Updated on 2022/03/31

写真a

 
KAWASAKI Akira
 
Organization
Institute of Materials and Systems for Sustainability Division of Systems Research (DS) Assistant Professor
Graduate School
Graduate School of Engineering
Title
Assistant Professor
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Degree 1

  1. 博士(工学) ( 2016.3   東京工業大学 ) 

Research Interests 4

  1. Space systems

  2. Magnetohydrodynamics

  3. Space propulsion

  4. Detonation

Research Areas 2

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

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

Research History 7

  1. Nagoya University   Institute of Materials and Systems for Sustainability Division of Systems Reserch   Assistant Professor

    2019.6

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    Country:Japan

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  2. Nagoya University   Department of Aerospace Engineering   Assistant Professor

    2016.12 - 2019.5

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  3. Texas A&M University   Department of Aerospace Engineering   Visiting Scholar

    2016.10 - 2016.11

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  4. 宇宙航空研究開発機構   宇宙科学研究所   日本学術振興会特別研究員

    2016.4 - 2016.11

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    Country:Japan

  5. Japan Aerospace Exploration Agency   Institute of Space and Astronautical Science   JSPS

    2016.4 - 2016.11

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  6. Tokyo Institute of Technology

    2015.4 - 2016.3

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    Country:Japan

  7. Tokyo Institute of Technology   Department of Energy Sciences   JSPS

    2015.4 - 2016.3

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Education 3

  1. Tokyo Institute of Technology   Interdisciplinary Graduate School of Science and Engineering   Department of Energy Sciences

    2013.4 - 2016.3

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  2. Tokyo Institute of Technology   Interdisciplinary Graduate School of Science and Engineering   Department of Energy Sciences

    2011.4 - 2013.3

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

    2007.4 - 2011.3

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Professional Memberships 1

  1. THE JAPANESE SCOIETY OF MECHANICAL ENGINERRS

    2020

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Committee Memberships 5

  1. The Japan Society for Aeronautical and Space Sciences   member  

    2020.4 - 2021.3   

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    Committee type:Academic society

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  2. The Japan Society of Mechanical Engineers   member  

    2020.4 - 2021.3   

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    Committee type:Academic society

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  3. 電気学会   電磁界応答流体によるエネルギー・環境技術の新展開に関する調査専門委員会 委員  

    2019.7   

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    Committee type:Academic society

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  4. 日本航空宇宙学会 中部支部   幹事(会計)  

    2018.3 - 2020.2   

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    Committee type:Academic society

  5. 日本航空宇宙学会   中部支部 幹事(会計)  

    2018.3 - 2020.2   

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Awards 7

  1. Flame Photo

    2020.12   Combustion Society of Japan  

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  2. 若手優良発表賞

    2020.3   電気学会 電力・エネルギー部門 新エネルギー・環境技術委員会  

    川﨑央

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  3. Best presentation award

    2018.8  

    KAWASAKI Akira, KASAHARA Jiro, INAKAWA Tomoya, MATSUOKA Ken, KAWASHIMA Hideto, MATSUO Akiko, FUNAKI Ikkoh

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  4. Excellent paper award A

    2014.3   The Institute of Electrical Engineers of Japan  

    KAWASAKI Akira

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  5. Best Presentation Award

    2013.12   Academy for Co-creative Education of Environment and Energy Science, Tokyo Institute of Technology  

    KAWASAKI Akira

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  6. American Astronautical Society Award

    2013.6   29th International Symposium on Space Technology and Science  

    KAWASAKI Akira

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  7. Student Presentation Award

    2013.3  

    KAWASAKI Akira

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Papers 25

  1. Cathode temperature measurement of a hydrogen self-field MPD thruster during 1 ms quasi-steady operation Reviewed

    Yuya Oshio, Shitan Tauchi, Akira Kawasaki, Ikkoh Funaki

    Journal of Applied Physics   Vol. 130 ( 17 ) page: 173306 - 173306   2021.11

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:AIP Publishing  

    DOI: 10.1063/5.0063942

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  2. Experimental investigation on a rotating detonation cycle with burned gas backflow Reviewed

    Ken Matsuoka, Masaya Tanaka, Tomoyuki Noda, Akira Kawasaki, Jiro Kasahara

    Combustion and Flame   Vol. 225   page: 13 - 19   2021.3

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:Elsevier BV  

    DOI: 10.1016/j.combustflame.2020.10.048

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  3. A study on detonation-diffraction reflection point distances in H-2/O-2, C2H2/O-2, and C2H4/O-2 systems Reviewed

    Han Sun, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   Vol. 38 ( 3 ) page: 3605 - 3613   2021

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:ELSEVIER SCIENCE INC  

    Recently, Kawasaki and Kasahara (2019) reported that reflection point distance , which is a detonation characteristic length relevant to the diffraction process, is a useful measure; i.e., the critical condition for detonation diffraction can be universally expressed in terms of the diffraction point distance, independent of mixture stability. However, their findings were limited to their experimental conditions only. In this study, we performed high-speed visualization of processes of cylindrical (line-symmetric) detonation diffraction around a 90-degree corner for two series of experiments to obtained reflection point distances, l(r), as a novel characteristic length, and examined critical conditions of reinitiation expressed in terms of the reflection point distance. In the first experimental series, stoichiometric C2H2/O-2 mixtures with 50% Ar dilution were employed, and the channel width l(c) was varied to 5, 10, 15, and 20 mm to investigate the influences of the boundary condition of the flow field. In the second experimental series, H-2/O-2, C2H2/O-2, or C2H4/O-2 mixtures with different equivalence ratios were employed to investigate influences of the reaction systems. Our results confirmed that the channel width does not affect the reflection point distance or the critical condition. The critical condition was also independent of fuel species and equivalence ratio, and can be uniquely expressed as l(r)/l(c) = 4.0 +/- 0.6 in terms of the reflection point distance. (C) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2020.06.371

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  4. Numerical analysis on behavior of dilute water droplets in detonation Reviewed

    Hiroaki Watanabe, Akiko Matsuo, Ashwin Chinnayya, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   Vol. 38 ( 3 ) page: 3709 - 3716   2021

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:ELSEVIER SCIENCE INC  

    Two-dimensional numerical simulations are conducted based on the Eulerian-Lagrangian method to model a gaseous detonation laden with monodispersed water droplets. The premixed mixture is a slightly diluted stoichiometric hydrogen oxygen mixture at low pressure. The outcome of the interactions of the droplet breakup with the cellular instabilities and the non-uniform flow behind the leading front is analyzed. The simulation results are also analyzed using instantaneous flow fields and Favre average profiles for water droplets. Breakup occurs mainly near the detonation front. The mean final diameter of the water droplets at the end of the breakup process is the same regardless of the initial strength of the leading shock or whether it is lower or greater than the Chapman-Jouguet value. The polydispersity comes from local phenomena behind the leading shock, such as forward jets coming from triple point collisions, transverse waves and a combination of both. The total breakup time is longer than that estimated from post-shock conditions and the present finding is in line with the previous experimental results on spray detonation.(c) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2020.07.141

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  5. Investigation of combustion modes and pressure of reflective shuttling detonation combustor Reviewed

    Masato Yamaguchi, Tomoya Taguchi, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara, Hiroaki Watanabe, Akiko Matsuo

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   Vol. 38 ( 3 ) page: 3615 - 3622   2021

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:ELSEVIER SCIENCE INC  

    Detonation combustors are considered promising alternatives to conventional combustors because they offer high thermal efficiency and fast combustion. However, especially for the rotating detonation combustor, the theoretical propulsive performance has not been confirmed in experimental studies because the highly unsteady flow field hinders the measurements process. To understand the involved phenomena in more detail, a reflective shuttling detonation combustor (RSDC) with a rectangular combustion chamber was developed. The interior of the chamber can easily be visualized owing to its two-dimensional quality. Utilizing the RSDC, several combustion tests with gaseous ethylene and oxygen were conducted for different values of mass flow rates and equivalence ratios. Combustion modes from the tests were classified into four types based on the fast Fourier transform (FFT) analysis of the luminous intensity of the CH* self-luminescence images captured by a high-speed camera and a band pass filter. Simultaneously, the theoretical total pressure of a conventional isobaric combustor was compared to the static pressure measured at the bottom of the RSDC chamber. For the detonation modes, the ratio between experimentally measured static pressure and the theoretical pressure varied depending on the location in the chamber owing to the distribution of the time-averaged static pressure. Furthermore, the pressure ratio of the detonation modes was up to 18% lower than that of the deflagration mode potentially owing to the flow velocity induced by the detonation waves. (C) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2020.07.064

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  6. Investigation into the effective injector area of a rotating detonation engine with impact of backflow Reviewed

    K. Goto, R. Yokoo, A. Kawasaki, K. Matsuoka, J. Kasahara, A. Matsuo, I. Funaki, H. Kawashima

    Shock Waves   Vol. 31 ( 7 ) page: 753 - 762   2021

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:Springer Science and Business Media LLC  

    For rotating detonation engines, the high-pressure region behind the detonation causes backflow into the plenum, making it difficult to properly design injectors to achieve the target pressure balance due to blockage of a part of the injector area during engine operation. In this paper, we present the pressure and thrust measurement of a rotating detonation engine with two different triplet injectors (fuel injector diameters of 0.8 mm and 1.0 mm) using gaseous methane, gaseous ethylene, and gaseous oxygen. The detonation wave propagation velocity with the fuel injector diameter of 0.8 mm was approximately 200 m/s higher than that with the fuel injector diameter of 1.0 mm. Combustor pressures and specific impulses were almost identical for both fuel injector diameters in this study. For our evaluation of the extent to which the available injector area can be utilized during engine operation, the effective injector area ratio was defined as the ratio of the plenum pressure during burn time to the pre-ignition value. Regardless of fuel species and fuel injector orifice diameter, the effective injector area ratio decreased proportionally with the ratio of combustor pressure to pre-ignition plenum pressure. This result implies that the pressure balance between the upstream plenum pressure and the combustor pressure can be roughly determined taking the effect of backflow into consideration.

    DOI: 10.1007/s00193-021-00998-9

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    Other Link: http://link.springer.com/article/10.1007/s00193-021-00998-9/fulltext.html

  7. Experimental study of internal flow structures in cylindrical rotating detonation engines Reviewed

    Ryuya Yokoo, Keisuke Goto, Jiro Kasahara, Venkat Athmanathan, James Braun, Guillermo Paniagua, Terrence R. Meyer, Akira Kawasaki, Ken Matsuoka, Akiko Matsuo, Ikkoh Funaki

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   Vol. 38 ( 3 ) page: 3759 - 3768   2021

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    The internal flow structures of detonation wave were experimentally analyzed in an optically accessible hollow rotating detonation combustor with multiple chamber lengths. The cylindrical RDC has a glass chamber wall, 20 mm in diameter, which allowed us to capture the combustion self-luminescence. A chamber 70 mm in length was first tested using C2H4-O-2 and H-2-O-2 as propellants. Images with a strong self-luminescence region near the bottom were obtained, confirming the small extent of the region where most of the heat release occurs as found in our previous research. Based on the visualization experiments, we tested RDCs with shorter chamber walls of 40 and 20 mm. The detonation wave was also observed in the shorter chambers, and its velocity was not affected by the difference in chamber length. Thrust performance was also maintained compared to the longer chamber, and the short cylindrical RDC had the same specific impulse tendency as the cylindrical (hollow) or annular 70-mm chamber RDC. Finally, we calculated the pressure distributions of various chamber lengths, and found they were also consistent with the measured pressure at the bottom and exit. We concluded that the short-chamber cylindrical RDC with equal length and diameter maintained thrust performance similar to the longer annular RDC, further expanding the potential of compact RDCs. (c) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2020.08.001

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  8. Propulsion Performance of Cylindrical Rotating Detonation Engine Reviewed

    Ryuya Yokoo, Keisuke Goto, Juhoe Kim, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki

    AIAA JOURNAL   Vol. 58 ( 12 ) page: 5107 - 5116   2020.12

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:AMER INST AERONAUTICS ASTRONAUTICS  

    This study evaluated the propulsion performance of a nozzleless, cylindrical rotating detonation engine (RDE). Using a C2H4-O-2 mixture, the RDE was tested in a low-back-pressure environment at propellant mass flow rates of 8-45 g/s. In high-speed imaging of the self-luminescence within the combustor, rotating luminous regions were observed at mass flow rates above 22 g/s. Measured pressure distributions suggest that burned gas reached sonic velocity at the combustion chamber outlet. This paper proposes the structure of internal flow in the RDE and confirms that calculated pressure distribution based on the structure was close to the experimental distribution. This study also estimated the RDE's thrust by pressure and momentum exchange and confirmed it by experimental measurement. Moreover, the theoretical thrust calculated under the assumption that exhaust is a sonic flow agreed with the load cell thrusts, suggesting that RDE combustion is perfectly completed inside the chamber. Specific impulses are 80-90% of specific impulses for ideal correct expanded flow for all mass flow rates, and its value was close to that of an annular RDE. In addition, RDE performance will increase by about 20% if the RDE is equipped with a divergent nozzle and the gas is correctly expanded to back pressure.

