2024/05/14 更新

オザワ コウヘイ
小澤 晃平
OZAWA Kohei
Scopus 論文情報  
総論文数: 0  総Citation: 0  h-index: 8

Citation Countは当該年に発表した論文の被引用数

所属
大学院工学研究院 機械知能工学研究系
職名
准教授
外部リンク

研究キーワード

  • 流体力学

  • 燃焼学

  • ロケット工学

研究分野

  • フロンティア(航空・船舶) / 航空宇宙工学

取得学位

  • 東京大学  -  博士(工学)   2017年03月

学内職務経歴

  • 2022年02月 - 現在   九州工業大学   大学院工学研究院   機械知能工学研究系     准教授

  • 2017年03月 - 2022年01月   九州工業大学   大学院工学研究院   機械知能工学研究系     助教

学外略歴

  • 2015年04月 - 2017年03月   東京大学   工学系研究科   日本学術振興会特別研究員   日本国

  • 2024年02月 - 2027年02月   ミシガン大学   航空宇宙工学科   客員研究員   客員研究員   アメリカ合衆国

  • 2022年02月 - 2024年02月   ミシガン大学   日本学術振興会 海外特別研究員   客員研究員   アメリカ合衆国

所属学会・委員会

  • 2020年07月 - 現在   火薬学会   日本国

  • 2020年06月 - 現在   宇宙工学委員会   日本国

  • 2018年03月 - 現在   日本機械学会   日本国

  • 2014年06月 - 現在   アメリカ航空宇宙学会   アメリカ合衆国

  • 2013年06月 - 現在   日本航空宇宙学会   日本国

研究経歴

  • 液化ガス燃料を用いた革新的爆轟推進の研究

    回転デトネーションエンジン、パルスデトネーションエンジン

    研究期間: 2024年04月  -  現在

  • 3次元プリンタを用いたリアルタイム性能の自己評価機能を持つハイブリッドロケットの研究

    ハイブリッドロケット、リアルタイム燃料後退速度測定、機能性固体燃料、エンジンフィードバック制御

    研究期間: 2019年04月  -  現在

  • デトネーションを用いた姿勢制御装置の研究開発

    デトネーション、姿勢制御スラスタ

    研究期間: 2018年09月  -  現在

  • 低粘性液化燃料を生ずる境界層燃焼の加速度環境下における燃焼機構解明

    加速度環境, 境界層燃焼, 混相流, 超臨界流体

    研究期間: 2018年04月  -  現在

論文

  • Visualization of the physical destruction process of additive-manufactured regression sensor structures 査読有り

    Kohei Ozawa, Tomoya Neki, Nobuyuki Tsuboi

    Science and Technology of Energetic Materials ( 火薬学会 )   84 ( 2 )   24 - 31   2023年05月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    DOI: 10.34571/stem.84.2_24

    DOI: 10.34571/stem.84.2_24

  • Experimental and Numerical Study on Disc-RDE: Relation between Number of Detonation Wave and Pressure 国際誌

    Koichi Hayashi A., Ohno K., Ishii K., Shimomura K., Tsuboi N., Ozawa K., Jourdaine N.H., Dzieminska E., Obara T., Maeda S., Mizukaki T.

    AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022   2022年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Disc-RDE is studied experimentally and numerically to figure out how single-wave mode and double-wave mode appear. From the measurement, we can pick up the better condition for the better pressure gain result. The experimental results are obtained that rotating detonation waves mainly propagate near the outer circumference of the combustion chamber. A transition of the wave number from single to double is found to exist at the equivalence ratio of 0.67 to 0.84. However, the normal transition of wave number is from double to single transition at the equivalence ratio of 1.45 and higher. Numerical work starts with a single-wave system to turn to double-wave system, then single-wave system back due to the ignition, bifurcation, and so on. Depending on the condition, a single-wave becomes a double-wave, which is confirmed at both the experiment and numerical analysis. The reduction of detonation velocity is confirmed at both experiment and numerical calculation up to 7 % when a single detonation becomes double one.

    DOI: 10.2514/6.2022-0837

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85123355591&origin=inward

  • Attempts of Time-Resolved Fuel Regression Measurements of an Altering-Swirling-Oxidizer-Flow-Type Hybrid Rocket Motor 国際誌

    Ozawa K., Omiya K., Tsuboi N., Ishii K.

    AIAA AVIATION 2022 Forum   2022年01月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    This paper presents efforts to acquire the time-resolved fuel regression data of an Altering-Swirling-Oxidizer-Flow-Type (A-SOFT) hybrid rocket motor using polylactic acid (PLA)-based segmented solid fuels as well as the steady-state fuel regression behavior of the A-SOFT with PLA-based solid fuel. The sensing segments are manufactured by the multi-material Fused Filament Fabrication. They have a ladder-shaped resistor structure to detect the fuel regression in the form of rapid voltage changes caused by its rung destruction. Despite the successful firings of the motor, the verification of the temporally resolved fuel regression was imperfect. The destructions of the first and final rungs provided rapid voltage change, but the other rungs did not provide a rapid voltage change when they were expected to be broken according to the time-averaged regression rate of the segment. The major factors causing the unexpected sensor voltages can include chamber pressure and thermal environments in the motor because the sensor structure had not been tested under these conditions. The spatiotemporal-averaged fuel regression behavior was approximated by a fuel-mass-flux-based linear regression, reflecting the variety of filling rates in additive-manufactured solid fuels. The approximation agreed with the resultant fuel regression behavior, with the residual standard deviation of 2.97×10-2 mm/s. The segmentation of the solid fuels provided the ease of post-experimental measurements of axial distribution of regression rates.

    DOI: 10.2514/6.2022-3565

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85135041692&origin=inward

  • Numerical Simulation on Rotating Detonation Engine: Effect of Number of Injection Ports in Non-Premixed H2-O2 Gases 国際誌

    Yoshidomi K., Kurita N., Ozawa K., Tsuboi N., Hayashi A.K., Kawashima H.

    AIAA Science and Technology Forum and Exposition, AIAA SciTech Forum 2022   2022年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    The effect of the number of injection ports in H2-O2 non-premixed gases on the performance of rotating detonation engine is estimated using two-dimensional numerical simulation with a detailed H2-O2 chemical reaction model. The differences in the instantaneous flow field, specific impulse, detonation velocity, and pressure gain are discussed among three configurations with the different numbers of injection ports. The detonation disappears for the 40-port configuration, while it continues to propagate stably for 60-and 80-port configurations. During the detonation propagation, the specific impulse has small, but significant difference depending on the number of injection ports. Detonation velocity is lower than the CJ velocity and its difference increases as the number of injection ports decreases. Pressure gain using oxygen has a maximum value of-0.084 at the 60-port configuration.

    DOI: 10.2514/6.2022-1112

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85123582558&origin=inward

  • Regression behavior of polylactic acid manufactured by fused filament fabrication for hybrid rocket propulsion 査読有り

    Kohei Ozawa, Han-wei Wang, Takefumi Inenaga, Nobuyuki Tsuboi

    Science and Technology of Energetic Materials ( 火薬学会 )   82 ( 6 )   170 - 177   2021年12月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    Advanced closed-loop control of thrust and mixture ratio of hybrid rockets is planned by combing a fuel mass flow rate control technique with a real-time fuel regression measurement. The latter technique is enabled by a multi-materialadditive-manufactured solid fuel with an integrated sensor probe structure. This work investigated the fuel regression behavior of polylactic acid (PLA) fuel used for the main material of this type of solid fuels. Rectangular slab fuels and cylindrical fuel grains were manufactured by Fused Filament Fabrication (FFF) and fired in an optically accessible slab burner and a lab-scale motor, respectively. The cylindrical fuel grains had more than 2 times larger regression rates than rectangular slab fuels even with the same PLA filament. One of the main factors causing this large difference can be an anisotropy in the fuel regression rate behavior of the solid fuels manufactured by FFF.

