoshi, KAMITA Toru

    航空原動機・宇宙推進講演会講演集(CD−ROM)   Vol. 50th   page: ROMBUNNO.AJCPP2010-118   2010

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  114. Experimental and numerical simulation study of liquid-propellant draining from rocket tanks Reviewed

    Kiyoshi Kinefuchi, Toru Kamita, Hideyo Negishi, Keisuke Yamada, Masanobu Fujimura

    Journal of Spacecraft and Rockets   Vol. 47 ( 5 ) page: 860 - 863   2010

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    The draining process for the circular-plate-type device, using both experiments and numerical simulations was investigated. The effect of turbulence was evaluated by conducting both turbulent and laminar flow calculations. The Navier-Stokes equation was spatially discretized by the finite-volume method, and the convective flux was evaluated by a second-order upwind scheme. The air and water gas-liquid two-phase flow was modeled by the volume-of-fluid (VOF) method, and the surface tension of water was assumed to have a 75° contact angle to the wall. The cell distribution of the normal 2-D grid and radial cross section of the 3-D grid were similar and the minimum grid size was 0.125 mm around the baffle edge and the maximum was approximately 5 mm at the tank equator for both grids. A numerical accuracy study based on Richardson statistics indicated that the fractional errors of the liquid remaining were roughly 19 and 1.2% for the normal and finer-grid systems, respectively.

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  115. Investigation of microwave attenuation by solid rocket exhausts

    Kiyoshi Kinefuchi, Ikkoh Funaki, Hiroyuki Ogawa, Teruo Kato, Sumitaka Tachikawa, Toru Shimada, Takashi Abe

    47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition     2009.1

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  116. Investigation of microwave attenuation by solid rocket exhausts

    Kiyoshi Kinefuchi, Ikkoh Funaki, Hiroyuki Ogawa, Teruo Kato, Sumitaka Tachikawa, Toru Shimada, Takashi Abe

    47th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition   Vol. 49th   page: B11   2009

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    In rocket flights, ionized exhaust plumes from solid rocket motors may interfere with RF transmission under some conditions. In order to clarify the important physical process involved, microwave attenuation and phase delay due to rocket exhaust plumes were measured during sea-level static firing tests conducted on two types of full-scale solid propellant rocket motors. The measured data were analyzed by comparing them with numerical results such as flowfield simulations of exhaust plumes and by employing a detailed analysis of microwave transmission by using a frequency-dependent finite-difference time-domain (FD2TD) method. 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. For the actual launch vehicle flight, the transmission process is dominated by the diffraction effect so that we applied a two-dimensional diffraction theory to analyze the communication between a vehicle and a ground station. The attenuation levels estimated using diffraction theory agree with the data recorded in-flight. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

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  117. Study for Validation of a Liquid Rocket Engine Ablative Chamber

    杵淵紀世志, 紙田徹, 平井研一

    航空原動機・宇宙推進講演会講演論文集(CD-ROM)   Vol. 49th   page: A35   2009

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  118. Laser absorption velocimetry of plasma flow in two-dimensional magnetoplasmadynamic arcjet Reviewed

    Kiyoshi Kinefuchi, Ikkoh Funaki, Kyoichiro Toki

    Journal of Propulsion and Power   Vol. 22 ( 5 ) page: 1085 - 1090   2006

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    Experimental velocimetry in the discharge chamber of a two-dimensional magnetoplasmadynamic (MPD) arcjet, fabricated for experimental internal flow measurement, was conducted to investigate the acceleration process for hydrogen propellant. In the experiment, we evaluated the neutral atom velocity and the temperature from the laser absorption spectroscopy using a tunable diode laser. The results using two types of anode, a flared-tvpe anode and a converging-diverging (C-D)-type anode, were compared for the case with a discharge current of 13 kA and a mass-flow rate of 0.65 g/s. It was found that a large velocity slip between the ions and the neutrals prevented the acceleration of the neutral particles. This velocity slip is expected to reduce thrust performance because the flow with ion-neutral slip requires additional electric power compared to the flow without velocity slip. The velocity slip was reduced in the case of the C-D anode compared to the flared anode because of strong ion-neutral momentum coupling in the throat region of the C-D anode. Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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  119. LE‐5Bエンジン起動特性の検討