    DOI: 10.2514/1.J058322

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  9. Numerical analysis of the mean structure of gaseous detonation with dilute water spray Reviewed

    Hiroaki Watanabe, Akiko Matsuo, Ashwin Chinnayya, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara

    Journal of Fluid Mechanics   Vol. 887   2020.3

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:Cambridge University Press (CUP)  

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    DOI: 10.1017/jfm.2019.1018

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  10. Optical Measurement of Fluid Motion in Semi-Valveless Pulse Detonation Combustor with High-Frequency Operation Reviewed

    Akiya Kubota, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara, Hiroaki Watanabe, Akiko Matsuo, Takuma Endo

    COMBUSTION SCIENCE AND TECHNOLOGY   Vol. 192 ( 2 ) page: 197 - 212   2020.2

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:TAYLOR & FRANCIS INC  

    The purge layer of a semi-valveless pulse detonation cycle (PDC) needs to be minimized for operating at a gas-dynamic upper frequency limit. Therefore, it is essential to better understand the process of burned gas backflow for minimizing the purge layer thickness. The flow field of the semi-valveless PDC was visualized to illustrate the movement of burned gas. A combustor of length of 95 mm with a 10-mm-square cross section was used. Supercritical ethylene and oxygen gas were used as fuel and oxidizer, respectively, and the operation frequency was 604 Hz. The unsteady refilling process of the detonable mixture was modeled by an isentropic flow. In addition, the detailed burned gas blowdown process with deflagration-to-detonation transition (DDT) and the backflow were captured. It was shown that the retonation wave generated by the DDT process was the primary trigger of the burned gas backflow. When the duration required for the DDT process was sufficiently shorter than that of the burned gas blowdown process, it was found the latter could be reproduced with approximately 90% accuracy by one-dimensional numerical analysis without the DDT process.

    DOI: 10.1080/00102202.2018.1559837

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  11. Investigation of the measurement characteristics of a multiple-ion-probe method for a propagating methane-oxygen-nitrogen flame Reviewed

    Tomoaki Yatsufusa, Keigo Kii, Naoya Miura, Hiroki Yamamoto, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara

    COMBUSTION AND FLAME   Vol. 211   page: 112 - 123   2020.1

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    The multiple-ion-probe measurement method is a method to measure a propagating flame using ion probes installed on the wall surface of a combustion chamber. The dynamic behavior of the propagating flame along the wall surface can be regenerated from the dataset of flame signals from individual ion probes. Although this method only captures flames near the wall surface, the flame propagation behavior can be indirectly visualized. Because this method can attain very high temporal resolution, it can provide precise measurements of high-speed phenomena such as knocking in spark-ignition engines and detonation in detonation combustors. This study aimed to investigate the ability of a developed 64-channel multiple-ion-probe measurement system to characterize a propagating flame. To this end, three flames with substantially different propagation velocities were measured using the proposed multiple-ion-probe measurement system. During the experiments, methane-oxygen stoichiometric mixtures diluted with different amounts of nitrogen were used. The flame propagation velocity varied within the range of several m/s for a turbulent flame to 2.4km/s for detonation by varying the dilution ratio of nitrogen. In the case where a mixture with a nitrogen mole fraction of 0.71 was used, a phenomenon of repeating stagnation and reacceleration of the propagating flame was observed. Furthermore, the phenomenon considered to be flame quenching was also observed near the wall. In the case of no dilution (nitrogen mole fraction = 0.00), multiple-ion probes with an installation interval of 1.5 mm indicated that the velocity fluctuated within the range of -500m/s to +2000 m/s with respect to the Chapman-Jouguet detonation velocity of 2390 m/s. Experiments involving soot foil recording conducted in parallel confirmed that this velocity fluctuation was derived from the detonation cell structure and that micro-explosions in the detonation front could be captured using the multiple-ion-probe method. (C) 2019 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.combustflame.2019.09.022

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  12. A novel characteristic length of detonation relevant to supercritical diffraction Reviewed

    A. Kawasaki, J. Kasahara

    Shock Waves   Vol. 30 ( 1 ) page: 1 - 12   2020.1

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    Authorship:Lead author   Language:English   Publishing type:Research paper (scientific journal)   Publisher:Springer Science and Business Media LLC  

    For stoichiometric C2H4–O2 and C2H2–O2 mixtures with or without argon dilution, the processes of detonation diffraction have been investigated in a two-dimensional setup through high-speed schlieren imaging, with the characteristic length and the stability of detonation varied by regulating the initial pressure and argon mole fraction of the mixture. In particular, a length relevant to the process of supercritical diffraction (i.e., distance from the channel end corner to reflection point of the transverse detonation on the channel end face, reflection point distance in short) was deduced from obtained sequential schlieren images and analyzed. The reflection point distance can be idealized for the infinitely wide donor channel, and thus, it can be a parameter in which properties intrinsic to each detonable mixture are manifested. Experimental results showed that the reflection point distance was roughly inversely proportional to the initial pressure for identical mixtures and independent of the width of the donor channel at high initial pressures. For a certain combination of the fuel and oxidizer, correlations between the reflection point distance and the initial partial pressure of fuel were very similar regardless of the argon mole fraction. Critical conditions of the diffraction problem could be given for the ratio of the reflection point distance to the channel width, and it was suggested that the critical value lies in a range of 3–5 and does not significantly depend on the stability of the mixture.

    DOI: 10.1007/s00193-019-00890-7

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    Other Link: http://link.springer.com/content/pdf/10.1007/s00193-019-00890-7.pdf

  13. Semi-valveless pulse detonation cycle at a kilohertz-scale operating frequency Reviewed

    Ken Matsuoka, Haruna Taki, Akira Kawasaki, Jiro Kasahara, Hiroaki Watanabe, Akiko Matsuo, Takuma Endo

    COMBUSTION AND FLAME   Vol. 205   page: 434 - 440   2019.7

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    A high operating frequency of a pulse detonation engine is required to increase the thrust-to-engine weight ratio or thrust density. The semi-valveless pulse detonation cycle (PDC) proposed by Matsuoka et al. (2017) can achieve a high operating frequency exceeding several kilohertz. For achieving a higher operating frequency close to the upper limit of gas dynamics, it is necessary to minimize the process in which the buffer layer is applied to avoid self-ignition of the detonable mixture. Experiments were conducted, and a one-dimensional numerical model was developed to investigate the minimum thickness of the buffer layer and the required duration for the stable PDC operation. Ethylene was used as a fuel and pure oxygen as an oxidizer. The total length of two combustors with an inner diameter of 10 mm was 40 and 80 mm. Therefore, the thickness of the buffer layer of approximately 20 mm was suggested for the stable PDC operation. This result indicated that 10% of the duration of one PDC was required to prevent self-ignition (SI). In the failed PDC, the early and late SI were confirmed. Interestingly, high-frequency PDC operation with a short combustor can suppress late SI and results in a higher success rate with the same thickness of the buffer layer. Furthermore, a stable PDC operation of a 1916 Hz with a combustor with a total length of 40 mm was demonstrated. (C) 2019 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.combustflame.2019.04.035

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  14. Propulsive Performance and Heating Environment of Rotating Detonation Engine with Various Nozzles Reviewed

    Keisuke Goto, Junpei Nishimura, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki, Daisuke Nakata, Masaharu Uchiumi, Kazuyuki Higashino

    JOURNAL OF PROPULSION AND POWER   Vol. 35 ( 1 ) page: 213 - 223   2019.1

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:AMER INST AERONAUTICS ASTRONAUTICS  

    Geometric throats are commonly applied to rocket combustors to increase pressure and specific impulse. This paper presents the results from thrust measurements of an ethylene/gas-oxygen rotating detonation engine with various throat geometries in a vacuum chamber to simulate varied backpressure conditions in a range of 1.1-104 kPa. For the throatless case, the detonation channel area was regarded to be equivalent the throat area, and three throat-contraction ratios were tested: 1, 2.5, and 8. Results revealed that combustor pressure was approximately proportional to equivalent throat mass flux for all test cases. Specific impulse was measured for a wide range of pressure ratios, defined as the ratio of the combustor pressure to the backpressure in the vacuum chamber. The rotating detonation engine could achieve almost the same level of optimum specific impulse for each backpressure, whether or not flow was squeezed by a geometric throat. In addition, heat-flux measurements using heat-resistant material are summarized. Temporally and spatially averaged heat flux in the engine were roughly proportional to channel mass flux. Heat-resistant material wall compatibility with two injector shapes of doublet and triplet injection is also discussed.

    DOI: 10.2514/1.B37196

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  15. Critical condition of inner cylinder radius for sustaining rotating detonation waves in rotating detonation engine thruster Reviewed

    Akira Kawasaki, Tomoya Inakawa, Jiro Kasahara, Keisuke Goto, Ken Matsuoka, Akiko Matsuo, Ikkoh Funaki

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   Vol. 37 ( 3 ) page: 3461 - 3469   2019

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    Authorship:Lead author   Language:English   Publishing type:Research paper (scientific journal)   Publisher:ELSEVIER SCIENCE INC  

    We describe the critical condition necessary for the inner cylinder radius of a rotating detonation engine (RDE) used for in-space rocket propulsion to sustain adequate thruster performance. Using gaseous C2H4 and O-2 as the propellant, we measured thrust and impulse of the RDE experimentally, varying in the inner cylinder radius r from 31 mm (typical annular configuration) to 0 (no-inner-cylinder configuration), while keeping the outer cylinder radius (r(o) = 39 nun) and propellant injector position (r(inj) = 35 mm) constant. In the experiments, we also performed high-speed imaging of self-luminescence in the combustion chamber and engine plume. In the case of relatively large inner cylinder radii (r(i) = 23 and 31 mm), rotating detonation waves in the combustion chamber attached to the inner cylinder surface, whereas for relatively small inner cylinder radii (r(i) = 0, 9, and 15 mm), rotating detonation waves were observed to detach from the inner cylinder surface. In these small inner radii cases, strong chemical luminescence was observed in the plume, probably due to the existence of soot. On the other hand, for cases where r(i) = 15, 23, and 31 mm, the specific impulses were greater than 80% of the ideal value at correct expansion. Meanwhile, for cases r(i) = 0 and 9 mm, the specific impulses were below 80% of the ideal expansion value. This was considered to be due to the imperfect detonation combustion (deflagration combustion) observed in small inner cylinder radius cases. Our results suggest that in our experimental conditions, r(i) = 15 mm was close to the critical condition for sustaining rotating detonation in a suitable state for efficient thrust generation. This condition in the inner cylinder radius corresponds to a condition in the reduced unburned layer height of 4.5-6.5. (C) 2018 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2018.07.070

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  16. Supersonic combustion induced by reflective shuttling shock wave in fan-shaped two-dimensional combustor Reviewed

    Masato Yamaguchi, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara, Hiroaki Watanabe, Akiko Matsuo

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   Vol. 37 ( 3 ) page: 3741 - 3747   2019

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:ELSEVIER SCIENCE INC  

    As a novel detonation combustor that differs from a pulse and a rotating detonation engine, a reflective shuttling detonation combustor (RSDC), in which detonation waves shuttle repeatedly, was proposed. In a fan-shaped two-dimensional combustor, detonation waves propagate, repeating attenuation and re-ignition by a shock reflection at the side wall. hi the demonstration experiment, chemiluminescence visualization and pressure measurement with ethylene-oxygen mixture were conducted at the same time. As the result, a single shuttling wave coupled with pressure rise was observed in the combustor. The tangential velocity of the wave was 1526 +/- 12 m/s and approximately 60% of the estimated Chapman-Jouguet velocity of 2513 m/s. The ratio of pressure in front of the wave to one behind the primary wave or the reflected wave was in good agreement with one-dimensional shock theory, and it was suggested that the rapid reaction behind the reflected shock wave sustained the continuous propagation of the shock wave. (C) 2018 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2018.06.210

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  17. Numerical investigation on propagation behavior of gaseous detonation in water spray Reviewed

    Hiroaki Watanabe, Akiko Matsuo, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   Vol. 37 ( 3 ) page: 3617 - 3626   2019

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    Language:English   Publishing type:Research paper (scientific journal)   Publisher:ELSEVIER SCIENCE INC  