    DOI: 10.34571/stem.82.6_170

    Kyutacar

    CiNii Article

    CiNii Research

    その他リンク: http://hdl.handle.net/10228/00008705

  • ハイブリッドロケットの推力 - 混合比制御のノミナル性能向上を詳細に評価

    小澤 晃平

    宇宙航空研究開発機構宇宙科学研究所年次要覧2020年度 ( 国立研究開発法人宇宙航空研究開発機構 宇宙科学研究所 )   23 - 23   2021年11月

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    担当区分:筆頭著者   記述言語:日本語   掲載種別:記事・総説・解説・論説等(大学・研究所紀要)

  • Time-resolved fuel regression measurement function of a hybrid rocket solid fuel integrated by multi-material additive manufacturing 査読有り 国際誌

    Kohei Ozawa, Han-wei Wang, Takuro Yoshino, Nobuyuki Tsuboi

    Acta Astronautica   187   89 - 100   2021年06月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    The function of time-resolved fuel regression measurement is integrated into additively manufactured hybrid rocket fuels and conceptually demonstrated through the comparison with time-resolved optical fuel regression measurements during burns. This function is enabled by multi-material additive manufacturing of mainly thermoplastic solid fuels with a ladder-shaped resistor made of a conductive thermoplastic. This functionalization has many advantages, as indicated by its wide range of applicability, including solid fuel grains with emerging complex geometries, cost-effectiveness, ease of implementation, and minimal influence on the original propulsive performance and fuel regression behavior. Fuel regression is detected by a step change in the voltage applied to the ladder-shaped resistor, indicating the breakage of its top fuel surface rung during a burn. For ignited cases, all the rung breaks were detected, while the fuel regression was also optically measured using high-speed video. Errors were observed between the fuel regression time-histories measured using the two methods; however, most of these errors could be explained by the distribution of the fuel regression in the camera direction and the camera resolution and the melted fuel layer thickness. Considering these uncertainties, the remaining errors that could not be explained by these factors were less than ±0.15 mm, corresponding to ±1 layer of additive manufacturing applied to the prototyping of the solid fuels.

    DOI: 10.1016/j.actaastro.2021.06.031

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85108881693&origin=inward

  • Experimental and numerical study on disc-rde: Flow structure and its performances 国際誌

    Hayashi A.K., Ishii K., Watanabe T., Tsuboi N., Ozawa K., Jourdaine N.H., Dzieminska E., Tang X., Obara T., Maeda S., Mizukaki T.

    AIAA Scitech 2021 Forum   1 - 12   2021年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    The present study discusses disc-type rotating detonation engine (DRDE) experimentally and numerically. The experimental work shows that the detonation propagates in three different modes; single, dual, and hybrid. The operating frequency of dual-wave mode is 1.8-2.1 times faster than that of single wave mode. The number of detonation wave can be predicted based on the pressure history and the operating frequency signal. The numerical work shows the performance of 3D numerical analysis of DRDE with uniform injection case and multiport injection case. By increasing the wave number from one to two, the detonation propagation velocity decreases by 18.7 %. The one-detonation head case gives some better performance to the flow than the two-detonation head case. The inlet flow angle to the radial turbine becomes about 50 degrees to the radial direction no matter how large the plenum chamber is.

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85100319111&origin=inward

  • 3D numerical study on flow field in disc-RDE 国際誌

    Hayashi A.K., Shimomura K., Tsuboi N., Ozawa K., Jourdaine N.H., Ishii K., Dzieminska E., Obara T., Maeda S., Mizukaki T.

    AIAA Propulsion and Energy Forum, 2021   2021年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Disc-type rotating detonation engine is studied numerically using the threedimensional compressible Navier-Stokes equations with 9 species and 21 reactions and with an multi-port injection system. Especially the grid resolution using two-detonation case and the process of stabilization for one and two-detonation cases are studied. As the results of grid resolution, the case of larger grid number provides higher pressure due to capturing the pressure oscilation better and that their rotating speed becomes faster. And the detonation velocity becomes faster when the number of grid becomes larger. As the results of the process of stabilization for one and two-detonation cases, two detonation wave propagate by changing back and forth their distance and no matter what the number of detonation wave is given, one or two, the final detonation wave number becomes two.

    DOI: 10.2514/6.2021-3665

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85123365877&origin=inward

  • Numerical simulations on propane/oxygen detonation in a narrow channel using a detailed chemical mechanism: formation and detailed structure of irregular cells 査読有り 国際誌

    Takeshima Naomi, Ozawa Kohei, Tsuboi Nobuyuki, Hayashi Koichi A., Morii Yuhi

    Shock Waves ( Springer )   30 ( 7-8 )   809 - 824   2020年12月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)

    Numerical simulations of two-dimensional inviscid detonations for a stoichiometric propane/oxygen gas mixture are performed using a detailed chemical reaction model. The UC San Diego model which includes 57 chemical species and 268 elementary reactions is mainly used in the present study. It is shown that a grid size of 3 µm can capture important features such as the unburned gas pocket behind the detonation when compared to larger grid sizes. The effects of channel width show that the detonation propagates with the CJ (Chapman–Jouguet) velocity for all cases and for more than 100 times the channel width of 4.5 mm. Increasing the channel width results in an irregular detonation cell structure. A transverse detonation forms with cross-hatching marks on the maximum pressure history. The irregular detonation cell structure forms because both the reduced activation energy and the stability parameter have a value of approximately 10; however, the maximum thermicity in the detonation is one. The free radicals C3H7 and H2O2 play an important role in the propane oxidation under the high temperature in the detonation. The maximum concentration exists at a temperature of 2000–3000 K. The fifth-order WCNS (weighted compact nonlinear scheme) scheme can resolve the contact surface and complicated flow structure behind the detonation front compared to the second-order MUSCL (Monotonic Upstream-centered Scheme for Conservation Laws).

    DOI: 10.1007/s00193-020-00978-5

    Scopus

    その他リンク: https://link.springer.com/article/10.1007/s00193-020-00978-5

  • Performance of Mixture-Ratio-Controlled Hybrid Rockets under Uncertainties in Fuel Regression 査読有り 国際誌

    Kohei Ozawa, Toru Shimada

    Journal of Propulsion and Power ( American Institute of Aeronautics and Astronautics )   37 ( 1 )   86 - 99   2020年10月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    This paper evaluates various sources of oxidizer to fuel mass ratio (O/F) shifts in hybrid rockets and paths (physical phenomena) through which these O/F shifts affect flight performance. Moreover, the performance increase of O/F control in hybrid rockets is evaluated. Vertical launches of O/F uncontrolled and O/F controlled of hybrid sounding rockets were simulated under two uncertainty models of fuel regression behavior based on experimental data: a) systematic errors with a constant deviation within 3σ and 2) random errors subject to a probability distribution. These simulations included all sources of O/F shifts that originated in the fuel regression behavior and all paths through which the O/F shifts affect flight performance. Residual propellant mass and decreases in specific impulse are found to be the dominant causes of performance loss under both uncertainty models. For both cases 1 and 2, the O/F-controlled hybrid rockets maintained the performance expected under nominal fuel regression behavior, whereas the O/F-uncontrolled hybrid rockets had a lower performance by upwards of 6.69 and 4.06% in ΔV for cases 1 and 2, respectively. For case 2, 3008 flight simulations revealed that the worst case of the O/F-controlled hybrid rocket had a 4.06 to 4.49% larger ΔV and 10.5 to 13.3% higher apogee than that of the O/F-uncontrolled hybrid rocket, and that the O/F-uncontrolled hybrid rocket had a 6.61 times larger standard deviation in ΔV. These results mean that the elimination O/F shift in hybrid rockets significantly improves performance, as well as the accuracy and reliability of performance predictions.

    DOI: 10.2514/1.B37970

    Kyutacar

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85098688616&origin=inward

  • Performance of Mixture-Ratio-Controlled Hybrid Rockets for Nominal Fuel Regression(共著) 査読有り 国際誌

    Kohei Ozawa, Toru Shimada

    Journal of Propulsion and Power ( American Institute of Aeronautics and Astronautics )   36 ( 3 )   400 - 414   2020年03月

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    担当区分:筆頭著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    This paper discusses the impacts of oxidizer-to-fuel mass ratio (O∕F) shifts on the flight performance of single-stage sounding rockets using the flight simulations of three scales of O∕F-controlled and O∕F-uncontrolled hybrid rockets under a nominal fuel regression behavior without uncertainty. The flight simulation code includes three factors dependent on the O∕F: Thermodynamic states of the burned gas (theoretical Isp), shifts in c* efficiency, and nozzle throat erosion. In the flight simulations, a thrust control law was applied to increase the apogee and evaluate the effects of O∕F shifts in the thrust curve including throttling. For the best cases in each scale, O∕F-controlled hybrid rockets slightly improved the performance by 2.03–2.42%in the averaged specific impulse. However, the performance of the O∕F-controlled sounding rockets is essentially the same as the O∕F-uncontrolled type under the median regression behavior: Especially when considering the slight increases in the mass and complexity of the oxidizer feed system needed for O∕F control. Considerable scale effects on the throat erosion and theoretical Isp were observed, but that of the c* efficiency was negligible. The improvement of the theoretical Isp was the primary contributor to flight performance, which was responsible for a larger than 70% share in the total Isp increase. The second largest contribution was the improvement of the c_ efficiency with a share of 21.8–24.3%. The O∕F control gave an improvement of throat erosion corresponding to 5.75% in the total Isp increase for the smallest scale; but, with increasing of the scale, the throat area increase ratio became small so that the throat erosion improvement contribution was reduced to 1.21%.

