2021/07/12 更新

写真a

コウチ トシノリ
河内 俊憲
KOUCHI Toshinori
所属
自然科学学域 教授
職名
教授
外部リンク

学位

  • 博士(工学) ( 2005年3月   東北大学 )

研究キーワード

  • 航空宇宙工学

  • 超音速流

研究分野

  • ものづくり技術(機械・電気電子・化学工学) / 熱工学

  • フロンティア(航空・船舶) / 航空宇宙工学

  • ものづくり技術(機械・電気電子・化学工学) / 流体工学

学歴

  • 東北大学    

    - 2005年

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    国名: 日本国

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  • 東北大学   Graduate School, Division of Engineering  

    - 2005年

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  • 東北大学   Graduate School, Division of Engineering  

    - 2003年

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  • 東北大学    

    - 2003年

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    国名: 日本国

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経歴

  • 東北大学工学部   Faculty of Engineering

    2007年

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  • 国立研究開発法人宇宙航空研究開発機構

    2005年 - 2007年

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  • 独立行政法人 宇宙航空研究開発機構 プロジェクト研究員

    2005年 - 2007年

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  • 東北大学 日本学術振興会 特別研究員

    2003年 - 2005年

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  • 東北大学 大学院工学研究科 航空宇宙工学専攻 航空宇宙システム工学講座   助教

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所属学協会

 

論文

  • Transverse jet mixing in a supersonic grid turbulence

    Toshinori Kouchi, Masaki Iwachido, Takahiro Nakagawa, Yasunori Nagata, Shinichiro Yanase

    AIAA Scitech 2020 Forum   1 PartF   2020年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    © 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Turbulence grids were applied to Mach 2 supersonic wind tunnel to increase turbulence in a mainstream. We measured wall pressure, velocity by Laser Speckle Velocimetry (LSV) using acetone condensation nanoparticle and density by acetone Planer Laser Induced Fluorescence (PLIF). In this study, the test section has 12 mm width and 10 mm height at the exit of the nozzle and the turbulence grids, which consisted of tungsten wires having sub-mm diameter, was installed at the exit of the nozzle. Combination of the wire grid and tunnel wall expansion increased mainstream turbulence without flow unstart in the test section flow. In the case with 0.4-mm-diameter grid at 3 by 3 arrangement having a blockage of 21% of the nozzle exit area, the mainstream turbulence reached 8% of the mainstream velocity. Nitrogen gas was perpendicularly injected into the grid-generated supersonic turbulent flow and it mixing performance was investigated by using acetone-PLIF. Installation of the wire grid affected not only mainstream turbulence but also wall-bounded flow, resulting in thickening boundary layer. As a result, jet penetration increased with installing the wire grid. However, no remarkable improvement of the jet mixing was observed with installing the wire grid.

    DOI: 10.2514/6.2020-2040

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  • Les analysis of transverse jet mixing in supersonic free-stream turbulence

    Yasunori Nagata, Seitaro Yokoi, Rei Aoki, Toshinori Kouchi, Shinichiro Yanase

    AIAA Scitech 2020 Forum   1 PartF   1 - 11   2020年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    © 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In the engine combustion test using the hypersonic wind tunnel facility, high temperature, high pressure, and high velocity flow is generated by combustion heating upstream of the nozzle. However, in this process, the turbulence of the flow and the contamination by the combustion gas occur unavoidably. Therefore, it is a concern that the different situation is observed between the ground test and actual flight because of these unavoidable factors. In this study, we focus on the mixing process of the injected fuel and the free-stream and aim to numerically evaluate the influence of the free-stream turbulence on the jet mixing. The LES calculation with free-stream turbulence is performed by the inflow boundary condition applying the velocity, temperature, and pressure fluctuations based on the Random Flow Generation (RFG) method, strong Reynolds analogy (SRA), and an isentropic process. The input parameter for RFG can control the generated free-stream turbulence characteristic inside the calculation domain. The difference between the cases with and without free-stream turbulence, whose intensity is up to 2.2% of mean free-stream velocity, is not observed clearly on the mean and instantaneous value fields in the present calculation.