    A two-dimensional (2D) numerical simulation is conducted to clarify the propagation behavior of gaseous detonation in a water spray and its structure. The computational target refers to the experiment conducted by G Jarsale et al., and C2H4 -air gaseous detonation propagates where the water droplets (WDs) are sprayed. The parameters used are the C2H4-air equivalence ratio and WD mass fraction. The flow field, Favre-averaged one-dimensional profile, and cellular structure are revealed in 2D simulations. Stable propagation of gaseous detonation is observed in the water spray, and the decrease in velocity relative to the Chapman-Jouguet velocity without WDs is as much as 3.2%. Adding WDs changes the cellular pattern, especially for leaner mixtures. The weak triple point decays. and the cell width increases because of the longer induction length due to decreased velocity. The WD presence changes the detonation flow field substantially, and evaporation occurs primarily at 10 mm behind the shock wave. The high-evaporation region propagates at the detonation speed, and the compression wave formed when the detonation reflects from the two-phase medium propagates backward. Furthermore, WD evaporation suppresses the velocity, vorticity, and temperature fluctuations. Rapid evaporation with WDs leads to lower hydrodynamic thickness than that without WDs or in the Zel'dovich-von Neumann-Doring model. (C) 2018 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2018.07.092

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  18. Numerical Analysis on the Plasma Behavior of a Hydrogen MPD Thruster at the Critical Current Reviewed

    Shitan Tauchi, Akira Kawasaki, Masakatsu Nakane, Kenichi Kubota, Ikkoh Funaki

    JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES   Vol. 67 ( 5 ) page: 159 - 166   2019

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    <p>For a self-field magnetoplasmadynamic (MPD) thruster using hydrogen propellant, plasma flows were numerically simulated with a model including the ion-slip effect. To clarify the thruster behavior near the critical current, the discharge current and the propellant mass flow rate were set to 5 or 10 kA (critical current) and 0.4 g/s, respectively. At the critical current, current paths protruded toward a downstream region due to an increased Hall parameter when compared with the lower current case. In conjunction with this, the pressure was higher in the vicinity of the cathode tip and the ion-slip parameter exceeded unity in the discharge chamber at the critical current. Significant ion-slip heating occurred in the supersonic region, which resulted in limited amount of gas dynamic thrust. </p>

    DOI: 10.2322/jjsass.67.159

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  19. Study on a Reflective Shuttling Detonation Wave in Fan-Shaped Two-Dimensional Combustor Reviewed

    M. Yamaguchi, K. Matsuoka, A. Kawasaki, J. Kasahara, H. Watanabe, A. Matsuo

    Proceedings of the Combustion Institute     page: ***   2018

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  20. The Effect of Anode Configuration on Hydrogen MPD Thruster Performance: A Numerical Study Reviewed

    Tauchi, S, Kawasaki. A, Nakene, M, Kubota, K, Funaki, I

    Trans. JSASS, Aerospace Technology Japan     page: ***   2018

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  21. The Effect of Anode Configuration on Hydrogen MPD Thruster Performance: A Numerical Study Reviewed

    TAUCHI Shitan, KAWASAKI Akira, NAKANE Masakatsu, KUBOTA Kenichi, FUNAKI Ikkoh

    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN   Vol. 16 ( 3 ) page: 274 - 279   2018

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    <p>The flowfields of a self-field magnetoplasmadynamic (MPD) thruster using hydrogen propellant were numerically simulated with a physical model incorporating the ion-slip effect. Thrust performance was investigated for two anode configurations, namely, straight anode and flared anode at discharge currents between 5 to 8 kA. Simulation results show that thrust efficiency increases with increased discharge current for the straight anode, while for the flared anode, thrust efficiency tends to decrease; this opposite trend is caused by the ion-slip effect. When comparing thrust characteristics, thrust for the flared anode was found to be larger than that for the straight anode, but the advantage of the flared anode diminishes at higher discharge currents due to strong pinching and consequent pressure depletion in the vicinity of the flared anode surface. This pressure depletion leads to large electric power consumption owing to the ion-slip heating. That is, at lower pressures, the ion-slip effect becomes more significant because collisions between ions and neutral atoms are not frequent.</p>

    DOI: 10.2322/tastj.16.274

  22. Numerical Study on Discharge Current Path and Performance of a Magnetoplasmadynamic Thruster Reviewed

    Kawasaki Akira, Kubota Kenichi, Funaki Ikkoh, Okuno Yoshihiro

    IEEJ Transactions on Fundamentals and Materials   Vol. 136 ( 3 ) page: 135 - 140   2016

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    The discharge current path and performance of a steady-state, self-field magnetoplasmadynamic (MPD) thruster using argon propellant have been examined under the assumption of temperature distributions on the cathode by axisymmetry two-dimensional magnetohydrodynamic (MHD) flow simulation with electrode sheath model as the boundary condition. The discharge current path in the thruster is affected not only by the Hall effect but also by the distribution of thermionic emission from the cathode. When the cathode temperature is decreased from the tip to the root, the discharge current shifts to the cathode tip, which mitigates the current concentration toward the cathode root due to the Hall effect. Then, the thrust is increased as well as the input power, and the thrust efficiency is almost the same as that under the constant temperature distribution on the cathode.

    DOI: 10.1541/ieejfms.136.135

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  23. Numerical Investigation of Discharge Current Path in a Hydrogen MPD Thruster Reviewed

    Kawasaki Akira, Kubota Kenichi, Funaki Ikkoh, Okuno Yoshihiro

    IEEJ Transactions on Fundamentals and Materials   Vol. 136 ( 3 ) page: 141 - 146   2016

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    Authorship:Lead author, Corresponding author   Language:Japanese   Publishing type:Research paper (scientific journal)   Publisher:The Institute of Electrical Engineers of Japan  

    For a 10N-class, steady-state, self-field magnetoplasmadynamic (MPD) thruster using hydrogen as the propellant, the distribution of discharge current path has been investigated by means of axisymmetry two-dimensional magnetohydrodynamic (MHD) flow simulation including an electrode sheath model with cathode temperature distributions. The discharge current path concentrates in the downstream region of the thruster, particularly on the anode edge and the cathode tip, because the ionization of hydrogen occurs after the dissociation. This feature is so dominant that the cathode sheath voltage is determined mainly by the temperature at the cathode tip and the discharge current path is hardly affected by the temperature gradient of the cathode. These characteristics are quite different from those for argon MPD thruster.

    DOI: 10.1541/ieejfms.136.141

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  24. MHD Simulation and Thermal Design of an MPD Thruster Reviewed

    Akira KAWASAKI, Kenichi KUBOTA, Ikkoh FUNAKI, Yoshihiro OKUNO

    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN   Vol. 12 ( ists29 ) page: Pb_19 - Pb_25   2014

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    Authorship:Lead author, Corresponding author   Language:English   Publishing type:Research paper (scientific journal)   Publisher:Japan Society for Aeronautical and Space Sciences  

    DOI: 10.2322/tastj.12.pb_19

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  25. Thermal Design of a Self-Field Argon MPD Thruster by Numerical Calculation Reviewed

    Akira KAWASAKI, Kenichi KUBOTA, Ikkoh FUNAKI, Yoshihiro OKUNO

    JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES   Vol. 61 ( 6 ) page: 167 - 173   2013

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    A 400-kW-class steady-state self-field magnetoplasmadynamic (MPD) thruster is numerically designed with a combination of magnetohydrodynamic (MHD) and thermal analyses, where a heat flux evaluated from the MHD analysis is imposed on the electrode as a boundary condition in the thermal analysis. The increase in the ratio of an anode radius to a cathode radius improves the thrust performance, but can rise the temperature locally at an anode downstream edge and a cathode tip due to the concentration of discharge current and/or insufficient heat removal. It is suggested, however, that a thruster without electrode melting is realizable even at such a high input power by setting an appropriate cathode radius and enhancing heat removal from the electrode by means of heat pipe. The thruster designed under the thermal constraint is expected to achieve a thrust of 17 N, a specific impulse of 990s, a thrust efficiency of 21% for argon propellant.

    DOI: 10.2322/jjsass.61.167

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MISC 120

  1. Experimental investigation on a rotating detonation cycle with burned gas backflow

    Ken Matsuoka, Masaya Tanaka, Tomoyuki Noda, Akira Kawasaki, Jiro Kasahara

    COMBUSTION AND FLAME   Vol. 225   page: 13 - 19   2021.3

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    To analyze a rotating detonation cycle (RDC) with burned gas backflow, simultaneous self-luminous visualization, pressure, and thrust measurements with gaseous ethylene and oxygen were performed. Three different geometric blockage ratios (bottom-wall-surface area to cross-sectional area of combustor) were set at 89.2, 70.2, and 51.7%. The fuel and oxidizer mass flow rates and equivalence ratio were constant at 20.6g/s, 41.2g/s, and 1.7, respectively. During the combustion test, the single detonation wave rotated at 1557, 1459, and 1353 m/s, and the propagation speed increased proportionally for the geometric blockage ratio. The estimated fuel-oxidizer-based specific impulse was in the range of 148 +/- 8s, and the impact of the geometric blockage ratio and propagation speeds on this specific impulse was not confirmed. The hydrodynamic blockage ratio of the oxidizer injector due to the detonation wave was estimated using the oxidizer plenum pressure. It was found that the hydrodynamic blockage ratio linearly decreased with an increase in the geometric blockage ratio. This important trend suggests that the RDC operation is limited in the region of the lower geometric blockage ratio. It is also predicted that a reduction in the hydrodynamic blockage ratio while maintaining the geometric blockage ratio is required for stable RDC operation and achievement of pressure gain combustion. Moreover, the whole RDC structure including the burned gas back flow successfully visualized at the frame rate of 0.5 and 1 mu s. The validity of estimated hydrodynamic blockage ratio was demonstrated by comparison with the visualization experiment. It was concluded that the hydrodynamic blockage ratio was primarily determined mainly by the time scale of the burned gas backflow. (C) 2020 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.combustflame.2020.10.048

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  2. Space Flight Demonstration of Detonation Kick Motor Using Sounding Rocket S-520

    KASAHARA Jiro, MATSUYAMA Koichi, MATSUOKA Ken, KAWASAKI Akira, WATANABE Hiroaki, ITOUYAMA Noboru, GOTO Keisuke, ISHIHARA Kazuki, MATSUO Akiko, FUNAKI Ikkoh, NAKATA Daisuke, UCHIUMI Masaharu, HABU Hiroto, TAKEUCHI Shinsuke, ARAKAWA Satoshi, MASUDA Junichi, MAEHARA Kenji@@WADA Asato, YAMADA Kazuhiko

        2021.3

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    Language:Japanese   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    3rd Sounding Rocket Symposium (March 24-25, 2021. Online Meeting)

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  3. 観測ロケット S-520 31号機プロジェクト:デトネーションエンジンシステムの宇宙実証

    松岡 健, 笠原 次郎, 松山 行一, 川﨑 央, 伊東山 登, 渡部 広吾輝, 後藤 啓介, ブヤコフ バレンティン, 石原 一輝, 野田 朋之, 松尾 亜紀子, 船木 一幸, 中田 大将, 内海 政春, 羽生 宏人, 竹内 伸介, 荒川 聡, 増田 純一, 前原 健次, 山田 和彦, 和田 明哲, MATSUOKA Ken, KASAHARA Jiro, Matsuyama Koichi, KAWASAKI Akira, ITOUYAMA Noboru, WATANABE Hiroaki, GOTO Keisuke, BUYAKOFU Valentin, ISHIHARA Kazuki, NODA Tomoyuki, MATSUO Akiko, FUNAKI Ikkoh, NAKATA Daisuke, UCHIUMI Masaharu, HABU Hiroto, TAKEUCHI Shinsuke, ARAKAWA Satoshi, MASUDA Junichi, MAEHARA Kenji, YAMADA Kazuhiko, WADA Asato

    観測ロケットシンポジウム2020 講演集 = Proceedings of Sounding Rocket Symposium 2020     2021.3

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    第3回観測ロケットシンポジウム(2021年3月24-25日. オンライン開催)著者人数: 21名資料番号: SA6000162001レポート番号: Ⅰ-1

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  4. Experimental Study on the Thruster Performance and Thermal Characteristics of 2MW-Class Self-Field MPD Thrusters

    TAUCHI Shitan, OSHIO Yuya, KAWASAKI Akira, FUNAKI Ikkoh

    Proceedings of 2021 Symposium on Laboratory Experiment for Space Science     2021.3

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    Language:Japanese   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    2021 Symposium on Laboratory Experiment for Space Science (March 5, 2021. Online Meeting)

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  5. Progress Status of a Detonation Engine System for Sounding Rocket S-520 No. 31: Rotating Detonation Engine

    MATSUOKA Ken, GOTO Keisuke, BUYAKOFU Valentin, ISHIHARA Kazuki, NODA Tomoyuki, ITOYAMA Noboru, KAWASAKI Akira, WATANABE Hiroaki, MATSUYAMA Koichi, KASAHARA Jiro, MATSUO Akiko, FUNAKI Ikkoh, NAKATA Daisuke, UCHIUMI Masaharu, TAKEUCHI Shinsuke, IWASAKI Akihiro, WADA Asato, MASUDA Junichi, ARAKAWA Satoshi, HABU Hiroto, YAMADA Kazuhiko

        2021.1

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    Language:Japanese   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. Study on System Operation Evaluation of Rotating Detonation Engine Using Liquid Oxygen

    伊藤志朗, 石原一輝, 米山健太郎, 伊東山登, 渡部広吾輝, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸, 中田大将, 内海政春, 松井康平, 北川幸樹, 中村秀一, 東野和幸, 福地亜宝郎, 長尾隆央

    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. 53rd-39th   2021

  7. Prediction of Pressure Loss in Injector for Rotating Detonation Engines Using Single-element Simulations

    鈴木寛人, 松尾亜紀子, 大門優, 川島秀人, 川崎央, 松岡健, 笠原次郎

    宇宙航空研究開発機構特別資料 JAXA-SP-(Web)   ( 20-008 ) page: 1 - 10   2021

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    This paper newly proposes a prediction method for estimation of pressure loss in injector for rotating detonation engines with lower cost calculation. This method is supposed to be used to design new injector shapes. Although the conventional full-scale simulation contains plenum, all injectors, and combustion chamber, the computational domain in the prediction method is limited to single-element of injector, which includes single injector and plenum. The effects of combustion chamber including the detonation propagation are given by the time-evolving boundary conditions at injector outlet. The time-evolving boundary conditions are prepared by the full-scale simulation or theoretical calculation of one cycle in RDE, and mass flow rate is determined for single-element simulations. The single-element simulation consists of two phases. First, a non-reactive steady flow is calculated for new injector shape, which is used as an initial condition. Then, time-evolving profile is periodically applied to injector outlet as a boundary condition. Eventually, plenum pressure converges to appropriate value. In this paper, the single-element simulations are carried out for the rectangle injector and the chamfered injector for validation of the new method. The error of pressure loss prediction is less than 0.6%, which is 50 times faster, in comparison with full-scale simulation.