    DOI: 10.2514/1.B37665

    Kyutacar

    Scopus

    その他リンク: https://arc.aiaa.org/doi/abs/10.2514/1.B37665

  • Three-dimensional numerical simulation on hydrogen/air rotating detonation engine with aerospike nozzle: Effects of nozzle geometries 国際誌

    Kurita N., Jourdaine N., Tsuboi N., Ozawa K., Hayashi A.K., Kojima T.

    AIAA Scitech 2020 Forum   1 PartF   2020年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Three-dimensional simulations with a detailed chemical reaction model were performed to evaluate the propulsive performance of rotating detonation engines (RDEs) for H2/Air with conic aerospike nozzles. The simulations aimed to clarify the propulsive performance for various configurations of aerospike nozzle: open/choked or complete/truncated configurations. The governing equations were the Euler equations, and the chemical model was for 9 species and 21 element chemical reactions. Various equivalent ratios from 0.9 to 1.4 were simulated. The results show that the specific impulse (Isp) and the efficiency of the choked nozzle were 4-6% higher than those for the open nozzle. The higher pressure in the chamber due to the decreasing in the throat area resulted in a gain of pressure thrust by the nozzle and consequently, an addition in Isp and the efficiency. The difference in thrust performance between the complete and the truncated nozzle was smaller than 2% because the pressure component on the base surface compensated for the thrust drop.

    DOI: 10.2514/6.2020-0688

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85091953963&origin=inward

  • Numerical simulation on disk rotating detonation engine: Influence of internal flow field structure on performance 国際誌

    Watanabe T., Shimomura K., Jourdaine N.H., Ozawa K., Tsuboi N., Kojima T., Ishii K., Hayashi A.K.

    AIAA Scitech 2020 Forum   1 PartF   1 - 7   2020年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Three-dimensional simulation for hydrogen-oxygen mixtures on Disk Rotating Detonation Engine (DRDE) are performed in order to understand the flowfields inside the DRDE. The propagation patterns of the detonation are visualized for various conditions such as the stagnation pressure, the number of detonation fronts, and the injection system. The numerical results show that the detonation velocity is larger than the C-J velocity for all conditions. The flow angle into the turbine location is approximately 50 degrees against the radial direction. Furthermore, this angle and the inflow velocity are independent of the stagnation pressure. In the case of two-wave detonation, it is found that the two-wave system merges into the one-wave system. The characteristics of the flow field in this study can be applied to the design of the radial turbines in future experiments and simulations.

    DOI: 10.2514/6.2020-2160

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85092418617&origin=inward

  • Recent experimental and numerical study on disc-type rdes 国際誌

    Hayashi K., Tsuboi N., Ozawa K., Watanabe T., Jourdaine N.H., Ishii K., Kawana H., Kuwata W., Ohno K., Obara T., Maeda S., Dzieminska E., Tang X., Mizukaki T.

    AIAA Scitech 2020 Forum   1 PartF   2020年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    The present study discusses disc-type rotating detonation engine (DTRDE) experimental ly and numerically. The experimental work is performed using hydrogen and air and measuring pressure records together with taking a high speed movie to see a direction of rotating detonation. Especially the relation between mass flow rate and operation frequency and that between equivalence ratio and operation frequency. The number of detonation head is figured out from the maximum operation frequency. Then the 3D numerical analysis of DTRDE will provide the flow structure in the DTRDE combustion chamber, high pressure profile, and low mach number profiles.

    DOI: 10.2514/6.2020-1169

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85091925389&origin=inward

  • Accuracy of real-time fuel regression measurement function of a 3d printed solid fuel 国際誌

    Ozawa K., Wang H.W., Inenaga T., Tsuboi N.

    AIAA Propulsion and Energy 2020 Forum   1 - 15   2020年01月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    The accuracy of a real-time fuel regression measurement by an additive manufactured solid fuel was discussed by comparing with optically measured regression acquired in firing experiments for the conceptual demonstration. The solid fuels that succeeded in fuel regression measurement were manufactured by a multi-material 3D printer and had the ladder-shaped resistor structure with the 0.15 mm thickness rungs. The fuel regression is detected by the voltage change in the reference resistor due to the break of rungs. For all experiments, the initial resistance of the ladder-shaped resistor structure was estimated to be as expected, and all rung break was successfully detected by the steep voltage change. The acquired measurement values by the two methods had errors like an offset for all runs. The major factors of these errors were the distribution of the fuel regression in the fuel width direction and the resolution of the high-speed video camera. As a result, it was found that the time histories of the values optically measured for comparison had relatively large uncertainties. The remaining errors that cannot be explained by these major error factors were estimated to be less than ±0.15 mm, which corresponds to two layers of the solid fuels manufactured in this study.

    DOI: 10.2514/6.2020-3741

    Scopus

    その他リンク: https://www.scopus.com/inward/record.uri?partnerID=HzOxMe3b&scp=85091296741&origin=inward

  • Boundary-layer combustion of wax-based fuels at various chamber pressures under two static acceleration environments 国際誌

    Ozawa K., Yoshino T., Wang H.W., Tsuboi N.

    AIAA Propulsion and Energy Forum and Exposition, 2019   2019年01月

     詳細を見る

    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Boundary layer combustion of wax-based fuels was visualized under atmospheric and supercritical chamber pressures under vertical and horizontal setups. The final goal of this research is to clarify the effects of acceleration environments on the performance of wax-based fuels. At the atmospheric pressure, the roll waves of the melted fuel excited propagated over the fuel top surface, and fuel droplets were entrained by the main gaseous flow for the horizontal setup. However the propagation of the roll waves covered at most 55% of the fuel to surface, and the entrained fuel droplets were hardly observed for the vertical setup. Above the critical pressure, the instabilities of the melted fuel were observed for both setups, but it was difficult to find any differences in the behaviors of the melted fuel due to the strong radiation of the fuel-rich flame on the melted fuels even after the various improvements of the optical system and the solid fuel. At supercritical pressures, the diffusion flames and the melted fuels oscillated at the frequency of approximately 600 Hz by sight, which is approximately the same with the natural frequency of the TCG coupling mode. The fast Fourier transformation up to 500 Hz output an intense peak of approximately 311 Hz for all runs with the horizontal setup and the sonic nozzle. A slight difference in the fuel regression behavior was found between the two burner setups. The vertically set burner showed the larger regression rates at supercritical conditions whereas the regression exponent for the horizontal setup agreed with that of existing data with a horizontally set motor. This result suggests body force increases the mass flow rate of the melted fuel film, which was not considered as a fuel mass transfer in the current liquefying hybrid theory.

    DOI: 10.2514/6.2019-4100

    Scopus

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  • Development of a high efficiency system with a rotating detonation engine for a gas turbine engine (RDE-GTE) using pressure gain combustion 国際誌

    Koichi Hayashi A., Tsuboi N., Ozawa K., Ishii K., Obara T., Maeda S., Dzieminska E., Mizukaki T.

    AIAA Scitech 2019 Forum   2019年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    The present study will show a new development of a gasturbine engine with rotating detonation engine. Five universities and a venture company will study characteristics of rotating detonation engines to develop a gasturbine engine with RDE system in order to improve its performance of gas turbine engine, especially its efficiency. We have three teams; the experimental team, numerical team, and gasturbine team. The experimental team will perform several categories; concentric as well as disk type RDE development and measurement of its stable conditions and controlabilities as well as its efficiency; the cooling system of RDE; injection and ignition system; and DDT performance. The numerical group will simulate above experimental categories by real size calculations with both DNS and AMR base. And a micro-gasturbine engine will be run together with RDE system in a near future. We prefer to choose a disc type RDE instead of concentric cylinder type one. Hence the midterm report like presentation will be given at the SciTech meeting, Jan 2019 and 2020.

    DOI: 10.2514/6.2019-1509

    Scopus

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  • Investigation of graphite nozzle erosion in hybrid rockets using N2O/HDPE 国際誌

    Kamps L., Sakurai K., Ozawa K., Nagata H.