    DOI: 10.2514/6.2020-1335

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  • Hydrodynamic Instability with Convective Heat Transfer through a Curved Channel with Strong Rotational Speed

    Mohammad Sanjeed Hasan, Rabindra Nath Mondal, Toshinori Kouchi, Shinichiro Yanase

    8TH BSME INTERNATIONAL CONFERENCE ON THERMAL ENGINEERING   2121   2019年

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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)   出版者・発行元:AMER INST PHYSICS  

    In this paper, a comprehensive numerical study on viscous incompressible fluid flow and heat transfer through a loosely coiled square duct has been presented. Spectral method is used as a basic tool to solve the system of non-linear partial differential equations. Numerical calculations are carried out for the Dean number Dn = 1000 with a temperature difference across the vertical sidewalls for the Grashof number Gr = 100, where the outer wall is heated and the inner wall cooled. A rotation of the duct about the center of curvature is imposed in the positive direction for the Taylor number 0 <= Tr <= 2000 and combined effects of the centrifugal, Coriolis and buoyancy forces are investigated. First, steady solutions are obtained by the Newton-Raphson iteration method. As a result, three branches of asymmetric steady solutions with two- to four-vortex solutions are obtained. Then, time evolution calculations as well as power spectrum density of the unsteady solutions are obtained and it is found that the unsteady flow undergoes through various flow instabilities, if Tr is increased in the positive direction. Nusselt numbers are calculated as an index of convective heat transfer, and it is found that convective heat transfer is significantly enhanced by the secondary flow. Finally, a comparison between the numerical and experimental investigations has been provided. It is found that there is a good agreement between the numerical and experimental investigations.

    DOI: 10.1063/1.5115851

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  • Numerical study of air-entraining and submerged vortices in a pump sump

    Shinichiro Yanase, Ryo Yamasaki, Toshinori Kouchi, Shunsuke Hosoda, Yasunori Nagata, Higuchi Shunji, Toshihiko Kawabe, Toshihiro Takami

    29TH IAHR SYMPOSIUM ON HYDRAULIC MACHINERY AND SYSTEMS   240 ( 3 )   2019年

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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)   出版者・発行元:IOP PUBLISHING LTD  

    Numerical detection of harmful vortices in pump sumps, such as an air-entraining vortex (AEV) and a submerged vortex (SMV), is crucially important to develop the drain pump machinery. We performed numerical simulations of the benchmark experiments of the pump sump conducted by Matsui et al. (2006 and 2016) using the OpenFOAM and compared the simulation results with the experimental data considering the effects of turbulence model, grid density and detection method of the vortices. We studied the threshold of the gas-liquid volume fraction of the VOF method and the second invariant of velocity gradient tensor to identify AEV and SMV. The methods proposed in the present paper were found to be very effective for the detection of the vortices, and the simulation results by RANS with the SST k-omega model successfully reproduced the experimental data. LES with the Smagorinsky model, however, was sensitive to the grid system and difficult to reproduce the experimental data even for the finest grid system having 3.7 million cells in the present study.

    DOI: 10.1088/1755-1315/240/3/032001

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  • Acetone-condensation nano-particle image velocimetry in a supersonic boundary layer

    Toshinori Kouchi, Seiya Fukuda, Syouma Miyai, Yasunori Nagata, Shinichiro Yanase

    AIAA Scitech 2019 Forum   2019年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    © 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. Condensation nanoparticle planar laser light scattering imaging was conducted in a suction type Mach 1.9 supersonic flow. A series of the nanoparticle image pair was obtained by using the same optical components of PIV. The velocity data were obtained by the image pairs using the image-based correlation procedure as same as PIV. Acetone vapor instead of tracer particles in a conventional PIV was added to a mainstream gas in a reservoir. Acceleration process through a Laval nozzle automatically generated condensation particles of the acetone which were uniformly seeded into the entire flowfield. The increase in the additive concentration increased the number density of the particle and enabled a detail visualization of the vortex structures in the boundary layer. The increase in the additive concentration also increased the mean molecular weight of the acetone-seeded air. This decreased the flow speed. However, this is not a big matter because the heat release due to the condensation was negligible and the decrease in the flow speed was easily predicted from the thermodynamic properties of the gas. The particle size was difficult to be measured directly, so the tracer response time was estimated by the oblique shock test. The Stokes diameter of the particle was estimated to be 160 nm. Such a small diameter particle provides high traceability resulting in capturing shock wave with a few vector spacing and also provides high spatial resolution image resulting in capturing sub-mm scale vortices within the boundary layer. The mean and turbulent velocity fields evaluated from such high spatially and temporally resolved image pairs fairly agreed with previous measurement in the boundary layer.