    DOI: 10.2514/6.2020-3879

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  8. Study on Optical Measurement of a Reflective Shuttling Detonation Phenomena

    松岡健, 田口知哉, 渡部広吾輝, 川崎央, 伊東山登, 笠原次郎, 松尾亜紀子

    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. 53rd-39th   2021

  9. A Data-driven Investigation for Prediction of Detonation Characteristic Length based on Observation of Diffracting Detonation Waves

    川崎央, 長谷川大樹, SUN Han, 伊東山登, 渡部広吾輝, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    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. 53rd-39th   2021

  10. Experimental study of Cylindrical Rotating Detonation Engine with Diverging Channel

    中田耕太郎, 太田光星, 石原一輝, 後藤啓介, 伊東山登, 渡部広吾輝, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2020   2021

  11. Experimental investigation of the effect of equivalence ratio on characteristic length of hydrogen-oxygen and hydrocarbon-oxygen mixtures with detonation diffraction

    SUN Han, 川崎央, 伊東山登, 渡部広吾輝, 松岡健, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2020   2021

  12. Experimental Study on the Initiation of Detonation Waves by Reflected Shock Waves in Laser Ignition

    佐藤朋之, 松岡健, 川崎央, 笠原次郎

    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. 53rd-39th   2021

  13. Study on Operative Conditions Determination of Rotating Detonation Combustor Using Ethanol

    米山健太郎, 石原一輝, 伊藤志朗, 渡部広吾輝, 伊東山登, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    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. 53rd-39th   2021

  14. An Experimental Investigation of the Rotating Detonation Rocket Engine Using Additive Manufacturing

    服部花凜, 太田光星, 石原一輝, 後藤啓介, 伊東山登, 渡部広吾輝, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2020   2021

  15. New Approach to Find Low Pressure Loss Injector for Rotating Detonation Engine Using Single-element Simulation

    鈴木寛人, 嶋英志, 松尾亜紀子, 大門優, 川島秀人, 川崎央, 松岡健, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2020   2021

  16. Propulsive performane of cylindrical rotating detonation engine with propellant injection cooing

    Keisuke Goto, Kosei Ota, Akira Kawasaki, Hiroaki Watanabe, Noboru Itouyama, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki

    AIAA Scitech 2021 Forum     page: 1 - 7   2021

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    © 2021, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. An engine cooling concept for cylindrical rotating detonation engine which had an injector surface on the combustor side wall has been tested with three combustor length of 21, 30, and 6 mm. Thrust measurement of the cylindrical RDE (24-mm-diameter) was conducted with monitoring K-type thermocouples inserted in combustor wall. Single rotating detonation wave was observed with the combustor length of 30 and 69 mm in this study. Cooling effect due to the propellant injection was confirmed as the nearly saturated temperature response in the combustor side wall. when the chamber length is more than 30 mm, the specific impulse maintained more than 80% of theoretical value assuming sonic condition at the chamber exit. The result indicated that modest combustor length as an efficient thruster exists in the range of 30 to 69 mm.

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  17. Numerical study on early-times laser controlled detonative propulsion

    Tomoyuki Sato, Andrea Alberti, Alessandro Munafò, Marco Panesi, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara

    AIAA Scitech 2021 Forum   Vol. 1 PartF   page: 1 - 9   2021

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    In this work we investigate applications of laser energy deposition for the control of deflagration-to-detonation transition and for the early stage propagation of flame in pulse detonation combustor. The plasma hydrodynamics are described by the system of chemically reactive Navier-Stokes equations and non-equilibrium effects are described with a two-temperatures model for heavy-particles and free-electrons. The non-equilibrium radiation model for the laser discharge is based on a kinetic approach for the photons (radiative transfer equation formulation). Inverse Bremsstrahlung, multi-photon ionization, breakdown chemical kinetics and shock wave dynamics are accounted for self-consistently. Preliminary simulations were performed for ethylene-oxygen mixture.

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  18. Experimental measurement of torque and force on a rotating detonation engine with six-axis force sensor

    Satoru Sawada, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki

    AIAA Scitech 2021 Forum     page: 1 - 15   2021

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    Rotating detonation Engine (RDE) is highly expected for the future propulsion systems due to its more compact structure than conventional internal combustion engines. It is because detonation waves circling in the order of km/s compress propellant instead of the mechanical complex compressor. Of interest on RDE, the torque around the z axis is important for practical use of the system. Due to the detonation waves, fluid inside the RDE produces friction on the chamber wall, which causes force and torque aside from thrust on RDE. In this study, we measured the torque by introducing the 6-axis force sensor which output the torque and axial force simultaneously. And we observed several modes, some of which were dominated positive or negative propagation duration, and others of which were contained both of propagation. From the results, we clarified the torque closely connected to the propagation of detonation waves in terms of the direction and strength. Moreover, we evaluated the effect on thrust performance of RDE. And we concluded that the contribution to RDE performance loss was effectively zero.

    DOI: 10.2514/6.2021-0295

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  19. Study of Cylindrical Rotating Detonation Engine with Lattice Structure Injector

    太田光星, 鈴木遼太郎, 中田耕太郎, 服部花凜, 伊藤志朗, 石原一輝, 後藤啓介, 伊東山登, 渡部広吾輝, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸, 川島秀人, 松山新吾, 丹野英幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2020   2021

  20. Numerical analysis of the mean structure of gaseous detonation with dilute water spray Reviewed

    Hiroaki Watanabe, Akiko Matsuo, Ashwin Chinnayya, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara

    JOURNAL OF FLUID MECHANICS   Vol. 887   2020.3

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  21. Experimental Study on Thruster Performance and Physical Phenomena of High Power Self-Field MPD Thruster

    TAUCHI Shitan, OSHIO Yuya, KAWASAKI Akira, FUNAKI Ikkoh

        2020.3

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    Language:Japanese   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

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  22. Characterization of a Quasi-Steady Self-Field MPD Thruster with Various Electrode Configurations

    Shitan Tauchi, Yuya Oshio, Akira Kawasaki, Ikkoh Funaki

    AIAA Scitech 2020 Forum     2020.1

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    DOI: 10.2514/6.2020-0191

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  23. Demonstration Study of Pulse Detonation Engine System for Sounding Rocket S-520-31

    BUYAKOFU Valentin, 野田朋之, 澤田悟, JOSEPH Victoria, 後藤啓介, 石原一輝, 渡部広吾輝, 伊東山登, 川崎央, 松岡健, 松山行一, 笠原次郎, 中田大将, 内海政春, 松尾亜紀子, 船木一幸, 竹内伸介, 和田明哲, 岩崎祥大, 羽生宏人

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

  24. Propulsion Performance Evaluation of Reflective Shuttling Detonation Combustor

    田口知哉, 山口聖人, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子

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

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    Detonation combustors, such as a rotating detonation combustor and a pulse detonation combustor, have advantages in its higher thermal efficiency and possibility to configure smaller scale combustor than that of conventional internal combustion systems. In a previous study, we proposed a new detonation combustor named a reflective shuttling detonation combustor (RSDC) in which detonation waves propagate in opposite directions repeating reflection at side walls. In the present study, a rectangular combustor (53 × 45 × 5 mm) with non-premixed triplet injectors was used to clarify the effect of equivalence ratios and mass flow rates on combustion modes and propulsive performance. As a result, both detonation and deflagration modes were observed. These modes were classified into four types (Single, Double, Single strong single weak, and Deflagration modes) based on CH* images captured by a high-speed camera and a band-pass filter whose peak value is 430 nm. For mass flow rates and equivalence ratios, it is suggested that a normalized fill height h/λ, which varies depending on these parameters, affects wave number transition as a rotating detonation combustor. For propulsive performance, static pressure measured at the bottom of the combustion chamber was normalized with the theoretical value of a conventional isobaric combustor. The normalized pressure for detonation modes were lower than that of deflagration modes. This might be attributed to higher dynamic pressure caused by the wave propagation and/or insufficient combustion.

    DOI: 10.2514/6.2020-1171

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  25. Experimental Study of the effect of Injector Orifice on Propulsive Performance of Rotating Detonation Engine

    後藤啓介, 横尾颯也, BUYAKOFU Valentin, 澤田悟, 野田朋之, JOSEPH Victoria, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸, 有松昂輝, 稲積慧, 中田大将, 内海政春, 川島秀人

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

  26. Proposing new Method to Predict Pressure Loss in Injector for Rotating Detonation Engines Development

    鈴木寛人, 松尾亜紀子, 大門優, 川島秀人, 川崎央, 松岡健, 笠原次郎

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

  27. Operating Characteristics of a Rotating Detonation Engine with a Large Inlet Cross Section

    野田朋之, 松岡健, 川崎央, 渡部広吾輝, 伊東山登, 笠原次郎

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 58th   2020

  28. An Experimental Investigation of Influence of Diluents on Reflection Point Distance of Gaseous Detonation

    川崎央, SUN Han, 伊東山登, 渡部広吾輝, 松岡健, 笠原次郎

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 58th   2020

  29. Thrust performance of rectangular-shaped reflective shuttling detonation combustor

    田口知哉, 松岡健, 川崎央, 渡部広吾輝, 伊東山登, 笠原次郎

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 58th   2020

  30. Combustion pressure distributions and thrust performances in small cylindrical rotating detonation engines

    Ryuya Yokoo, Keisuke Goto, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki

    AIAA Scitech 2020 Forum   Vol. 1 PartF   2020

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    The internal flow and wave structure of a nozzle-less, cylindrical rotating detonation engine (RDE) was investigated in this research. For a C2H4–O2 mixture, pressure distributions and chemiluminescence inside the chamber were obtained by combustion experiments. Pressure distributions suggest that combustion region is finished near 20 mm from bottom of the cylindrical RDE, and Mach number distributions obtained by Rayleigh flow theory also reveal flow reaches the sonic speed at the exit of the cylindrical RDE. Chemiluminescence images taken from the side of the cylindrical RDE show that strong luminance area ends at approximately 15-20 mm, which also means that combustion in the RDE finishes around that point. Moreover, a forward-tilting detonation wave which stably rotated at 1414 m/s was observed in the images, and it extended to the downstream of burned gas. From these results of strong luminescence at bottom and shock wave extending to the exit, the flow and wave structure inside cylindrical RDEs are proposed.

    DOI: 10.2514/6.2020-0202

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  31. Experimental Study on a Process of Refilling Mixture in Pulse Detonation Engine

    野田朋之, 松岡健, 川崎央, 笠原次郎

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

  32. Pressure and visualization measurements on pulsed combustion thrustor

    Motomu Asahara, Jiro Kasahara, Ken Matsuoka, Akira Kawasaki., Akiko Matsuo, Ikkoh Funaki

    AIAA Scitech 2020 Forum   Vol. 1 PartF   2020

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    Pulsed combustor, it has features such as taking out work without a compressor and mixing promotion by a complicated flow field. The purpose of my research is to evaluate the thrust of a pulse combustor that substitutes this check valve with an ejector mechanism without moving parts, and to clarify the pressure-gain of pulsed combustion and grasp the combustion state with pressure and visualization measurement. Although the total pressure of exhaust gas was lower than the oxygen plenum pressure in all shots, these were almost equal in case in which the pulse combustion was stably continued. It was also found that the expansion wave and the compression wave reciprocate in the combustion section, and that the behavior of the pressure wave and the flame were synchronized. A pulse combustion cycle due to this pressure wave reciprocation was estimated, and the sound velocity obtained from this was in agreement at the order level with theoretical value.