    AIAA Propulsion and Energy Forum and Exposition, 2019   2019年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Seventeen static firing tests are carried out on a small-scale hybrid rocket motor using liquid nitrous oxide as the oxidizer to investigate the chemical erosion characteristics of graphite nozzles. Over 200 data are obtained by employing an innovative data reduction method to determine time-resolved values for nozzle throat diameter, nozzle throat pressure, equivalence ratio, nozzle throat wall temperature and more. An analytical model is formed based on previous research, and used to develop an informed empirical formula for experimental correlations. Empirical correlations are shown to predict nozzle erosion rates with a coefficient of determination upwards of 0.81, whereas the analytical model results in a coefficient of determination of only 0.26. Nozzle erosion rates reached upwards of 0.25 mm/s for equivalence ratios around unity, and chamber pressures around 5 MPa. Lastly, the conditions at the onset of erosion are examined, and used to demonstrate, quantitatively, that chemical erosion can be more easily mitigated when using nitrous oxide as the oxidizer than oxygen.

    DOI: 10.2514/6.2019-4264

    Scopus

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  • Three-dimensional numerical simulation of disk rotating detonation engine: Unsteady flow structure 国際誌

    Watanabe T., Jourdaine N.H., Ozawa K., Tsuboi N., Kojima T., Hayashi A.K.

    AIAA Scitech 2019 Forum   2019年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Disk Rotating Detonation Engine (DRDE) simulation using a mixture of hydrogen and oxygen was performed by three-dimensional numerical analysis with detailed chemical reaction model. The computational grid is created based on the DRDE combustor shape designed by Huff et al. and this grid system use three regions. In zone 1 and 2, the inside of the combustor is simulated and in zone 3, the flow behind the nozzle exit is simulated. The combination of coarse and fine computational grids can calculate the propagation of rotating detonation wave inside the DRDE with small computational resources. The present results show that the detonation wave front is not normal to the outer circular wall but has approximately 60 degrees. The maximum pressure near the detonation head on the outer wall has approximately 10 MPa. The Mach number of the exhaust flow becomes approximately 2.5 in this condition.

    DOI: 10.2514/6.2019-1498

    Scopus

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  • Real-time regression rate measurement of an additive-manufactured functional hybrid rocket fuel 国際誌

    Ozawa K., Wang H.W., Yoshino T., Tsuboi N.

    Proceedings of the International Astronautical Congress, IAC   2019-October   2019年01月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    The function of real-time fuel regression integrated into an additively-manufactured hybrid rocket fuel was conceptually demonstrated by firing experiments. This fuel has a complex ladder resistor structure made of a plastic filament including conductive additives in an electrically insulated main plastic fuel. The real-time regression rate measurement is enabled by the change of the voltage applied to the ladder resistor structure when the top rung is broken by the fuel regression. This type of fuel has advantages in the easy acquisition of the distribution of the local regression rates with a greatly small pitch, minimizing the influences on the structural mass, propulsive performance, and fuel regression behavior especially when loading such a measurement system on space vehicles. Prototype slab fuels outputted steep changes in the voltage indicating the break of the rungs when using rungs with the thickness of 0.15 mm in firing experiments. The acquired breaking signals with the positions of the rungs were compared with the results of the fuel surface detection in the high-speed videos. In 3 of 4 successful experiments, the sensing system of the proposing measurement method followed the fuel regression detected by the high-speed videos within 0.5 mm errors, which are in the same range as the roughness of the solid fuels after the burns. The time history of the breaking signals showed a fuel regression at an almost constant rate. The difference of the steady-state regression rates was less than 0.06 mm/s between these two methods, except for run #4, 0.094 mm/s. It was also confirmed that the ladder resistor structures can be a heater to warm the main solid fuel when applying a voltage larger than 5 V to the resistor structure. This function is also useful for the thermal management of solid fuels in space.

    Scopus

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  • A Theoretical Study on Throttle Ranges of O/F Controllable Hybrid Rocket Propulsion Systems(共著) 査読有り 国際誌

    Kohei OZAWA, Toru SHIMADA

    Journal of Fluid Science and Technology   13 ( 4 )   1 - 18   2018年11月

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    担当区分:責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    The characteristics of several O/F control methods for hybrid rocket propulsion have been discussed and theoretically analyzed from the physical properties of propellants and fuel regression behavior. In this research, comparisons have been made among different oxidizer injection methods of Altering-intensity Swirling Oxidizer Flow Type (A-SOFT), Aft-chamber Oxidizer Injection Method (AOIM), and Swirling-AOIM for the throttle range with a constant O/F, design restrictions of the fuel grain, penalties on the adoption of the methods, and suitable scales of the engine. Theoretical analysis on regression rates has revealed that A-SOFT has upper and lower limits of throttle while maintaining a constant O/F whereas AOIM does not have any lower limit, and Swirling-AOIM covers both the throttle ranges. The designing restriction of the fuel grain derived from the regression rate behavior has indicated that A-SOFT using paraffin and oxygen has a potential to maintain 50- 100% throttle range over a burn. The penalties for the adoption of these O/F control methods have also been discussed from the aspects of the increase in the complexity of the system, structural mass, and pressure drop at the injector for the methods using gaseous injection. The pressure drop has quantitatively been evaluated by relating the available swirl strength with the cross-sectional area and gaseous oxidizer mass flux at the injector. This analysis has revealed 5 times difference in the available swirl strength between the gaseous oxygen and the decomposed gas of 90% hydrogen peroxide. The sizing of the 1st stage of the satellite launcher has revealed that A-SOFT and Swirling-AOIM are suitable for small-scale engines with a propellant mass of 100-102 [ton] using paraffin and liquid oxygen whereas AOIM and Swirling-AOIM are suitable for engines with paraffin and 90% hydrogen peroxide.

    DOI: 10.1299/jfst.2018jfst0031

    Kyutacar

    Scopus

    その他リンク: https://www.jstage.jst.go.jp/browse/jfst/-char/en

  • Three-dimensional numerical thrust performance analysis of hydrogen fuel mixture rotating detonation engine with aerospike nozzle(共著) 査読有り

    Nicolas Jourdaine, Nobuyuki Tsuboi, Kohei Ozawa, Takayuki Kojima, A.Koichi Hayashi

    Proceedings of the Combustion Institute   2018年10月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    UK   Dublin   2018年07月29日  -  2018年08月03日

    DOI: 10.1016/j.proci.2018.09.024

    その他リンク: https://www.sciencedirect.com/science/article/pii/S1540748918306291

  • Hybrid Rocket Firing Experiments at Various Axial-Tangential Oxidizer-Flow-Rate Ratios(共著) 査読有り 国際誌

    Kohei OZAWA, Koki KITAGAWA, Shigeru ASO, Toru SHIMADA

    Journal of Propulsion and Power   35 ( 1 )   94 - 108   2018年09月

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    担当区分:責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    A breadboard model (BBM) of an altering-intensity swirling-flow-type (A-SOFT) hybrid rocket engine was recently developed, and static firing experiments of the BBM were performed under various axial and tangential oxidizer mass-flow rates. A-SOFTs are intended to control thrust at an optimal oxidizer-to-fuel-mass ratio and achieve a high baseline of regression rates with a fixed motor configuration. This is possible by controlling both the oxidizer mass-flow rate and the effective geometric swirl number. A simple model based on a continuous and monotonic function of oxidizer mass flux and effective geometric swirl number was able to accurately predict the performance of experiments conducted on the A-SOFT BBM. The local fuel regression behavior in the axial direction of the A-SOFT BBM was shown to be like that of a swirling-oxidizer-flow-type hybrid rocket engine. Combustion efficiency was evaluated indirectly using equations for the efficiencies of thrust and specific impulse to eliminate errors due to local pressure shifts and the centrifugal force in swirling flows. In most cases, this indirect method compensated for the overestimations of c efficiency that resulted by directly using chamber-pressure data.