    DOI: 10.2514/6.2019-1821

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  • Laminar forced convective heat transfer in helical pipe flow

    Anup Kumer Datta, Shinichiro Yanase, Toshinori Kouchi, Mohammed M. E. Shatat

    INTERNATIONAL JOURNAL OF THERMAL SCIENCES   120   41 - 49   2017年10月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:ELSEVIER FRANCE-EDITIONS SCIENTIFIQUES MEDICALES ELSEVIER  

    Laminar forced convective heat transfer in a helical pipe with circular cross section subjected to wall heating was investigated numerically by three-dimensional (3D) direct numerical simulations (DNS) comparing with the experimental data obtained by Shatat (2010). The study was performed for three Prandtl numbers, Pr = 8.5, 7.5 and 4.02, over the wide range of torsion. In 3D steady calculations, we found the appearance of fully-developed axially invariant flow regions, where the averaged Nusselt number (averaged over the peripheral of the pipe cross section) were calculated, being in good agreement with the experimental data. Because of the effect of torsion on the heat transfer characteristics, the averaged Nusselt number exhibits repetition of decrease and increases as torsion increases from zero for all Reynolds numbers. It was found that there exists two maximums and two minimums of the averaged Nusselt number. It is interesting that the global minimum of the Nusselt number occurs at beta congruent to 0.1 and the global maximum at beta congruent to 0.55. (C) 2017 Elsevier Masson SAS. All rights reserved.

    DOI: 10.1016/j.ijthermalsci.2017.05.026

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  • Numerical Study of Turbulent Helical Pipe Flow With Comparison to the Experimental Results

    Anup Kumer Datta, Yasutaka Hayamizu, Toshinori Kouchi, Yasunori Nagata, Kyoji Yamamoto, Shinichiro Yanase

    JOURNAL OF FLUIDS ENGINEERING-TRANSACTIONS OF THE ASME   139 ( 9 )   2017年9月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:ASME  

    Turbulent flow through helical pipes with circular cross section is numerically investigated comparing with the experimental results obtained by our team. Numerical calculations are carried out for two helical circular pipes having different pitches and the same nondimensional curvature delta (=0.1) over a wide range of the Reynolds number from 3000 to 21,000 for torsion parameter beta (=torsion /root 2 delta = 0.02 and 0.45). We numerically obtained the secondary flow, the axial flow and the intensity of the turbulent kinetic energy by use of three turbulence models incorporated in OpenFOAM. We found that the change to fully developed turbulence is identified by comparing experimental data with the results of numerical simulations using turbulence models. We also found that renormalization group (RNG) k-epsilon turbulence model can predict excellently the fully developed turbulent flow with comparison to the experimental data. It is found that the momentum transfer due to turbulence dominates the secondary flow pattern of the turbulent helical pipe flow. It is interesting that torsion effect is more remarkable for turbulent flows than laminar flows.

    DOI: 10.1115/1.4036477

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  • Extracting dominant turbulent structures in supersonic flow using two-dimensional Fourier transform 査読

    Toshinori Kouchi, Goro Masuya, Shinichiro Yanase

    EXPERIMENTS IN FLUIDS   58 ( 8 )   2017年8月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:SPRINGER  

    A new image process for quantifying both convection velocities (U-C) and scales (lambda(d)) of turbulent structures captured in a fast-framing schlieren movie is presented. We obtained 90 time-series schlieren images of a transverse jet into a Mach 2 supersonic flow with 1-MHz sampling. The schlieren images captured not only the shock and expansion waves but also the turbulent structures within the jet and the boundary layer. The image intensities were extracted along the outer edges of the jet and the boundary layer and were remapped as a time-space intensity map. The time-space map exhibited swept stripe patterns, indicating that stable turbulent structures were periodically generated and convected downstream. The angle and interval of the stripe patterns were efficiently extracted using the two-dimensional Fourier transform, which corresponded to U-C and lambda(d) of the dominant structures. The zero-padding fast Fourier transform and the sub-pixel estimation of the spectral peak positions in the Fourier domain improved the accuracy for evaluating the angle and interval of the stripes, which resulted in the accurate evaluation of U-C and lambda d. The proposed method was validated by comparing U-C obtained using the proposed method to those obtained via schlieren image velocimetry for both the transverse jet and the supersonic boundary layer.