    杵淵紀世志, 内海政春, 長谷川恵一, 沖田耕一, 真子弘泰, 恩河忠興

    日本航空宇宙学会北部支部講演会講演論文集   Vol. 2006   page: 151-154   2006

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  120. Laser absorption velocimetry of plasma flow in two-dimensional magnetoplasmadynamic arcjet

    Kinefuchi Kiyoshi, Funaki Ikkoh, Toki Kyoichiro

    JOURNAL OF PROPULSION AND POWER   Vol. 22 ( 5 ) page: 1085 - 1090   2006

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  121. Measurement of Velocity and Power Balance in a Two-Dimensional MPD Arcjet Reviewed

    杵淵紀世志, 船木一幸, 都木恭一郎, 清水幸夫

    日本航空宇宙学会論文集   Vol. 53 ( 616 ) page: 215 - 223   2005

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    Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

    Velocity and temperature measurements were conducted for a two-dimensional magnetoplasmadynamic arcjet with hydrogen propellant. To obtain the velocities of both atoms and ions, laser absorption spectroscopy was employed for atom, and time-of-flight technique was used for ions. In a quasi-steady operation at 13kA/0.65g/s, larger ions velocity (33km/s) than that of the atoms (13km/s) was found in the case of flared anode configuration, which implies that large mean free path between the ions and atoms prohibited momentum transfer from the ions to the neutral particles. This velocity differen...

    DOI: 10.2322/jjsass.53.215

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  122. LE-7A FTP Full Load Cold Run

    KINEFUCHI K, UCHIUMI M, INOUE M, HIRATA K

    航空原動機・宇宙推進講演会講演集(CD−ROM)   Vol. 45th   page: 22033   2005

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  123. Diagnosis of plasma flow field of MPD arcjet.

    杵淵紀世志, 船木一幸, 清水幸夫, 都木恭一郎

    宇宙輸送シンポジウム 平成14年度     page: 257-260   2003

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Research Project for Joint Research, Competitive Funding, etc. 5

  1. 中温超伝導体による液体水素モビリティイノベーション

    2021.7 - 2022.3

    未来社会創造プロジェクト 

    飯田和昌, 土屋雄司, 畑野敬史, 杵淵紀世志, 吉田隆, 生田博志

  2. 宇宙を誰もが自由にアクセス・利用できる空間へ

    2021.2 - 2021.6

    ムーンショット型研究開発事業 新たな目標検討のためのビジョン公募 

    チームリーダ:稲守孝哉 サブリーダ:杵淵紀世志

  3. 高比推力・高推力を両立する高効率多層ヒータによる電熱型電気推進

    2020.7 - 2021.3

    2020年度宇宙工学委員会戦略的開発研究費 

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    Grant type:Competitive

  4. 薄肉タングステン合金のレーザ溶融積層造形法の開発と宇宙エンジン用電熱ヒータへの応用

    2020 - 2023

    一般研究開発助成 <レーザプロセッシング> 

    杵淵紀世志

  5. 高効率熱交換チューブとボイルオフガス/冷凍機ループによるタンク冷却システム

    2020 - 2021

    宇宙探査イノベーションハブ第6回研究提案募集<B.アイデア型> 

    杵淵紀世志

KAKENHI (Grants-in-Aid for Scientific Research) 5

  1. 極低温マイクロジェット/トランスピレーション冷却による熱防護の革新

    Grant number:21K18779  2021.7 - 2024.3

    日本学術振興会  科学研究費助成事業 挑戦的研究(萌芽)  挑戦的研究(萌芽)

    杵淵 紀世志

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    Authorship:Principal investigator  Grant type:Competitive

    Grant amount:\6370000 ( Direct Cost: \4900000 、 Indirect Cost:\1470000 )

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  2. 前方流速場nudgeによる衝撃波変調~波面消失原理実証と応用展開

    Grant number:21H04589  2021.4 - 2025.3

    日本学術振興会  科学研究費助成事業 基盤研究(A)  基盤研究(A)