    DOI: 10.2514/1.B36889

    Kyutacar

    Scopus

    その他リンク: https://arc.aiaa.org/doi/abs/10.2514/1.B36889

  • Experimental Investigation of Fuel Regression Rate of Low-Melting-Point Thermoplastic Fuels in the Altering-Intensity Swirling-Oxidizer-Flow-Type Hybrid Rocket Engine (共著) 査読有り

    Yo KAWABATA, Ayana BANNO, Yutaka WADA, Kohei OZAWA, Toru SHIMADA, Nobuji KATO, Keiichi HORI, Ryo NAGASE

    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN ( 一般社団法人 日本航空宇宙学会 )   16 ( 3 )   267 - 273   2018年05月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)

    <p>In this study, we developed a high-performance and high-thrust hybrid rocket motor using low-melting-point thermoplastic (LT) fuel and swirling oxidizer flow. LT fuel has excellent mechanical and adhesive properties, as well as a high regression rate compared to conventional hybrid rocket fuel. In this study, we conducted several firing tests using swirling oxidizer flow to obtain the fuel regression rate and evaluate its effects on the geometric swirl number (<i>S<sub>g</sub></i>). We determined that the average regression rate of the LT fuel with <i>S<sub>g</sub></i> = 37.3 was ~2.9 times larger than the axial flow test value. The LT fuel was more susceptible to swirling flow than polypropylene, presumably due to the different physical properties of the fuels. In the swirl flow experiment, we confirmed that the local fuel regression rate behind the fuel is uniform, and it differs from the regression rate seen in the axial flow experiment. For the range of oxygen mass flux values <i>G<sub>oxlo</sub></i> = 30–72, <i>ṙ<sub>loave</sub></i> was fitted to a conventional formula. The results of this fit suggested that the local regression rate at the head region of low-melting-point fuel, such as the LT fuel, cannot be represented only by chemical reactions; therefore, the fluid dynamics of liquefied fuel must be included in the model.</p>

    DOI: 10.2322/tastj.16.267

    CiNii Article

    CiNii Research

    その他リンク: https://www.jstage.jst.go.jp/article/tastj/16/3/16_267/_pdf

  • Visualization of boundary layer combustion of wax-based fuels in vertical and horizontal configurations 国際誌

    Ozawa K., Yoshino T., Tsuboi N.

    2018 Joint Propulsion Conference   2018年01月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Efforts have been carried out from both the theoretical and experimental approaches to reveal the effects of body force parallel to the main flow on the diffusion flame in hybrid rocket propulsion with entrainment of liquid fuel droplets. A theoretical study on liquefying fuels such as wax fuels were partially extended for static acceleration environments and acquired interesting results that the static body force scarcely affects the internal ballistics of the entrained fuel droplets but it can affect the thickness of the liquid fuel layer and the fuel regression rate of the entrained droplets. Diffusion flame of two wax fuels at the atmospheric pressure has been visualized with a slab burner and a high-speed camera in the horizontally and vertical configurations to change the body force loaded to the fluid in the slab burner. The instability of the diffusion flames looked similar to each other over all the experiments, and the high-speed video has observed a few burning droplets entrained from the liquid fuel layer in each of the experiments. The monitoring camera observed a lot of unburned fuel droplets from the slab burner vertically installed.

    DOI: 10.2514/6.2018-4928

    Scopus

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  • Effects of O/F shifts on flight performances of vertically launched hybrid sounding rockets 国際誌

    Ozawa K., Shimada T.

    53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017   2017年01月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    Impacts of various types of O/F shifts on flight performances of a single stage sounding rocket with a scale of S-520 sounding rocket series were comprehensively evaluated by flight simulations of O/F controlled and uncontrolled hybrid rockets. Before the simulation, sources of O/F shifts and factors affected by O/F shifts were classified and discussed in order to clarify the respective sets of simulations. O/F shifts along with the median fuel regression rate behaviors, systematic errors of the behaviors, and random errors of fuel regression rates were statistically modelled using multiple regression theroy. Shifts of thermodynamic states of productive gases after combustion, shifts of c∗ efficiency, and nozzle throat erosion were modelled as factors affected by O/F shifts. Operations of propulsion systems were assumed to include throttling. In the presence of O/F shifts along with a median regression rate equation, shifts of thermodynamic states of productive gases were the dominant factor causing performance losses of O/F uncontrolled rockets. In the presence of systematic or random errors of fuel regression rates, residuals of propellants were the other dominant factor to decrease flight performances. Especially in the random error cases, as a result of 3000 times of flight performances, the guaranteed highest altitude of the O/F controlled rocket was 5.2% higher than that of uncontrolled rocket within ±3σ. This result shows that the O/F controlled hybrid rockets have about 2.6% higher acceleration than the O/F uncontrolled rockets. Accuracy of acceleration of the O/F controlled hybrid rocket was 10 times higher than that of the O/F uncontrolled hybrid rocket. These results indicate that the significance of O/F shifts elimination of hybrid rockets from both the aspects of expectancy and accuracy of flight performances.

    DOI: 10.2514/6.2017-5051

    Scopus

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  • Online Magnetometer Calibration in Consideration of Geomagnetic Anomalies Using Kalman Filters in Nanosatellites and Microsatellites 査読有り

    Takaya Inamori, Ryuhei Hamaguchi, Kouhei Ozawa, Phongsatorn Saisutjarit, Nobutada Sako, Shinichi Nakasuka

    Journal of Aerospace Engineering   29 ( 6 )   2016年06月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)

    DOI: 10.1061/(ASCE)AS.1943-5525.0000612

    その他リンク: https://ascelibrary.org/doi/pdf/10.1061/%28ASCE%29AS.1943-5525.0000612

  • Theoretical prediction of regression rates in swirl-injection hybrid rocket engines(共著) 査読有り

    K. Ozawa, T. Shimada

    Progress in Propulsion Physics   2016年06月

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    担当区分:責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    DOI: 10.1051/eucass/201608283

  • Static burning tests on a bread board model of altering-intenisty swirling-oxidizer-flow-type hybrid rocket engine 国際誌

    Ozawa K., Usuki T., Mishima G., Kitagawa K., Yamashita M., Mizuchi M., Katakami K., Maji Y., Aso S., Tani Y., Wada Y., Shimada T.

    52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016   2016年01月

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    担当区分:筆頭著者, 責任著者   記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    A bread board model of Altering-number Swirling Flow Type (A-SOFT) hybrid rocket engine has been newly developed and the static burning tests have been conducted. A-SOFT hybrid rocket engines have controllability of both thrust and O/F without any replacement of components in the engine. This advantage is acquired by controlling oxidizer mass flow rate and effective swirl number. The purpose of this set of experiments is to confirm the continuity, monotonousness and predictability of the performances of A-SOFTs. The A-SOFT BBM showed a favorable fuel regression behavior. The fuel regression averaged along spatial direction fit the shape of regression rate function proposed before the experiments within ±3.5% errors, and the fuel regression rates are continuous and monotonous along swirl number and oxidizer mass flux. The O/F and thrust data also respectively fit the prediction formulas within ±4.2% and ±4.7% errors. c* efficiency is evaluated with the Isp efficiency - nozzle efficiency ratio ηIsp/ ηCF in order to compensate the pressure sensing errors caused by the centrifugal forces of swirling flows. Though ηIsp/ ηCF in the cases of weak swirl injection was clearly larger than in the cases of axial injection, its dependence on effective geometric swirl number was not clear in strong swirl conditions.

    Scopus

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  • Flame emission spectroscopy in a paraffin-based hybrid rocket 国際誌

    Stober K.J., Leccese G., Narsai P., Ozawa K., Cantwell B.

    Proceedings of the International Astronautical Congress, IAC   0   2016年01月

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    記述言語:英語   掲載種別:研究論文(国際会議プロシーディングス)

    A spectrometer was used to measure the emissions emanating from the plume and combustion chamber of a paraffin-based hybrid rocket. Flame emissions were captured between 200-900 nm at numerous points during the 3-7 second ground tests. Time-resolved blackbody emissions were obtained, as well as emission and absorption peaks associated with combustion products, emanating from the plume and combustion chamber. Plume measurements were taken seven inches aft of the nozzle exit plane at the center of the plume. The rocket is described in Narsai et al. [1] and utilizes paraffin and additives as fuel along with gaseous oxygen as the oxidizer. Chamber pressure and oxidizer flow rates were varied in order to validate results over a wide range of operating conditions. The blackbody emissions were fit to Planck's law in order to estimate flame temperature. The observed temperatures matched well with simulations run via a chemical equilibrium solver [2]. Further validation of the algorithm used to estimate flame temperature was provided by capturing emissions from a tungsten filament. Additionally, numerous species of interest were identified, including igniter materials and combustion products which contributed to erosion of the nozzle. Two other techniques for estimating flame temperature were investigated: (1) the analysis of OH emission bands near 315 nm and (2) the analysis of the sodium D-lines around 590 nm. Trace amounts of sodium chloride were added to the melted paraffin wax prior to spin-casting in order to attempt to produce D-line emissions. Absorption spectroscopy was conducted on paraffin and a blackener dye used to improve the absorption of thermal radiation at the exposed surface of the fuel grain. Absorption spectra were acquired between 0.2 and 22 microns. These absorption spectra were compared with flame emission spectra within the combustion chamber in order to gain insight into how heat is transferred from the flame to the fuel surface.