    DOI: 10.2514/6.2020-0923

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  33. Cylindrical rotating detonation engine cooling by means of propellant injection

    Keisuke Goto, Kosei Ota, Akira Kawasaki, Hiroaki Watanabe, Noboru Itouyama, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki

    AIAA Propulsion and Energy 2020 Forum     page: 1 - 9   2020

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    © 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. An engine cooling concept for cylindrical rotating detonation engine which had an injector surface on the combustor side wall has been tested and demonstrated. Thrust measurement of the cylindrical RDE (24-mm-diameter) was conducted with monitoring K-type thermocouples inserted in combustor wall. Single rotating detonation wave was observed in the testing conditions ranging from 31 to 59 g/s in this study. Combustion tests for 4.0 ~ 4.9 s were successfully done, and all injector side wall temperature increases were suppressed compared to that of combustor base plate, which had no cooling structure. This could be due to the cooling effect by the heat exchange of propellant injection. In the 4.9 s combustion test with 31 g/s, all thermocouples inserted in the combustor side wall which had the propellant injector surface showed a temperature decreasing 2.5 s after ignition even though the combustion was continuing, and implied the combustion mode shift.

    DOI: 10.2514/6.2020-3855

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  34. Development of Space Flight Detonation Engine Demonstrator Using Sounding Rocket S520-31 and Its Evolution

    Kasahara Jiro, Matsuoka Ken, Kawasaki Akira, Goto Keisuke, Yokoo Ryuya, Buyakofu Valentin, Matsuo Akiko, Funaki Ikkoh, Nakata Daisuke, Uchiumi Masaharu, Habu Hiroto, Takeuchi Shinsuke, Yamada Kazuhiko, Kitagawa Koki, Tobe Hirobumi, Iwasaki Akihiro, Wada Asato

        2019.8

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    Language:Japanese   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    2nd Sounding Rocket Symposium (August 5-6, 2019. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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  35. Orbit Injection Demonstration Experiment of Detonation Kick Motor Using Sounding Rocket SS520

    Kasahara Jiro, Matsuoka Ken, Kawasaki Akira, Matsuyama Koichi, Goto Keisuke, Matsuo Akiko, Funaki Ikkoh, Maru Yusuke, Habu Hiroto, Takeuchi Shinsuke, Yamada Kazuhiko, Kitagawa Koki, Tobe Hirobumi, Yamada Kazuhiko, Arakawa Satoshi, Iwasaki Akihiro, Wada Asato, Nakata Daisuke, Uchiumi Masaharu, Endo Takuma, Ishii Kazuhiro, Tokudome Shinichiro, Nonaka Satoshi, Kojima Takayuki, Kawashima Hideto, Shouji Takeshi

        2019.8

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    Language:Japanese   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    2nd Sounding Rocket Symposium (August 5-6, 2019. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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  36. Research and development of rotating detonation engine system for the sounding rocket flight experiment S520-31

    Jiro Kasahara, Akira Kawasaki, Ken Matsuoka, Akiko Matsuo, Ikkoh Funaki, Daisuke Nakata, Masaharu Uchiumi

    AIP Conference Proceedings   Vol. 2121   2019.7

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    A detonation is a combustion wave that propagates at supersonic speed (2∼3 km/s) in a combustible mixture. There are many fundamental studies of detonation waves and detonation engine systems. The detonation cycle has a higher thermal efficiency than a conventional constant-pressure combustion cycle. Therefore, it is expected that a high-efficiency propulsion system can be realized using detonation waves.A rotating detonation engine (RDE) uses continuous detonation propagating at a bottom in an annular combustor. As detonation waves propagate at a supersonic speed only in the bottom region of the RDEs, the combustor can be shortened. However, the combustor needs cooling system due to high heat flux to the combustor wall. In this experimental study, we performed combustion tests of RDE system using gaseous ethylene and oxygen as the propellant. This RDE system performance will also be demonstrated in space environment by the sounding rocket. We measured the combustor pressure, temperatures, heat flus, mass flow rate and thrust. The RDE system used in this study is shown in Figure 1. We performed the long-duration rotating detonation engine combustion tests for at sea level condition. The stable trust histories were obtained.

    DOI: 10.1063/1.5115842

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  37. A Development Research of Methane-Oxygen Pulse Detonation Engine for a Sounding Rocket S-520-31

    BUYAKOFU Valentin, 横尾颯也, 後藤啓介, 川崎央, 松岡健, 笠原次郎, 中田大将, 内海政春, 松尾亜紀子, 船木一幸

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  38. An Experimental Study of an Axial Distribution Structure of Combustion Pressure in a Small Cylindrical Rotating Detonation Engine

    横尾颯也, 後藤啓介, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  39. Research on modes of combustion and propulsive performance of Reflective Shuttling Detonation Rocket Engine

    松岡健, 山口聖人, 田中聖也, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子

    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. 51st-37th   2019

  40. Experimental measurement of torque on rotating detonation engines

    澤田悟, 笠原次郎, 松岡健, 川崎央, 松尾亜紀子, 船木一幸

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  41. 回転デトネーションエンジンにおける内壁の熱制御実験

    KIM Ju-Hoe, 横尾颯也, 後藤啓介, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2018   2019

  42. 回転デトネーションエンジンのインジェクタにおける圧力損失に関する数値解析

    鈴木寛人, 松尾亜紀子, 大門優, 川島秀人, 川崎央, 松岡健, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2018   2019

  43. Numerical simulation on modeled combustion chamber for rotating detonation engines

    鈴木寛人, 松尾亜紀子, 大門優, 川島秀人, 川崎央, 松岡健, 笠原次郎

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  44. An Experimental Investigation of the Effect of Inner Cylinder on the Performance of a Rotating Detonation Rocket Engine

    川崎央, 笠原次郎, 稲川智也, 松岡健, 川島秀人, 松尾亜紀子, 船木一幸

    宇宙航空研究開発機構特別資料 JAXA-SP-(Web)   ( 18-005 )   2019

  45. Combustion mode in rectangular-shaped two-dimensional detonation combustor and its propulsive performance

    山口聖人, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  46. 水液滴を含む混合気中を伝播する気相デトネーションの特性長に関する数値解析

    渡部広吾輝, 松尾亜紀子, CHINNAYYA Ashwin, 松岡健, 川崎央, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2018   2019

  47. An Experimental Study of a Miniature Cylindrical Rotating Detonation Engine with a Film-Cooled Wall

    川崎央, 横尾颯也, KIM Ju-Hoe, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    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. 51st-37th   2019

  48. Heat Transfer Measurement of a Disk-shaped Rotating Detonation Combustor with a Thin Film Resistance Temperature Detector

    堀田貢太郎, 川崎央, 松岡健, 笠原次郎, 松尾亜希子, 船木一幸

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  49. Numerical Investigation for Characteristics of Rotating Detonation Engine with Hollow Combustion Chamber

    山口貴史, 松尾亜紀子, 川崎央, 松岡健, 笠原次郎

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  50. Experimental Study on Shortening Deflagration-to-Detonation Transition by Nanosecond Pulsed Laser Ignition

    佐藤朋之, 松岡健, 川崎央, 笠原次郎

    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. 51st-37th   2019

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    The effect of the ignition position of laser ignition and the initial pressure to the deflagration-to-detonation transition distance and to the early stage of flame propagation in pulse detonation combustor is evaluated by Schlieren visualization method using Nd:YAG laser. Changing these two parameters under pre-mixed, stoichiometric, static ethylene – oxygen gas, it is found that laser ignition can generate strong spherical blast wave before the combustion wave, that igniting at the center of the combustor can form more reflected shock wave’s jumbled area which intrigues flame acceleration, and that flame decelerates before the spherical flame reaches the combustor walls and accelerates after its reflection.

    DOI: 10.2514/6.2019-3872

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  51. Application of Detonation Combustion Technology in Space Propulsion Systems

    川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    電気学会研究会資料   Vol. 2019 ( FTE-19-020-025.027-032 ) page: 51 - 56   2019

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  52. An experimental study on the characteristic length appearing during detonation wave diffraction

    SUN Han, 川崎央, 松岡健, 笠原次郎

    燃焼シンポジウム講演論文集(CD-ROM)   Vol. 57th   2019

  53. セルサイズオーダーの希釈率擾乱がH<sub>2</sub>-O<sub>2</sub>-Arデトネーションの内部構造に与える影響の数値解析

    大平直矢, 松尾亜紀子, 川崎央, 松岡健, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2018   2019

  54. エジェクタ構造を有するデトネーション燃焼器に関する基礎研究

    朝原元夢, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2018   2019

  55. Propulsion performance of inner-cylinder-less rotating detonation engine

    Ryuya Yokoo, Kesisuke Goto, Juhoe Kim, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki

    AIAA Scitech 2019 Forum     2019

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    We evaluated the propulsion performance of a nozzle-less, inner-cylinder-less rotating detonation engine (RDE). For a C2H4–O2 mixture, the RDE was tested in a low-back-pressure environment at several propellant mass flow rates ranging from 8 to 45 g/s. In high-speed imaging of the self-luminescence within the combustor, rotating luminous regions were observed at a mass flow rates greater than 22 g/s. The specific impulse efficiency was greater than 80% for all the mass flow rate, and approximately 90% in some cases, which is comparable with ones in conventional RDEs having inner cylinders. By the control surface analysis, it was clarified that propellants injection from the injector holes and pressure on the bottom of the combustion chamber account for the thrust. It was also suggested that the design of the outer cylinder of the combustion chamber and the injector arrangement may affect the thrust performance. Axial Mach number distributions within the engine were calculated under an assumption of isentropic expansion. As a result, the burned gas reached a sonic or supersonic velocity at the outlet of the combustion chamber. Axial pressure distributions also suggested that the acceleration of the burned gas was completed at a far-upstream region within the combustion chamber.

    DOI: 10.2514/6.2019-1500

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  56. Numerical investigation on characteristic lengths for gaseous detonation with dilute water spray

    Hiroaki Watanabe, Akiko Matsuo, Ashwin Chinnayya, Ken Matsuoka, Akira Kawasaki, Jiro Kasahara

    AIAA Propulsion and Energy Forum and Exposition, 2019     2019

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    Two-dimensional (2D) numerical simulations based on Eulerian-Lagrangian method are conducted to clarify the droplet behavior within its lifetime within the detonation cell. The simulation results are analyzed via 2D instantaneous flow fields and Favre spatiotemporal average technique, by applying the recycling block method. Gaseous detonation with dilute water droplets (WDs) propagates stably with a 4% velocity decrease compared to dry CJ velocity in the simulation conditions. From the instantaneous flow field analysis, the droplet breakup occurs primarily in jets, downstream of the transverse wave, nearby the collision of transverse waves, and the interaction between the transverse wave and the jets. The Favre average one-dimensional profiles by grouping WDs based on the initial shock strength that WDs experience reveal the droplet life inside the cellular structure. The mean equilibrium diameter after the breakup is not affected by the initial shock strength.

    DOI: 10.2514/6.2019-4132

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  57. Experimental performance validation of a rotating detonation engine toward a flight demonstration

    Keisuke Goto, Ryuya Yokoo, Juhoe Kim, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki, Daisuke Nakata, Masaharu Uchiumi

    AIAA Scitech 2019 Forum     2019

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    Thrust measurements of rotating detonation engine of (1) ethylene / gas-oxygen and (2) methane / gas-oxygen with various throat geometries in a vacuum chamber to simulate different back-pressure conditions ranging from 1.1-104 kPa were conducted. For throatless rotating detonation engine, we defined equivalent throat area as the detonation channel area, and then tested four nozzle contraction ratios of 1, 1.5, 2.5, and 8. Engines could be successfully ignited by electric ignitors when initial pressure was high enough to have, at least, one detonation cell in RDE channel. We measured the combustor pressure and reveled that it was almost proportional to the throat mass flux regardless of contraction ratios and the propellant combinations. The specific impulse of methane / gas-oxygen case could achieve 84 ± 1% of ideal specific impulse at the optimum expansion for each back pressure.