    佐宗 章弘, 太田 匡則, 北村 圭一, 長田 孝二, 杵淵 紀世志, 市原 大輔, 中村 友祐

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    Authorship:Coinvestigator(s)  Grant type:Competitive

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  3. 宇宙プラズマと衛星磁場のイオンスケール相互作用の解明に基づく軌道維持機能の創出

    Grant number:21H01531  2021.4 - 2024.3

    日本学術振興会  科学研究費助成事業 基盤研究(B)  基盤研究(B)

    稲守 孝哉, 杵淵 紀世志, 川嶋 嶺

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  4. 二酸化炭素ホールスラスタの基礎放電特性および推薬供給方法

    Grant number:21K04492  2021.4 - 2024.3

    日本学術振興会  科学研究費助成事業 基盤研究(C)  基盤研究(C)

    張 科寅, 渡邊 裕樹, 杵淵 紀世志

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    Authorship:Coinvestigator(s)  Grant type:Competitive

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  5. ナノ秒パルス放電による超音速流れの剥離抑制に向けた渦生成と過熱機構の解明

    Grant number:20H02350  2020.4 - 2023.3

    日本学術振興会  科学研究費助成事業 基盤研究(B)  基盤研究(B)

    杵淵 紀世志

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    Authorship:Principal investigator 

    Grant amount:\18460000 ( Direct Cost: \14200000 、 Indirect Cost:\4260000 )

    本研究では、超音速飛翔体や空気吸込エンジンの実現に向け、ナノ秒放電プラズマアクチュエータによる衝撃波・境界層干渉による剥離の抑制メカニズムを明らかにし、その実装に向け独自の基盤的貢献を果たす。超音速流中のナノ秒放電現象は、渦生成と流れの過熱の2つの主要メカニズムに支配される。前者は剥離を抑制するよう作用するが、後者は逆に剥離を促進する。これらの現象把握のため、超音速風洞内にナノ秒放電電極を配し、熱線流速計による渦生成の計測と、平行光レーザBOS法による流れの過熱の把握に取り組む。剥離抑制に影響するナノ秒放電電極の流れに対する配置との関係を整理し、基礎理論を構築して実用に資す。

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Industrial property rights 4

  1. 多層断熱材及びそれを用いた断熱方法

    宮北 健, 北本 和也, 杵淵 紀世志, 斎藤 雅規

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    Applicant:国立研究開発法人宇宙航空研究開発機構

    Application no:特願2017-226987  Date applied:2017.11

    Announcement no:特開2019-094016  Date announced:2019.6

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  2. 電熱ヒータ、噴射装置及び宇宙機

    杵淵 紀世志

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    Applicant:国立研究開発法人宇宙航空研究開発機構

    Application no:特願2017-005844  Date applied:2017.1

    Announcement no:特開2018-116803  Date announced:2018.7

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  3. 宇宙航行体用の推進薬タンク及び宇宙航行体

    石川 佳太郎, 青山 太一, 杵淵 紀世志, 沖田 耕一, 更江 渉

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    Applicant:三菱重工業株式会社

    Application no:特願2015-074308  Date applied:2015.3

    Announcement no:特開2016-193662  Date announced:2016.11

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  4. 気液二相の流量計測方法及び二相流量計測装置

    小林 弘明, 田口 秀之, 杵淵 紀世志, 佐藤 哲也, 大平 勝秀

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    Applicant:独立行政法人 宇宙航空研究開発機構, 国立大学法人東北大学

    Application no:特願2013-111961  Date applied:2013.5

    Announcement no:特開2014-232007  Date announced:2014.12

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Teaching Experience (On-campus) 7

  1. 圧縮性流体力学

    2020

  2. 推進システム特論

    2020

  3. 航空宇宙推進工学

    2020

  4. 機械・航空宇宙工学序論

    2020

  5. 推進システム特論

    2019

  6. 航空宇宙推進工学

    2019

  7. 圧縮性流体力学

    2019

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Teaching Experience (Off-campus) 2

  1. 将来型推進工学特論

    2020.12 Muroran Institute of Technology)

  2. エネルギー科学とマネージメントⅢ

    2019.7 Kyushu University)