    Scopus

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  • A Theoretical Study of Combustion Stability in Vortex Injection Hybrid Rocket Engine(共著) 査読有り

    Kohei OZAWA, Toru SHIMADA

    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN   2014年07月

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    担当区分:責任著者   記述言語:英語   掲載種別:研究論文(学術雑誌)

    DOI: 10.2322/tastj.12.Pa_53

    その他リンク: https://www.jstage.jst.go.jp/browse/tastj/-char/en

▼全件表示

著書

  • Hybrid Propulsion Technology Development in Japan for Economic Space Launch 査読有り

    Shimada T., Yuasa S., Nagata H., Aso S., Nakagawa I., Sawada K., Hori K., Kanazaki M., Chiba K., Sakurai T., Morita T., Kitagawa K., Wada Y., Nakata D., Motoe M., Funami Y., Ozawa K., Usuki T.(共著)

    Springer Aerospace Technology  2017年01月 

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    記述言語:英語

    The demand for the economic and dedicated space launchers for vast amount of lightweight, so-called nano-/microsatellites, is now growing rapidly. There is a strong rationale for the usage of the hybrid propulsion for economic space launch as suggested by the assessment conducted here. A typical concept of development of such an economic three-stage launcher, in which clustering unit hybrid rocket engines are employed, is described with a development scenario. Thanks to the benefits of hybrid rocket propulsion, assuring and safe, economic launcher dedicated to lightweight satellites can be developed with a reasonable amount of quality assurance and quality control actions being taken in all aspects of development such as raw material, production, transportation, storage, and operation. By applying a multi-objective optimization technique for such a launch system, examples of possible launch systems are obtained for a typical mission scenario for the launch of lightweight satellites. Furthermore, some important technologies that contribute strongly to economic space launch by hybrid propulsion are described. They are the behavior of fuel regression rate, the swirling-oxidizer-flow-type hybrid rocket, the liquid oxygen vaporization, the multi-section swirl injection, the low-temperature melting point thermoplastic fuel, the thrust and O/F simultaneous control by altering-intensity swirl-oxidizer-flow-type (A-SOFT) hybrid, the numerical simulations of the internal ballistics, and so on.

    DOI: 10.1007/978-3-319-27748-6_22

    Scopus

  • Chemical Rocket Propulsion: A Comprehensive Survey of Energetic Materials

    De Luca, L.T., Shimada, T., Sinditskii, V.P., Calabro, M. (Eds.), Ozawa, K.( 範囲: Part VI Hybrid Rocket Propulsion, Hybrid Propulsion Technology Development in Japan for Economic Space Launch, pp. 545-576.)

    Springer  2016年08月  ( ISBN:978-3319277462

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    記述言語:英語

口頭発表・ポスター発表等

  • Accuracy of real-time fuel regression measurement function of a 3d printed solid fuel

    Ozawa K., Wang H.W., Inenaga T., Tsuboi N.

    AIAA Propulsion and Energy 2020 Forum  American Institute of Aeronautics and Astronautics

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    開催期間: 2020年08月24日 - 2020年08月26日   記述言語:英語  

  • Real-time Regression Rate Measurement of an Additive-manufactured Functional Hybrid Rocket Fuel

    Kohei Ozawa, Han Wei Wang, Takuro Yoshino, Nobuyuki Tsuboi

    70th International Astronautical Congress 

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    開催期間: 2019年10月21日 - 2019年10月25日   記述言語:英語  

  • Investigation of Graphite Nozzle Erosion in Hybrid Rockets Using N2O/HDPE

    Kohei Ozawa

    AIAA Propulsion and Energy 2019 Forum 

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    開催期間: 2019年08月19日 - 2019年08月22日   記述言語:英語  

  • Boundary-Layer Combustion of Wax-based Fuels at Various Chamber Pressures under Two Static Acceleration Environments

    Kohei Ozawa

    AIAA Propulsion and Energy 2019 Forum 

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    開催期間: 2019年08月19日 - 2019年08月22日   記述言語:英語  

  • Simulation of a Detonation Propagation in a Two-phase Gas/liquid Cross Flow Injection

    Nicolas Jourdaine

    27th International Colloquium on the Dynamics of Explosions and Reactive Systems 

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    開催期間: 2019年07月28日 - 2019年08月02日   記述言語:英語  

  • Two-dimensional Numerical Simulations on Unstable Propagation of Propane/Oxygen Detonation Using a Detailed Chemical Mechanism

    Naomi Takeshima

    27th International Colloquium on the Dynamics of Explosions and Reactive Systems 

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    開催期間: 2019年07月28日 - 2019年08月02日   記述言語:英語  

  • Two-Dimensional Numerical Simulation of Flame Acceleration and Deflagration-to-Detonation Transition in Channels with Obstacles: Effects of Blockage Ratio and Chanel Size

    Kanta Iwasaki

    27th International Colloquium on the Dynamics of Explosions and Reactive Systems 

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    開催期間: 2019年07月28日 - 2019年08月02日   記述言語:英語  

  • Numerical Study on Cryogenic Jet/Crossflow Interaction Structures under a Supercritical Pressure

    Taishi AMANO

    International Symposium on Space Technology and Science  The Japan Society for Aeronautical and Space Sciences

     詳細を見る

    開催期間: 2019年06月15日 - 2019年06月21日   記述言語:英語  

  • Feasibility Study of a Real-time Fuel Regression Rate Measurement using an Electrostatic Capacitive Probe for Hybrid Rocket Engines

    Han Weig WANG

    International Symposium on Space Technology and Science  The Japan Society for Aeronautical and Space Sciences

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    開催期間: 2019年06月15日 - 2019年06月21日   記述言語:英語  

  • Visualization of Boundary Layer Combustion of Wax-Based Fuels in Horizontal and Vertical Chamber Configurations

    Kohei OZAWA

    International Symposium on Space Technology and Science  The Japan Society for Aeronautical and Space Sciences

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    開催期間: 2019年06月15日 - 2019年06月21日   記述言語:英語  

  • カーボンブラックを用いたメタン/酸素予混合気におけるデ トネーションセル構造の可視化とギャッロッピングデトネーシ ョンの速度振動の詳細計測

    久門昂平

    日本機械学会九州支部 第72期総会講演会 講演論文集 

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    開催期間: 2019年03月14日   記述言語:日本語  

  • Development of a high efficiency system with a rotating detonation engine for a gas turbine engine (RDE-GTE) using pressure gain combustion

    Koichi Hayashi A., Tsuboi N., Ozawa K., Ishii K., Obara T., Maeda S., Dzieminska E., Mizukaki T.

    AIAA Scitech 2019 Forum 

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    開催期間: 2019年01月01日   記述言語:英語  

  • Three-dimensional numerical simulation of disk rotating detonation engine: Unsteady flow structure

    Watanabe T., Jourdaine N., Ozawa K., Tsuboi N., Kojima T., Hayashi A.

    AIAA Scitech 2019 Forum 

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    開催期間: 2019年01月01日   記述言語:英語  

  • ハイブリットロケットの推力・酸燃比制御に向けた新しい瞬時後 退速度計測手法

    王 瀚緯

    第62回宇宙科学技術連合講演会 

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    開催期間: 2018年10月24日 - 2018年10月26日   記述言語:日本語  

  • 数値解析及び風洞試験によるオービター搭載Waveriderの空力 特性調査

    宇崎友規

    第62回宇宙科学技術連合講演会 

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    開催期間: 2018年10月24日 - 2018年10月26日   記述言語:日本語  

  • 加速度環境におけるハイブリッドロケット用ワックス燃料の境界層燃焼可視化に関する研究

    吉野 拓郎

    第62回宇宙科学技術連合講演会 

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    開催期間: 2018年10月24日 - 2018年10月26日   記述言語:日本語  

  • 軸・接線噴射を用いたハイブリッドロケットエンジンのインジェクタ特性解析

    吉野 拓郎

    第62回宇宙科学技術連合講演会 

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    開催期間: 2018年10月24日 - 2018年10月26日   記述言語:日本語  

  • Visualization of boundary layer combustion of wax-based fuels in vertical and horizontal configurations

    Ozawa K., Yoshino T., Tsuboi N.

    2018 Joint Propulsion Conference 

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    開催期間: 2018年07月09日 - 2018年07月11日   記述言語:英語  

  • 加速度環境におけるワックス燃料の境界層燃焼可視化実験へ向けた着火方法の検討

    吉野 拓郎, 小澤 晃平, 坪井 伸幸

    日本機械学会九州支部講演論文集 

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    開催期間: 2018年01月   記述言語:日本語  

    CiNii Article

  • 強度可変酸化剤流旋回型ハイブリッドロケットにおける酸化剤軸・接線噴射の非定常数値解析

    後藤 祥太, 小澤 晃平, 宇崎 友規, 西川 佳希, 坪井 伸幸

    日本機械学会九州支部講演論文集 

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    開催期間: 2018年01月   記述言語:日本語  

    CiNii Article

  • Monitoring of natural disaster based on Synthetic Aperture Radar (SAR) satellite in Southeast Asia

    Seo S., Nomura S., Ariu K., Yoshihara Y., Kumse K., Funabiki N., Ozawa K., Tanaka E., Yuki T.