    DOI: 10.2514/6.2019-1501

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  58. Analysis of thrust performance and cathode phenomena on a megawatt-class mpd thruster

    Shitan Tauchi, Yuya Oshio, Akira Kawasaki, Kenichi Kubota, Ikkoh Funaki

    AIAA Scitech 2019 Forum     2019

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    The relationship between the thrust performance and the cathode temperature in a megawatt-class, self-field magnetoplasmadynamic (MPD) thruster was investigated experimentally and numerically. For various propellants, i.e. argon, hydrogen, nitrogen, and helium, the thrust performance and cathode temperature were measured at discharge currents ranging from 5 to 12 kA. Measured thrust and thrust efficiency increased with the discharge current. For hydrogen propellant, the highest thrust and thrust efficiency of 28 N and 30%, respectively, were attained at a mass flow rate of 0.4 g/s and a discharge current of 12 kA. Cathode surface temperature also increased with the discharge current. For the hydrogen propellant, the tip of the cathode was particularly heated and the temperature exceeded 4000 K. On the other hand, for the argon and helium propellants, the cathode was heated relatively entirely. Numerical results showed that the current density at the cathode tip increased significantly at high discharge currents because of high hall parameter. This can be a main reason why the cathode surface was heated particularly near the tip for the hydrogen propellant.

    DOI: 10.2514/6.2019-1241

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  59. An experimental study of in-space rotating detonation engine with cylindrical configuration

    Akira Kawasaki, Ju Hoe Kim, Ryuya Yokoo, Keisuke Goto, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki

    AIAA Propulsion and Energy Forum and Exposition, 2019     2019

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    Utilizing detonation in a combustor potentially leads to a compact and propellant-efficient rocket engine. A rotating detonation engine (RDE) is a form of such detonation engine. For a practical realization of the RDE, maturation of its thermal design is essential. In this study, a miniature cylindrical rotating detonation engine was tested in a low-pressure environment (~ 6 kPa) with a film-cooled wall. For the combustor, measurement of lateral wall temperature and high-speed imaging were conducted during combustion. As a result, it has been clarified that there exist an appropriate range in the coolant flow rate to maintain detonation combustion. Additionally, it has been also clarified that film cooling is effective even under detonation combustion.

    DOI: 10.2514/6.2019-4298

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  60. メタン-酸素を用いた回転デトネーションエンジンの推進性能に関する実験研究

    後藤啓介, 横尾颯也, KIM Juhoe, 佐藤朋之, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸, 安田一貴, 八木橋央光, 有松昂輝, 中田大将, 内海政春, 川島秀人

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2018   2019

  61. レーザー点火時に発生する球状衝撃波の挙動に関する実験研究

    佐藤朋之, 松岡健, 川崎央, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2018   2019

  62. Flight Experiment Research of Detonation Engines Using Sounding Rocket

    Kasahara Jiro, Matsuoka Ken, Kawasaki Akira, Matsuo Akiko, Funaki Ikkoh, Nakata Daisuke, Uchiumi Masaharu, Higashino Kazuyuki

        2018.7

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    Language:Japanese   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    1st Sounding Rocket Symposium (July 17-18, 2018. Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency (JAXA)(ISAS)), Sagamihara, Kanagawa Japan

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  63. Research and Development and Flight Experiment for Rotating Detonation Engines

      Vol. 49   page: 3p   2018.4

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  64. 高周波数パルスデトネーションサイクルの可視化研究

    久保田祥矢, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子, 遠藤琢磨

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  65. Experimental Study on Shortening Deflagration to Detonation Transition by Laser Ignition

    佐藤朋之, 松岡健, 川崎央, 笠原次郎

    日本航空宇宙学会中部・関西支部合同秋期大会講演論文集(CD-ROM)   Vol. 55th   2018

  66. An Experimental Study of Detonation Propagation in a Disk Shaped Combustor with Low-Loss Inlet

    東純一, 佐藤芳孝, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    日本航空宇宙学会中部・関西支部合同秋期大会講演論文集(CD-ROM)   Vol. 55th   2018

  67. Experimental Study on Combustion Characteristics of Small Rotating Detonation Engines with Cylindrical Structure

    横尾颯也, 後藤啓介, KIM Juhoe, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    燃焼シンポジウム講演論文集   Vol. 56th   2018

  68. Optical Measurement of a Reflective Shuttling Detonation Cycle

    松岡健, 山口聖人, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子

    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. 50th-36th   2018

  69. 回転デトネーションエンジンのサイドホールを用いた推力方向制御に関する実験的研究

    速水雄規, 西村純平, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  70. 回転デトネーションタービンエンジン自律作動システムに関する研究

    渡邊俊, RHEE Hyun-Seung, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  71. An Experimental Investigation of the Effect of Inner Cylinder on the Performance of a Rotating Detonation Rocket Engine

    川崎央, 笠原次郎, 稲川智也, 松岡健, 川島秀人, 松尾亜紀子, 船木一幸

    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. 50th-36th ( 18 ) page: 91 - 95   2018

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    50th Fluid Dynamics Conference /the 36th Aerospace Numerical Simulation Symposium (July 4-6, 2018. Miyazaki Citizen's Plaza), Miyazaki, JapanIn this study, an inner-cylinder-replaceable rotating detonation rocket engine is tested in a vacuum environment for various inner-cylinder radii r(sub i) ranging from 31 mm (typical lab scale) to 0 mm (no inner cylinder). As a result, it was clarified that there exists a critical inner-cylinder radius to sustain adequate thrust performance in the engine. Detonation waves attached to the inner-cylinder wall for r(sub i) = 23 and 31 mm (supercritical cases), whereas detachment of detonation waves from the inner-cylinder wall was observed for r(sub i) = 0 and 9 mm (subcritical cases), and 15 mm (critical case). In the cases of r(sub i) = 0, 9 mm (subcritical), and 15 mm (critical), strong chemical luminescence was observed significantly in the exhaust plume. For r(sub i) = 15 mm (critical cases), and 23 and 31 mm (supercritical cases), the measured specific impulses were greater than 80% of theoretical values. However, for r(sub i) = 0 and 9 mm (subcritical cases), the measured specific impulses were below 80% of the theoretical values. This is to be attributed to the imperfect detonation combustion observed significantly in the subcritical cases. From these results, we have concluded that the inner-cylinder radius is, at 15 mm, close to the critical condition to keep rotating detonation waves in a favorable state in the combustor.Physical characteristics: Original contains color illustrations

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  72. 回転デトネーションロケットエンジンの推力性能に与える内筒半径の影響に関する実験的検討

    川崎央, 稲川智也, 笠原次郎, 後藤啓介, 松岡健, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  73. 宇宙機ロール制御用S型パルスデトネーションスラスタの低背圧推進性能評価

    鵜飼貴斗, 瀧春菜, 後藤啓介, 西村純平, 東純一, 速水雄規, 松岡健, 川崎央, 笠原次郎, 安田一貴, 森謙太, 八木橋央光, 中田大将, 内海政春, 東野和幸, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  74. 扇形平行2平面燃焼器内を往復伝播するデトネーション波に関する実験的研究

    山口聖人, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  75. 水液滴が噴霧された混合気中を伝播するデトネーションの伝播挙動に関する数値解析

    渡部広吾輝, 松尾亜紀子, 松岡健, 川崎央, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  76. Numerical Investigation on the Effect of Evaporation Behavior on Gaseous Detonation with Water Droplet

    渡部広吾輝, 松尾亜紀子, 松岡健, 川崎央, 笠原次郎

    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. 50th-36th   2018

  77. High Frequency Pulse Detonation Operation with Liquid Fuel

    瀧春菜, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子, 遠藤琢磨

    燃焼シンポジウム講演論文集   Vol. 56th   2018

  78. Experimental study on high-frequency pulse detonation cycle with the liquid fuel

    瀧春菜, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子, 遠藤琢磨

    日本航空宇宙学会中部・関西支部合同秋期大会講演論文集(CD-ROM)   Vol. 55th   2018

  79. 湾曲した壁面と不活性気体に閉じられた予混合気中を伝播するデトネーションに関する数値解析

    重岡俊輔, 松尾亜紀子, 川崎央, 松岡健, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  80. 観測ロケットによるデトネーションエンジン推進飛行実証実験

    笠原次郎, 松岡健, 川崎央, 松尾亜紀子, 船木一幸, 中田大将, 内海政春, 東野和幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  81. 低損失インレットを備える円盤型回転デトネーション燃焼器に関する実験的研究

    東純一, 川崎央, 松岡健, 笠原次郎, 佐藤芳孝, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  82. An Experimental Investigation of a Characteristic Length Scale of the Detonation Through Visualization of Diffraction Process

    川崎央, 笠原次郎

    燃焼シンポジウム講演論文集   Vol. 56th   2018

  83. チャネル内のデトネーション波と燃料液滴との干渉実験

    山田泰平, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  84. An Experimental Study of the Detonation Wave Interacting with Fuel Droplets in a Channel

    山田泰平, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸, 長尾隆央

    日本航空宇宙学会中部・関西支部合同秋期大会講演論文集(CD-ROM)   Vol. 55th   2018

  85. Visualization Study of the Injection and the Structure of Detonation Waves in a Disk-Shaped Combustor by Using Schlieren and Self-luminescence Simultaneous Photograph

    LIU Tailong, 堀田貢太郎, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    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. 50th-36th   2018

  86. シュリーレン光と自発光の同時撮像による平板型回転デトネーション燃焼器内流動の可視化

    堀田貢太郎, LIU Tailong, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2017   2018

  87. Study on novel Detonation Detection Technology using Ion Probe

    田中聖也, 瀧春菜, 山口聖人, 松岡健, 笠原次郎, 川崎央, 八房智顯

    燃焼シンポジウム講演論文集   Vol. 56th   2018

  88. Preliminary experiments on rotating detonation rocket engine for flight demonstration using sounding rocket

    Keisuke Goto, Junpei Nishimura, Junichi Higashi, Haruna Taki, Takato Ukai, Yuki Hayamizu, Koyo Kikuchi, Taihei Yamada, Shun Watanabe, Koutaro Hotta, Tomoya Inakawa, Akiya Kubota, Masato Yamaguchi, Toshiki Daicho, Akira Kawasaki, Ken Matsuoka, Jiro Kasahara, Akiko Matsuo, Ikkoh Funaki, Kazuki Yasuda, Kenta Mori, Hiromitsu Yagihashi, Daisuke Nakata, Masaharu Uchiumi, Kazuyuki Higashino

    AIAA Aerospace Sciences Meeting, 2018   ( 210059 )   2018

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    Rotating detonation engine (RDE) is a promising candidate for future upper stage motor because of its high theoretical efficiency and short combustor length. Technological demonstration in space will be necessary to evaluate the performance and flight dynamics of RDE as a space propulsion system. Toward a flight test, we built an open-structure, rocket RDE-powered vehicle to measure thrust to validate a feasibility of an onboard propulsion system. We carried out 4.4 s combustion test, and achieved over 73% of ideal specific impulse above 1.1 MPa of static pressure in the RDE combustor.

    DOI: 10.2514/6.2018-0157

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  89. Rotating Detonation Engine Experiments during 10-Second Operation and in Low Ambient Pressure

      Vol. 48   page: 4p   2017.4

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  90. Numerical Study on Anode Configuration Effect on Hydrogen MPD Thruster Performance

    Tauchi Shitan, Nakane Masakatsu, Kawasaki Akira, Kubota Kenichi, Funaki Ikkoh

        2017.1

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    Language:Japanese   Publisher:Institute of Space and Astronautical Science, Japan Aerospace Exploration Agency(JAXA)(ISAS)  

    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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  91. 遠心コンプレッサ-軸流タービン付き回転デトネーションエンジンに関する実験的研究

    RHEE Hyun-Seung, 石山勢, 東純一, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2016   2017

  92. Experimental Study of Disk-Shaped Rotating Detonation Combustor with a Cylindrical Wall Injector

    東純一, 川崎央, 松岡健, 笠原次郎, 佐藤芳孝, 松尾亜紀子, 船木一幸

    燃焼シンポジウム講演論文集   Vol. 55th   2017

  93. Magnetohydrodynamic simulation of an applied-field magnetoplasmadynamic thruster

    矢野慶人, 川崎央, 奥野喜裕

    電気学会新エネルギー・環境研究会資料   Vol. FTE-17 ( 16-29 ) page: 67 - 72   2017

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  94. 圧縮機-燃焼器-タービンを単円盤に配置した回転デトネーションエンジンの自律作動に関する実証研究

    石山勢, 東純一, RHEE Hyun-Seung, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2016   2017