    DOI: 10.1007/s00348-017-2377-z

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  • Effect of Torsion on the Friction Factor of Helical Pipe Flow

    Anup Kumer Datta, Shinichiro Yanase, Yasutaka Hayamizu, Toshinori Kouchi, Yasunori Nagata, Kyoji Yamamoto

    JOURNAL OF THE PHYSICAL SOCIETY OF JAPAN   86 ( 6 )   2017年6月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:PHYSICAL SOC JAPAN  

    Three-dimensional direct numerical simulations of a viscous incompressible fluid flow through a helical pipe with a circular cross section were conducted for three Reynolds numbers, Re (= 80, 300, and 1000), and two nondimensional curvatures, delta (= 0.1 and 0.05), over a wide range of torsion parameters, beta (= nondimensional torsion= /root 2 delta), from 0.02 to 2.8. Well-developed axially invariant regions were obtained where the friction factors were calculated, in good agreement with the experimental data obtained by Yamamoto et al. [Fluid Dyn. Res. 16, 237 (1995)]. It was found that the friction factor sharply increases as beta increases from zero, then decreases after taking a maximum, and finally slowly approaches that of a straight pipe when beta tends to infinity. It is interesting that a peak of the friction factor exists in the region 0.2 <= beta <= 0.3 for all the Reynolds numbers and curvatures studied in the present paper, which manifests the importance of the torsion parameter in helical pipe flow.

    DOI: 10.7566/JPSJ.86.064403

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  • Supersonic Combustion using Multiple Stinger-shaped Injectors 査読

    Toshinori Kouchi, Sadatake Tomioka, Kohshi Hirano, Akiko Matsuo, Goro Masuya

    TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES   60 ( 1 )   56 - 59   2017年

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:JAPAN SOC AERONAUT SPACE SCI  

    Four stinger injectors arrayed at s/D -6:7 were tested in a Mach 2.44 supersonic combustor with a 1.66° diverging section at T0 = 2060 K. Wall pressure and gas sampling measurements confirmed the ability of the multiple stinger injectors to improve combustor performance. Wall pressure and gas sampling measurements confirmed the ability of the multiple stinger injectors to improve combustor performance. For ψ̎0:3, the combustion region was limited in the diverging section of the combustor, so the combustor was operated in the scramjet mode. For this mode, compared with the airflow between the multiple circular injectors, the airflow much more easily passed between the djacent stinger injectors due to the low blockage effect. As a result, the pressure thrust in the stinger case was 5% higher than that in the circular case. The effects of the injector port shape became unclear when the precombustion shock waves appeared far upstream of the injectors (dual-mode scramjet operation). The fuels from the multiple injectors were uniformly distributed in the span-wise direction, and combustor performance was mainly dominated by the change in airflow momentum flux and disturbance due to the precombustion shock wave.

    DOI: 10.2322/tjsass.60.56

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  • Wavelet analysis of transonic buffet on a two-dimensional airfoil with vortex generators

    Toshinori Kouchi, Shingo Yamaguchi, Shunske Koike, Tsutomu Nakajima, Mamoru Sato, Hiroshi Kanda, Shinichiro Yanase

    EXPERIMENTS IN FLUIDS   57 ( 11 )   2016年11月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:SPRINGER  

    We visualized the shock buffets on a two-dimensional transonic airfoil with and without vortex generators (VGs) by using a fast-framing focusing schlieren imaging. The focusing schlieren visualization showed that the flow three-dimensionality around the airfoil became remarkable with installing the VGs. This implies that narrow depth of focus of imaging systems was a key to accurately capture the characteristics of the shock oscillation due to the buffet for the cases with VGs. The time-resolved imaging also revealed that non-periodic components were included in the shock oscillation due to the buffet for the cases with VGs. This prevented Fourier analysis from being applied. We used wavelet analysis to extract the characteristic of the shock oscillation for the cases with VGs. The wavelet spectrograms revealed that the low-frequency oscillation having the buffet frequency was still included intermittently in the shock oscillation even when VG controlled the buffet. The rate of appearing the low-frequency oscillation increased with increasing both the interval between VGs and the angle of attack.