    Proceedings of the International Astronautical Congress, IAC 

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    開催期間: 2017年09月25日 - 2017年09月29日   記述言語:英語  

    Southeast Asia is a region that is severely affected by a variety of natural disasters, such as typhoons, heavy rain, earthquakes, and volcanic eruptions. Synthetic-Aperture Radar (SAR) has a large advantage in speed of data acquisition and observation because it can observe in the night or through clouds. SAR has been implemented only on large satellites so far. However, recently, a new concept of SAR and a high-speed downlink system using single small satellites has been developed. This technology can drastically decrease development costs of SAR satellites, increase the frequency of observations using a constellation of SAR satellites, and enable many developing countries to have SAR satellites. Moreover, this technology enables Southeast Asian countries to construct a shared rapid disaster observation system internationally. In this paper, a possible business model from the viewpoints of Japanese manufacturer and governments that utilize a feasible SAR satellite system based on MicroXSAR is described. A cost-benefit analysis is conducted and presented from the perspective of the manufacturer. Besides, the price of the whole observation systems is also calculated using a cost-estimation method of satellites, and is evaluated from the aspects of economic powers of Southeast Asian countries.

  • Effects of O/F shifts on flight performances of vertically launched hybrid sounding rockets

    Ozawa K., Shimada T.

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

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    開催期間: 2017年01月01日   記述言語:英語  

    Impacts of various types of O/F shifts on flight performances of a single stage sounding rocket with a scale of S-520 sounding rocket series were comprehensively evaluated by flight simulations of O/F controlled and uncontrolled hybrid rockets. Before the simulation, sources of O/F shifts and factors affected by O/F shifts were classified and discussed in order to clarify the respective sets of simulations. O/F shifts along with the median fuel regression rate behaviors, systematic errors of the behaviors, and random errors of fuel regression rates were statistically modelled using multiple regression theroy. Shifts of thermodynamic states of productive gases after combustion, shifts of c ∗ efficiency, and nozzle throat erosion were modelled as factors affected by O/F shifts. Operations of propulsion systems were assumed to include throttling. In the presence of O/F shifts along with a median regression rate equation, shifts of thermodynamic states of productive gases were the dominant factor causing performance losses of O/F uncontrolled rockets. In the presence of systematic or random errors of fuel regression rates, residuals of propellants were the other dominant factor to decrease flight performances. Especially in the random error cases, as a result of 3000 times of flight performances, the guaranteed highest altitude of the O/F controlled rocket was 5.2% higher than that of uncontrolled rocket within ±3σ. This result shows that the O/F controlled hybrid rockets have about 2.6% higher acceleration than the O/F uncontrolled rockets. Accuracy of acceleration of the O/F controlled hybrid rocket was 10 times higher than that of the O/F uncontrolled hybrid rocket. These results indicate that the significance of O/F shifts elimination of hybrid rockets from both the aspects of expectancy and accuracy of flight performances.

  • Flame emission spectroscopy in a paraffin-based hybrid rocket

    Stober K., Leccese G., Narsai P., Ozawa K., Cantwell B.

    Proceedings of the International Astronautical Congress, IAC 

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    開催期間: 2016年01月01日   記述言語:英語  

    Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved. A spectrometer was used to measure the emissions emanating from the plume and combustion chamber of a paraffin-based hybrid rocket. Flame emissions were captured between 200-900 nm at numerous points during the 3-7 second ground tests. Time-resolved blackbody emissions were obtained, as well as emission and absorption peaks associated with combustion products, emanating from the plume and combustion chamber. Plume measurements were taken seven inches aft of the nozzle exit plane at the center of the plume. The rocket is described in Narsai et al. [1] and utilizes paraffin and additives as fuel along with gaseous oxygen as the oxidizer. Chamber pressure and oxidizer flow rates were varied in order to validate results over a wide range of operating conditions. The blackbody emissions were fit to Planck's law in order to estimate flame temperature. The observed temperatures matched well with simulations run via a chemical equilibrium solver [2] . Further validation of the algorithm used to estimate flame temperature was provided by capturing emissions from a tungsten filament. Additionally, numerous species of interest were identified, including igniter materials and combustion products which contributed to erosion of the nozzle. Two other techniques for estimating flame temperature were investigated: (1) the analysis of OH emission bands near 315 nm and (2) the analysis of the sodium D-lines around 590 nm. Trace amounts of sodium chloride were added to the melted paraffin wax prior to spin-casting in order to attempt to produce D-line emissions. Absorption spectroscopy was conducted on paraffin and a blackener dye used to improve the absorption of thermal radiation at the exposed surface of the fuel grain. Absorption spectra were acquired between 0.2 and 22 microns. These absorption spectra were compared with flame emission spectra within the combustion chamber in order to gain insight into how heat is transferred from the flame to the fuel surface.

  • Static burning tests on a bread board model of altering-intenisty swirling-oxidizer-flow-type hybrid rocket engine

    Ozawa K., Usuki T., Mishima G., Kitagawa K., Yamashita M., Mizuchi M., Katakami K., Maji Y., Aso S., Tani Y., Wada Y., Shimada T.

    52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016 

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    開催期間: 2016年01月01日   記述言語:英語  

    © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. A bread board model of Altering-number Swirling Flow Type (A-SOFT) hybrid rocket engine has been newly developed and the static burning tests have been conducted. A-SOFT hybrid rocket engines have controllability of both thrust and O/F without any replacement of components in the engine. This advantage is acquired by controlling oxidizer mass flow rate and effective swirl number. The purpose of this set of experiments is to confirm the continuity, monotonousness and predictability of the performances of A-SOFTs. The A-SOFT BBM showed a favorable fuel regression behavior. The fuel regression averaged along spatial direction fit the shape of regression rate function proposed before the experiments within ±3.5% errors, and the fuel regression rates are continuous and monotonous along swirl number and oxidizer mass flux. The O/F and thrust data also respectively fit the prediction formulas within ±4.2% and ±4.7% errors. c* efficiency is evaluated with the Isp efficiency - nozzle efficiency ratio η Isp / η C F in order to compensate the pressure sensing errors caused by the centrifugal forces of swirling flows. Though η Isp / η C F in the cases of weak swirl injection was clearly larger than in the cases of axial injection, its dependence on effective geometric swirl number was not clear in strong swirl conditions.

  • Flight performance simulations of vertical launched sounding rockets using altering-intensity swirling-oxidizer-flow-type hybrid motors

    Ozawa K., Shimada T.

    51st AIAA/SAE/ASEE Joint Propulsion Conference 

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    開催期間: 2015年01月01日   記述言語:英語  

    © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In practical usage of conventional hybrid rocket engines, the oxidizer-to-fuel ratio (O/F) shift occurs by either the fuel port diameter increase or throttling because the fuel regression rate is not proportional to the oxidizer mass flux. As a promising technique to eliminate the O/F shift in a wide throttling range, Altering-intensity-Swirling-Oxidizer-Flow-Type (A-SOFT) hybrid rocket engines are proposed. A-SOFTs control O/F, independently of thrust, with the swirl intensity of oxidizer from the injector, as well as the mass flow rate of the oxidizer. In this paper, the increase rates of engine performance caused by O/F shift eliminating technique are evaluated with a vertical launch simulation for single stage sounding rockets. This simulation includes the throat erosion and c* efficiency models which can be affected by O/F shifts. The statistical uncertainty of fuel regression model is also included to evaluate the robustness of A-SOFTs and SOFTs. The increase rates of total impulse and maximum altitude of A-SOFTs compared to SOFTs depends on maximum oxidizer mass flow rate and are about 2% and 4% respectively. The most effective indicators in this evaluation to the flight performance are residuals of propellants and c* efficiency. Owing to the sensitivity of the flight performances to residuals, the fuel regression errors can cause risks of large losses of the highest altitude in SOFTs, and it is found that the feedback control of A-SOFTs have robustness to the fuel regression errors to some extent. c* efficiency dependent on L* is also sensitive to O/F shifts because O/F shifts affect combustion chamber volume and increase of throat area.

  • Performance calculations and burning tests on altering-intensity swirling oxidizer flow type hybrid rocket engines

    Ozawa K., Kitagawa K., Shimada T.