  95. Numerical Parametric Study of Electrode Configuration for Hydrogen MPD Thruster Performance

    田内思担, 川崎央, 中根昌克, 窪田健一, 船木一幸

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

  96. 液体燃料の相転移を伴うパルスデトネーションサイクルのキロヘルツ作動

    瀧春菜, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子, 遠藤琢磨

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2016   2017

  97. 液体燃料・酸化剤噴霧を伴うパルスデトネーションサイクルに関する実験的研究

    廣田成俊, 川崎央, 松岡健, 笠原次郎

    衝撃波シンポジウム講演論文集(CD-ROM)   Vol. 2016   2017

  98. ディスク型回転デトネーションタービンエンジンのサイクル特性に関する実験的研究

    東純一, 川崎央, 松岡健, 笠原次郎, 松尾亜紀子, 船木一幸, 森合秀樹

    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. 49th-35th   2017

  99. パルスデトネーション燃焼器の高周波数作動に関する実験的研究

    瀧春菜, 松岡健, 川崎央, 笠原次郎, 渡部広吾輝, 松尾亜紀子, 遠藤琢磨

    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. 49th-35th   2017

  100. Numerical Study on Scale Effect of an MPD Thruster

    佃麻里子, 川崎央, 奥野喜裕

    電気学会新エネルギー・環境研究会資料   Vol. 2016 ( 39 ) page: 53 - 58   2016.11

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  101. Coupled MHD and Thermal Simulation of a Self-Filed MPD Thruster

      Vol. 60   page: 6p   2016.9

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  102. Numerical Analysis on Anode Shape Dependence of Hydrogen MPD Thruster Performance

      Vol. 60   page: 6p   2016.9

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  103. A Numerical Investigation on Scaling of a Self-field MPD Thruster

      Vol. 60   page: 6p   2016.9

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  104. 電磁流体・陰極シース連成解析によるMPDスラスタの性能スケーリングに関する検討

    佃麻里子, 川崎央, 奥野喜裕

    電気学会全国大会講演論文集(CD-ROM)   Vol. 2016   2016

  105. Performance Prediction of MPD Thruster by MHD Simulation Considering Electrode Phenomena

    Kawasaki Akira, Kubota Kenichi, Funaki Ikkoh, Okuno Yoshihiro

    宇宙航空研究開発機構特別資料 JAXA-SP-   Vol. 46th-32nd ( 14 ) page: 179 - 184   2015.3

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    46th Fluid Dynamics Conference / 32nd Aerospace Numerical Simulation Symposium (July 3-4, 2014. Hirosaki Bunka Center), Hirosaki, Aomori, JapanFor a steady-state self-field magnetiplasmadynamic (MPD) thruster (ZT3 thruster, which was investigated experimentally at the Institute of Space System of the University of Stuttgart), a magnetohydrodynamic (MHD) simulation of a plasma flow is conducted under consideration for electrode phenomena by incorporating a theoretical cathode sheath/presheath model into an MHD fluid model as a boundary condition. The influence of the incorporation of the cathode sheath/presheath model on numerical performance prediction is discussed for the operation in a propellant (argon) flow rate of 2.0 g/s and a discharge current of 10 kA. By incorporating the cathode sheath/presheath model, the predicted discharge voltage (21.0 V) agreed well with the experimental result (ca. 20 V). Estimated averagevoltage drop within the cathode sheath/presheath accounts for 31% of the discharge voltage. It was therefore confirmed by numerical simulation that the existence of the cathode sheath significantly affects the operation of the MPD thruster.Physical characteristics: Original contains color illustrations

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  106. Numerical Investigation of Discharge Current Path in a Magnetoplasmadynamic Thruster

    川崎央, 藤本悠太, 窪田健一, 船木一幸, 奥野喜裕

    電気学会新エネルギー・環境研究会資料   Vol. FTE-15 ( 36-51 )   2015

  107. MPDスラスタ内の放電電流経路に関する数値シミュレーション

    藤本悠太, 川崎央, 奥野喜裕

    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. 47th-33rd   2015

  108. Coupled MHD and Thermal Simulation of an MPD Thruster

    Akira Kawasaki, Kenichi Kubota, Ikkoh Funaki, Yoshihiro Okuno

    51st AIAA/SAE/ASEE Joint Propulsion Conference   Vol. 59th   page: 6p   2015

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    Operation of a steady-state, self-field magnetoplasmadynamic (MPD) thruster was simulated numerically by a plasma flow solver coupled with a cathode thermal model through a cathode sheath model. In this paper, influences of the coupling of magneto-hydrodynamic (MHD) and thermal models are discussed. By the coupling, self-adjusted flow field of propellant and temperature distribution of the cathode was obtained. Cathode temperature is almost constant in the downstream part of the thruster, where most of discharge current is attached.

    DOI: 10.2514/6.2015-3727

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  109. MHD Flow Simulation of an MPD Thruster with Electrode Phenomena

      Vol. 58   page: 6p   2014.11

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  110. 電極現象を考慮した電磁流体シミュレーションによるMPDスラスタの性能予測

    川崎央, 窪田健一, 船木一幸, 奥野喜裕

    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. 46th-32nd   2014

  111. Analysis of self-field MPD thrusters for designing a megawatt-class in-space propulsion system

    Ikkoh Funaki, Ken'ichi Kubota, Akira Kawasaki, Yoshihiro Okuno, Kenji Miyazaki, Shun Takenaka, Hideyuki Horisawa

    50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference 2014     2014

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    Preliminary analysis of a mega-watt-class self-field MPD thruster head is conducted to conceptually design an MPD thruster system. It was found that thruster design was severely limited by the temperatures of electrode materials, and hence, thrust efficiency of the thruster head is restricted by heat ejection capability. To improve heat rejection capability, heat pipes are employed to thermally connect the electrodes and radiation panels. Through parametric survey of various thruster configurations, thrust efficiency as much as 38% was obtained for an Isp of 3,900s for hydrogen propellant.

    DOI: 10.2514/6.2014-3418

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  112. Numerical simulation of plasma flow in a self-field MPD thruster coupled with electrode sheath

    Akira Kawasaki, Kenichi Kubota, Ikkoh Funaki, Yoshihiro Okuno

    50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference 2014     2014

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    Plasma flows in a 100-kWe-class, steady-state, self-field magnetoplasmadynamic (MPD) thruster were simulated by a plasma flow solver coupled with an electrode sheath model, which enables us to evaluate electrode fall voltages quantitatively. In this paper, influences of the coupling with the electrode sheath model on discharge pattern are discussed as well as dependences of thruster performances on the propellant mass flow rate and the discharge current. By the coupling, it is shown that a thrust is not significantly affected while a discharge voltage is increased attributed to a cathode fall voltage comparable with a potential fall just in the bulk plasma. The thrust and discharge voltage evaluated with the electrode sheath roughly agree with existing experimental results. For an argon mass flow rate of 2.0 g/s and a discharge current of 8 kA, the average cathode fall voltage was estimated to be 7.1 V, which is comparable with the average bulk fall voltage (7.1 V). Thus, it can be said that energy consumption within the cathode sheath is a significant loss factor of the MPD thruster.

    DOI: 10.2514/6.2014-3696

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  113. Influence of Cathode Dimensions on Thrust Performance and Electrode Temperature of an MPD Thruster

    KAWASAKI Akira, KUBOTA Kenichi, FUNAKI Ikkoh, OKUNO Yoshihiro

    電気学会新エネルギー・環境研究会資料   Vol. 2013 ( 35 ) page: 41 - 46   2013.9

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  114. MPDスラスタの熱設計とその検証実験のための予備的検討

    川﨑 央, 宮崎 兼治, 佐藤 博紀, 窪田 健一, 堀澤 秀之, 船木 一幸, 奥野 喜裕, Kawasaki Akira, Miyazaki Kenji, Sato Hiroki, Kubota Kenichi, Horisawa Hideyuki, Funaki Ikko, Okuno Yoshihiro

    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     page: 1 - 5   2013.1

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    Language:Japanese   Publisher:宇宙航空研究開発機構宇宙科学研究所(JAXA)(ISAS)  

    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, 神奈川県自己誘起磁場型MPDスラスタの実機開発において最も重要な課題の1つとなる熱設計について、アルゴン推進剤を対象に推進剤流れの電磁流体解析とスラスタヘッドの熱解析を連携して行うことで数値的に検討を行った。推進剤流れの電磁流体解析とスラスタヘッドの熱解析を連携させる本研究の解析手法は、スラスタの形状パラメータ変化による電極への着弧様態の変化に追従した放電室の温度分布の詳細な解像を可能とする。スラスタヘッドでは特に陽極出口端および陰極先端で高温化がみられた。陰極半径を減少させると、推進性能は向上する一方で、Hall効果の顕著化に起因する陽極出口端の更なる高温化および熱通過率の低下に起因する陰極先端の更なる高温化に至り、陰極径の減少は熱的には厳しくなる状況を招くことがわかった。形態: カラー図版あり形態: PDF資料番号: AA0061856085レポート番号: STEP-2012-002

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  115. MPDスラスタの数値的性能評価と熱解析

    川崎央, 窪田健一, 船木一幸, 奥野喜裕

    航空原動機・宇宙推進講演会講演論文集(CD-ROM)   Vol. 53rd   2013

  116. Magnetohydrodynamic Plasma Flow and Thermal Design in an MPD Thruster

    KAWASAKI Akira, KUBOTA Kenichi, FUNAKI Ikkoh, OKUNO Yoshihiro

    電気学会新エネルギー・環境研究会資料   Vol. 2012 ( 17 ) page: 43 - 48   2012.9

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  117. 宇宙用MPDスラスタのプラズマ電磁流体解析と熱設計

    川崎央, 窪田健一, 船木一幸, 奥野喜裕

    数値流体力学シンポジウム講演論文集(CD-ROM)   Vol. 26th   2012

  118. Experiments of High Temperature Inert Gas Plasma MHD Power Generation with a Shock-Tunnel Driven Faraday Type Generator

    ZHUANG Yunqin, KAWASAKI Akira, MURAKAMI Tomoyuki, OKUNO Yoshihiro

      Vol. 2011 ( 9 ) page: 19 - 24   2011.8

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    CiNii Books

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  119. Experimental studies of seed-free pure-inert-gas working MHD power generation

    Y. Okuno, K. Watanabe, A. Kawasaki, T. Murakami

    42nd AIAA Plasmadynamics and Lasers Conference     2011

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    We describe the first MHD power generation experiment with seed-free high temperature inert gas (argon) plasma in a linear shaped Faraday type generator. The working gas with a total pressure and total temperature of 0.15 MPa and 9000 K, and a thermal input to the generator of 0.18 MW was produced using a single-pulsed shock tunnel. A magnetic flux density of 4.0 T was applied by a superconducting magnet. The almost nofluctuating electric output power has been obtained and the maximum power output is 19.7 kW, which corresponds to an enthalpy extraction ratio of 11.0% and a power density of about 240 MW/m3 (at the 4th electrode region). These values are comparable to those in the conventional seeded plasma MHD generator. The higher performance can be expected under optimizations of generator shape and working conditions. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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  120. Faraday 形発電機を用いた高温希ガスプラズマMHD発電実験

    川崎央

    平成23年電気学会全国大会講演論文集   Vol. 7 ( 7 ) page: 42 - 42   2011

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Presentations 4

  1. Research and Development of Rotating Detonation Engine System for the Sounding Rocket Flight Experiment S520-31 International conference

    Jiro Kasahara, Akira Kawasaki, Ken Matsuoka, Akiko Matsuo, Ikkoh Funaki, Daisuke Nakata, Masaharu Uchiumi

    8TH BSME INTERNATIONAL CONFERENCE ON THERMAL ENGINEERING  2019  AMER INST PHYSICS

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    Event date: 2019

    Language:English   Presentation type:Oral presentation (general)  

    A detonation is a combustion wave that propagates at supersonic speed (2 similar to 3 km/s) in a combustible mixture. There are many fundamental studies of detonation waves and detonation engine systems. The detonation cycle has a higher thermal efficiency than a conventional constant-pressure combustion cycle. Therefore, it is expected that a high-efficiency propulsion system can be realized using detonation waves. A rotating detonation engine (RDE) uses continuous detonation propagating at a bottom in an annular combustor. As detonation waves propagate at a supersonic speed only in the bottom region of the RDEs, the combustor can be shortened. However, the combustor needs cooling system due to high heat flux to the combustor wall. In this experimental study, we performed combustion tests of RDE system using gaseous ethylene and oxygen as the propellant. This RDE system performance will also be demonstrated in space environment by the sounding rocket. We measured the combustor pressure, temperatures, heat flus, mass flow rate and thrust. The RDE system used in this study is shown in Figure 1. We performed the long-duration rotating detonation engine combustion tests for at sea level condition. The stable trust histories were obtained.