    Proceedings of the International Astronautical Congress, IAC 

     詳細を見る

    開催期間: 2015年01月01日   記述言語:英語  

    As a technique to eliminate O/F shifts, Altering-intensity Swirling Oxidizer Flow Type (A-SOFT) hybrid rocket is proposed using the combination axial and tangential oxidizer injection. In this paper, the possible geometric design point of the motor is considered about 2-tons class single stage A-SOFT hybrid rockets, and the nominal performance increase by eliminating O/F shifts is evaluated. The highest altitude of the A-SOFT is 450 [km] and the averaged ISP is 284[s] with 100[kg] payload, and this performance is several percent higher than the ones of S-520 with 95[kg] payload, Japanese 2-tons class solid sounding rocket. Though A-SOFTs increase the nominal flight performance by 1% from Swirling Oxidizer Flow Type hybrid rockets, which cannot O/F shift, their feedback control prevents the large performance losses caused by fuel regression errors and residuals, and this function is not found in the conventional hybrid rockets. In this paper, a planning of steady state burning tests of A-SOFT bread board model is also explained. The purpose of the burning tests are to demonstrate the concept of A-SOFTs. The test motor is 250[N] scale and uses polypropylene and GOX and its burning duration is 5[s] . In more than 10 times steady state tests, 50% throttling conditions with effective geometric swirl intensity between 0 and 37.3 are inculded.

  • Linear stability analysis of uni-directional vortex injection hybrid rocket engines

    Ozawa K., Shimada T.

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

     詳細を見る

    開催期間: 2014年01月01日   記述言語:英語  

    © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. In this paper, first, a theoretical prediction method of regression rates and heat flux to solid fuels of uni-directional vortex injection hybrid rocket engines is developed by introducing a new swirl intensity decline model toward axial direction. Next, a linear propagative relation of heat flux to solid fuels with disturbances of oxidizer mass flux, fuel regression, and initial swirl intensity is derived. The couple of this response model and another unteady response model of solid fuel gasification amplifies oxidizer mass flux disturbance in the form of regression rate oscillation. This is the basic mechanism of low frequency instability unique to hybrid rocket engines. The linear stability analysis for uni-directional vortex types simulating both ILFI amplification source and main stream model is conducted. The result of this analysis shows that uni-directional vortex injection hybrid rocket engines have the same linear unstable mode as axial hybrid rocket engines.

▼全件表示

工業所有権

  • 燃焼装置及びその製造方法、並びに、ハイブリッドロケットエンジン

    小澤晃平, オウ カンイ, 吉野拓郎

     詳細を見る

    出願番号:2019-190285  出願日:2019年10月17日

講演

  • Combustion Flow of Wax-based Hybrid Rockets in Various Acceleration Environments

    JST さくらサイエンスプラン 2019 年度 熱流体工学分野に関する国際共同セミナー  2019年08月  九州工業大学

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    講演種別:基調講演   開催地:九州工業大学戸畑キャンパス  

報道関係

  • CLOSE-UP JKA   新聞・雑誌

    小澤晃平

    西日本新聞  2021年08月06日

学術関係受賞

  • 火薬学会奨励賞

    火薬学会   ハイブリッドロケット用多素材3Dプリント燃料についての研究   2022年05月24日

    小澤晃平

     詳細を見る

    受賞国:日本国

  • JAICI賞

    一般社団法人化学情報協会   2022年05月24日

    小澤晃平

     詳細を見る

    受賞国:日本国

  • AIAA Hybrid Rocket Best Student Paper

    AIAA   2016年07月27日

    Kohei Ozawa, Toru Shimada

     詳細を見る

    受賞国:アメリカ合衆国

  • Japanese Rocket Society Award

    International Symposium on Space Technology and Science   2013年06月09日

    Kohei Ozawa

     詳細を見る

    受賞国:日本国

科研費獲得実績

  • 液化ガス燃料を用いた革新的爆轟推進の研究

    研究課題番号:24K01082  2024年04月 - 2028年03月   基盤研究(B)

  • 不均一気体爆轟伝播の核心:垂直衝撃波-異成分噴流間相互干渉の解明

    研究課題番号:23KK0083  2023年10月 - 2027年03月   国際共同研究加速基金・国際共同研究強化(B)

  • 多素材3Dプリンタ製機能性固体燃料によるハイブリッドロケットの閉ループ制御

    研究課題番号:21K14346  2021年04月 - 2024年03月   若手研究

  • 低粘性液化燃料を生ずる境界層燃焼の加速度環境下における燃焼機構解明

    研究課題番号:18K13926  2018年04月 - 2022年03月   若手研究

  • ハイブリッドロケットの酸化剤旋回流を用いた最適混合比と推力の同時制御

    研究課題番号:15J08028  2015年04月 - 2017年03月   特別研究員奨励費

受託研究・共同研究実施実績

  • パルスデトネーションエンジンを応用した姿勢制御スラスタの研究開発

    2018年09月 - 現在

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    研究区分:受託研究

    目的: デトネーションを応用し、小型ロケットに搭載する無毒・低コスト・高効率な姿勢制御スラスタを研究開発する準備段階として、エンジンの研究開発ノウハウを獲得する。また、エンジニアリングモデル開発に向けた課題を洗い出す。

    研究内容:
    小型のパルスデトネーションエンジンを試作し大気中で燃焼実験を行い、所望の性能が出るか検証する。

その他競争的資金獲得実績

  • 円盤型回転爆轟エンジンの爆轟波特性決定機構の解明

    2022年02月 - 2024年02月

    日本学術振興会 海外特別研究員 滞在費・研究活動費  

  • 低粘性液化燃料を生ずる境界層燃焼の加速度環境下における燃焼機構解明

    2021年04月 - 2022年03月

    令和 3 年度 Linear Hyper G 公募共同研究  

  • 3Dプリンタ製機能性固体燃料を用いたハイブリッドロケットの閉ループ推力制御高度化実証

    2020年04月 - 2021年03月

    機械振興補助事業 若手研究  

  • 燃料流量計測機能を持つ固体燃料の高性能化に向けたグラフェン混合導電性樹脂の研究

    2020年04月 - 2021年03月

    公益財団法人 火薬工業技術奨励会 研究助成金  

  • 低粘性液化燃料を生ずる境界層燃焼の加速度環境下における燃焼機構解明

    2020年04月 - 2021年03月

    令和 2 年度 Linear Hyper G 公募共同研究  

  • 3Dプリンタ製機能性固体燃料を用いた超小型衛星用スラスタの高度化

    2019年10月 - 2021年03月

    公益財団法人服部報公会 工学研究奨励援助金  

  • 2通りの静的加速度環境における多様な燃焼圧での黒色ワックス燃料の境界層燃焼特性

    2019年08月

    公益財団法人宇宙科学振興会 国際学会出席旅費支援  

  • 爆轟波を用いた環境性に優れたロケット上段用姿勢制御スラスタの研究

    2019年04月 - 2020年03月

    平成31年度新成長戦略推進研究開発事業(シーズ創出・実用性検証事業)  

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    デトネーションを応用し、小型ロケットに搭載する無毒・低コスト・高効率な姿勢制御スラスタを研究開発する準備段階として、エンジンの研究開発ノウハウを獲得する。また、エンジニアリングモデル開発に向けた課題を洗い出す。

  • 3次元プリンタを用いたリアルタイム性能の自己評価機能を持つハイブリッドロケットの研究

    2019年04月 - 2020年03月

    公益財団法人 火薬工業技術奨励会 研究助成金  

▼全件表示

海外研究歴

  • 円盤型回転爆轟エンジンの爆轟波特性決定機構の解明

    ミシガン大学アナーバー校  アメリカ合衆国  研究期間:  2022年02月28日 - 現在

  • 小型ハイブリッドロケットエンジン内部燃焼流のパラメトリック研究および可視化

    スタンフォード大学  アメリカ合衆国  研究期間:  2016年07月03日 - 2016年10月02日

担当授業科目(学内)

  • 2023年度   機械工学実験Ⅰ

  • 2023年度   機械工学実験Ⅰ

  • 2023年度   機械工学実験Ⅰ

  • 2019年度   宇宙システム工学入門

  • 2017年度   宇宙工学入門

  • 2017年度   機械工学実験Ⅰ

  • 2017年度   機械工学実験Ⅱ

▼全件表示

教育活動に関する受賞・指導学生の受賞など

  • ミスミ学生ものづくり支援

    株式会社ミスミグループ本社  

    2018年12月18日

    吉野拓郎

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    学生のものづくり・研究に対する支援

学会・委員会等活動

  • 日本航空宇宙学会西部支部   常任幹事  

    2020年03月 - 2021年02月

社会貢献活動(講演会・出前講義等)

  • NHK高専ロボコン 九州沖縄地区大会 主審

    2017年10月20日 - 2017年10月21日

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    種別:その他