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  2. ファラデー形発電機を用いた衝撃波管駆動高温希ガスプラズマMHD発電実験

    庄 雲欽, 川崎 央, 村上 朝之, 奥野 喜裕

    電気学会研究会資料. FTE, 新エネルギー・環境研究会  2011.8.26 

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    Language:English   Presentation type:Oral presentation (general)  

  3. Influence of Cathode Dimensions on Thrust Performance and Electrode Temperature of an MPD Thruster International conference

    KAWASAKI Akira, KUBOTA Kenichi, FUNAKI Ikkoh, OKUNO Yoshihiro

    2013.9.26 

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    Language:Japanese   Presentation type:Oral presentation (general)  

  4. Magnetohydrodynamic Plasma Flow and Thermal Design in an MPD Thruster International conference

    KAWASAKI Akira, KUBOTA Kenichi, FUNAKI Ikkoh, OKUNO Yoshihiro

    2012.9.6 

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Works 1

  1. 観測ロケットS-520-31号機 DES

    2016.12
    -
    2021.7

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Research Project for Joint Research, Competitive Funding, etc. 3

  1. デトネーション燃焼を応用した小型・高性能な二元推進剤ロケットエンジンの開発

    2017.10

    日東学術振興財団研究助成 

    川﨑央

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    Grant type:Competitive

  2. 観測ロケット・ランダー用革新的デトネーション推進機構の研究

    2016.12

    宇宙工学委員会戦略的研究費 

    笠原次郎

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    Grant type:Competitive

  3. 革新的な高熱効率を有する自発予圧縮機構付き回転デトネーションエンジンの研究開発

    2016.12 - 2017.2

    エネルギー・環境新技術先導プログラム 

    笠原次郎

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    Grant type:Competitive

KAKENHI (Grants-in-Aid for Scientific Research) 14

  1. Application of Detonation to High-performance, Miniature Rocket Engine

    Grant number:19K15209  2019.4 - 2022.3

    Japan Society for the Promotion of Science  Grants-in-Aid for Scientific Research Grant-in-Aid for Early-Career Scientists  Grant-in-Aid for Early-Career Scientists

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    Authorship:Principal investigator 

    Grant amount:\4290000 ( Direct Cost: \3300000 、 Indirect Cost:\990000 )

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  2. デトネーション燃焼を応用した小型・高性能な二元推進剤ロケットエンジンの開発

    2017.12 - 2018.11

    日東学術振興財団  研究助成 

    川﨑 央

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  3. Downsizing of efficient bi-propellant rocket engine by using detonation combustion

    Grant number:17H06741  2017.8 - 2019.3

    Japan Society for the Promotion of Science  Grants-in-Aid for Scientific Research Grant-in-Aid for Research Activity Start-up  Grant-in-Aid for Research Activity Start-up

    Kawasaki Akira

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    Authorship:Principal investigator 

    Grant amount:\2860000 ( Direct Cost: \2200000 、 Indirect Cost:\660000 )

    This study aimed to downsize bi-propellant rocket engines by utilizing detonation combustion characterized by significantly fast combustion completion. In rotating detonation engines (RDEs), which are a type of rocket engines utilizing the detonation combustion, the combustors typically had annular configurations and consisted of two coaxial cylinders to stabilize detonation waves within the combustor. However, through investigation of the effects of combustor sizes on the engine performance, it has been revealed that a simple cylindrical configuration, in which the combustor consists only of one cylinder, can exert equivalent engine performances to conventional annular combustors by setting the propellant flow rate appropriately for the combustor cross-sectional area.

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  4. 宇宙用MPDスラスタのプラズマ電磁流体・熱連成モデルの構築と実機設計への展開

    2015.4 - 2017.3

    日本学術振興会  科学研究費補助金 特別研究員奨励費 

    川﨑 央

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  5. Study on Self-compression Type Detonation Propulsion: Evolutionary Space-Flight Demonstration Study Using Sounding Rockets

    Grant number:19H05464  2019.4 - 2024.3

    Japan Society for the Promotion of Science  Grants-in-Aid for Scientific Research Grant-in-Aid for Specially Promoted Research  Grant-in-Aid for Specially Promoted Research

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  6. 航空宇宙機用タービン革新:デトネーション半径方向末広超音速タービンの物理解明

    2018.10 - 2022.3

    日本学術振興会  科学研究費補助金 国際共同研究加速基金(国際共同研究強化(B)) 

    笠原 次郎

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  7. 航空宇宙機用タービン革新:デトネーション半径方向末広超音速タービンの物理解明

    Grant number:18KK0127  2018.10 - 2020.3

    日本学術振興会  科学研究費助成事業 国際共同研究加速基金(国際共同研究強化(B))  国際共同研究加速基金(国際共同研究強化(B))

    笠原 次郎, 松尾 亜紀子, 船木 一幸, 松岡 健, 川崎 央

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    申請者らの超小型デトネーション燃焼器による超音速ジェットを、米国研究者らが着想した超音速流ラジアルタービン(半径方向外向きに流れながら、流路は拡大し,速度は増加する)に吹き込むことで,極めて小型で高効率な航空宇宙用(超軽量・高効率)タービンが実現可能との見通しを得ている。しかしながら、そのような小型デトネーション燃焼器+超音速流ラジアルタービンの物理機構は十分理解できていないない。そこで、本研究では、(1) 申請者らは小型デトネーション燃焼器を製作し、スロートなしの超音速ジェット生成メカニズムの解明を担当し、(2) 米国側は新規着想の超音速タービンの機構を解明し、流れ係数、段負荷ファクターなどタービン特性を求め、さらに、(3) 申請者らと米国側研究者で,米国側研究室にて装置を結合し、小型デトネーション燃焼器+超音速流ラジアルタービンのシステム物理実証を行い、システムとしての物理特性を解明する。
    平成30年度、令和元年度は、月1回の会議を開催しつつ日本側は小型デトネーション燃焼器の設計・製作を行い高い燃焼効率を達成した。米国側は、新規着想の超音速タービンの機構を設計し、供給部等を含め、可視化研究を実施した。さらに、2019年9月19日、20日の米国(Purdue大)での共同での実験実施に向けて独自インターフェイスの検討・設計・製作を行った。日本側では、予定された流量での燃焼試験(透明円筒管による可視化・軸方向の圧力分布計測等)を完全に完了し、AIAA Journalへ論文投稿中である。

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  8. Experimental Study on Detonation Wave Actuator

    Grant number:17K18937  2017.6 - 2019.3

    Japan Society for the Promotion of Science  Grants-in-Aid for Scientific Research Grant-in-Aid for Challenging Research (Exploratory)  Grant-in-Aid for Challenging Research (Exploratory)

    Kasahara Jiro

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    We developed a rotating-detonation-engine actuator without an inner cylinder and measured the thrust characteristics of a high-speed jet of the engine, and confirmed that a thrust of 90% or more of the theoretical performance (assuming a proper expansion) was generated. When the control signal is input ON, the valves of the fuel (ethylene) and the oxidizer (oxygen) are opened, and the gas is injected into the actuator combustor and ignited to generate a rotational detonation wave. A jet can be generated, and it is configured to stop when it is turned off. We achieved that the response time of the engine was 100 ms, and the Isp was 242 sec. The results were also confirmed by visualization measurement. In addition, a nitrogen film cooling mechanism was developed, and it was experimentally confirmed that the heat flux to the wall can be suppressed.

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  9. 革新的な自律圧縮爆轟物理機構の解明:多孔壁噴射器付円盤回転デトネーションエンジン

    2017.4

    科学研究費補助金  基盤研究(B)

    笠原次郎

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  10. 機体とエンジンの融合を目指す:デトネーションアクチュエータの研究

    2017.4

    科学研究費補助金  研究成果公開促進費 (研究成果公開発表)

    笠原次郎

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  11. 革新的な自律圧縮爆轟物理機構の解明:多孔壁噴射器付円盤回転デトネーションエンジン

    Grant number:17H03480  2017.4 - 2020.3

    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)

    笠原 次郎, 松尾 亜紀子, 船木 一幸, 松岡 健, 川崎 央

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    Authorship:Coinvestigator(s)  Grant type:Competitive

    (1)自律圧縮爆轟現象の昇圧メカニズムの解明研究に関しては、低圧力損失インレットの設計を実施した。(2)自律的な圧力増加の限界値の解明研究に関しては、圧力ゲインを目指した内側噴射型の回転デトネーション燃焼器の設計を実施した。(3)多孔冷却壁面構造のデトネーションエンジンの熱的特性の解明研究に関しては、フィルム冷却型のインジェクターを搭載した回転デトネーション燃焼器の設計を実施した。以上の設計と並行して、ロケットシステムの検討をあわせて実施し、革新的な回転デトネーションロケットエンジンの実現への道筋を明かにした。これらの研究成果は、ICDERS等の国際会議の基調講演・招待講演で発表し、高く評価されている。

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  12. 革新的な高熱効率を有する自発予圧縮機構付き回転デトネーションエンジンの研究開発

    2015.2 - 2017.2

    新エネルギー・産業技術総合開発機構  エネルギー・環境新技術先導研究プログラム 

    笠原 次郎

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    Grant type:Competitive

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  13. 観測ロケット・ランダー用革新的デトネーション推進機構の研究

    2014.8 - 2019.3

    宇宙航空研究開発機構  戦略的開発研究(工学) 

    笠原 次郎

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    Grant type:Competitive

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  14. Study of MPD Thruster System for Future High-power Space Explorations

    Grant number:26289328  2014.4 - 2017.3

    Japan Society for the Promotion of Science  Grants-in-Aid for Scientific Research Grant-in-Aid for Scientific Research (B)  Grant-in-Aid for Scientific Research (B)

    Funaki Ikkoh, OSHIO Yuya, KUBOTA Kenichi, KAWASAKI Akira, MIYAZAKI Kenji, TONOOKA Satoshi, TAUCHI Shitan

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    For an orbital transfer from the Earth to Mars, a high power and high performance spacecraft electric propulsion system is required in particular for a cargo-vehicle since electric propulsion's high specific impulse enables a large payload ratio, and hence a large mass can be conveyed to Mars within limited launch vehicle capability. In this study, 100-N-class Magnetoplasmadynamic (MPD) arcjet system was conceptually designed, and the thruster head design was experimentally demonstrated by using a pulsed operation, in which a capacitor bank as a power source. Such an test can prevent a steady operation that requires large and continuous power and vacuum facilities.

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Teaching Experience (On-campus) 5

  1. 航空宇宙創造設計

    2018

  2. 熱力学及び演習

    2018

  3. 熱力学及び演習

    2017

  4. 機械・航空宇宙工学実験第2

    2017

  5. 航空宇宙創造設計

    2017

Teaching Experience (Off-campus) 9

  1. Design Practice 4

    2020 Nagoya University)

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    Level:Undergraduate (specialized)  Country:Japan

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  2. Exercises on Aircraft International Development Project

    2020 - 2021 Nagoya University)

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    Level:Postgraduate  Country:Japan

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  3. Exp. and Workshop Practice in Mech. and Aerospace Eng. 2

    2019 Nagoya University)

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    Exp. and Workshop Practice in Mech. and Aerospace Eng. 2(1.0 credits)
    Code : 11045
    Course Type :
    Class Format : Experiment
    Course Name : Department of Mechanical and Aerospace Engineering
    Starts 1 : 3 Autumn Semester
    Elective/Compulsory : Compulsory

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  4. Exp. and Workshop Practice in Mech. and Aerospace Eng. 1

    2019 Nagoya University)

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    Level:Undergraduate (specialized)  Country:Japan

    Exp. and Workshop Practice in Mech. and Aerospace Eng. 1(1.0 credits)
    Code : 11044
    Course Type :
    Class Format : Experiment
    Course Name : Department of Mechanical and Aerospace Engineering
    Starts 1 : 3 Spring Semester
    Elective/Compulsory : Compulsory

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  5. Seminar on Propulsion Energy Systems Engineering

    2017 Nagoya University)

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    Level:Postgraduate 

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  6. Advanced Experiments and Exercises in Propulsion Energy Systems Engineering

    2017 Nagoya University)

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    Level:Postgraduate  Country:Japan

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  7. Aerospace Innovative Design

    2017 - 2019 Nagoya University)

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    Level:Undergraduate (specialized)  Country:Japan

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  8. Thermodynamics with Exercises

    2017 - 2019 Nagoya University)

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    Level:Undergraduate (specialized)  Country:Japan

    Thermodynamics with Exercises(2.5 credits)
    Code : 420
    Course Type : Basic Specialized Courses
    Class Format : Lecture and Exercise
    Course Name : Department of Mechanical and Aerospace Engineering
    Starts 1 : 2 Spring Semester
    Elective/Compulsory : Compulsory

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  9. Exp. and Workshop Practice in Mech. and Aerospace Eng. 2

    2016 - 2018 Nagoya University)

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    Level:Undergraduate (specialized)  Country:Japan

    Exp. and Workshop Practice in Mech. and Aerospace Eng. 2(1.0 credits)
    Code : 629
    Course Type :
    Class Format : Experiment
    Course Name : Aerospace Engineering
    Starts 1 : 3 the latter term
    Elective/Compulsory : Compulsory

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Social Contribution 1

  1. 2020年度 JAXA相模原キャンパス 特別公開

    Role(s):Organizing member

    JAXA 宇宙科学研究所  2020年度 JAXA相模原キャンパス 特別公開  2021.3

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    Type:University open house

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Academic Activities 1

  1. ASME Turbo Expo

    Role(s):Peer review

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    Type:Peer review 

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