    DOI: 10.1007/s00348-016-2261-2

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  • Development of DEM-CFD Simulation of Combustion Flow in Incinerator with the Representative Particle Model

    Kenya Kuwagi, Toshihiro Takami, Azri Bin Alias, Degang Rong, Hiroshi Takeda, Shinichiro Yanase, Toshinori Kouchi, Toru Hyakutake, Kaoru Yokoyama, Yoshiyuki Ohara, Nobuo Takahashi, Noritake Sugitsue

    JOURNAL OF CHEMICAL ENGINEERING OF JAPAN   49 ( 5 )   425 - 434   2016年5月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:SOC CHEMICAL ENG JAPAN  

    A simulation code based on the discrete element method (DEM) and computational fluid dynamics (CFD) coupling model was developed to simulate the behavior of radioactive cesium in waste incinerators. The waste lump was represented by particles in the simulation. The energy equation for a mixed gas, diffusion equation for each gas component, as well as the energy, drying, pyrolysis, and combustion equations for each particle were solved in the simulation by adding a combustion model to the standard DEM-CFD coupling model. The particle size of the waste changed as drying, pyrolysis, and combustion progressed. At the end of the combustion process, particle waste became ash, and the number of ash particles was enormous. To avoid an excessive computational load due to the high particle number, a similar assembly model was adopted to reduce the particle number in the calculation. There was a good agreement between the simulation and experimental results for the temperature at the outlet of the furnace and the flue gas composition.

    DOI: 10.1252/jcej.15we099

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  • Turbulent diffusion flux of transverse jet into pseudo-shock wave

    Taekjin Lee, Toshinori Kouchi, Yoshinori Oka, Goro Masuya

    54th AIAA Aerospace Sciences Meeting   0   2016年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved. An experiment to investigate the effect of pseudo-shock wave (PSW) on turbulent diffusion of a transverse jet in a Mach 2.0 supersonic crossflow was conducted. The vertical jet from a 2.5 mm-diameter orifice was injected into pseudo-shock wave and jetto-crossflow momentum flux ratio was 2.2. To measure turbulent diffusion flux, stereoscopic particle image velocimetry (SPIV) and acetone planar laser-induced fluorescence (PLIF) system were combined and velocity and concentration fields were simultaneously measured. Measurement plane was a cross section at 10 mm downstream from the injection port. The position of PSW was controlled by a flow plug at the duct exit driven by a stepping motor. The front of PSW was set at 30 mm upstream from the orifice. Without PSW, turbulent diffusion flux (TDF) of the injectant was extended upward and the shape of injectant plume was deformed by a counter-rotating vortex pair (CVP) which was formed behind the injected gas. Under PSW, the CVP made horseshoe-shape concentration distribution. TDF directed from spur of concentration to both free stream side and lee side of the jet. Most of no-PSW condition and outward of under-PSW TDF was concentration-fluctuation-dominant, only inward TDF under PSW emerged as turbulent-dominant. The effect of PSW that enhances diffusion of injected gas was confirmed.

    DOI: 10.2514/6.2016-0662

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  • Wavelet analysis of unsteady shock-wave motion on two-dimensional airfoil with vortex generators

    Toshinori Kouchi, Shingo Yamaguchi, Shinichiro Yanase, Shunsuke Koike, Tsutomu Nakajima, Mamoru Sato, Hiroshi Kanda

    54th AIAA Aerospace Sciences Meeting   0   2016年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA, All rights reserved. We visualized the shock wave-boundary layer interaction on a two-dimensional transonic airfoil with and without vortex generators (VGs) by using a fast-framing focusing schlieren imaging. Image processing extracted a time-space trajectory of the shock wave motion from a time-series of the schlieren images. The shock trajectories were analyzed by using the Morlet wavelet. The shock motions in the cases without VGs were quite periodic. The Fourier analysis well extracted the characteristics of the shock motion such as frequency etc. However, the shock motions in the cases with VGs were not periodic. Therefore, the Fourier analysis is not applicable for this case. The wavelet analysis with the statistical significance test gave quantitative measure of change in the shock oscillation around the buffet frequency due to installation of VGs, such as amplitude and intermittency. The wavelet spectrograms revealed that the installation of VGs did not prevent from generating the buffet but just reduced the amplitude of the buffet oscillation.

    DOI: 10.2514/6.2016-1766

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  • Focusing-schlieren visualization in a dual-mode scramjet

    Toshinori Kouchi, Christopher P. Goyne, Robert D. Rockwell, James C. McDaniel

    EXPERIMENTS IN FLUIDS   56 ( 12 )   1 - 14   2015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)   出版者・発行元:SPRINGER  

    Schlieren imaging is particularly suited to measuring density gradients in compressible flowfields and can be used to capture shock waves and expansion fans, as well as the turbulent structures of mixing and wake flows. Conventional schlieren imaging, however, has difficulty clearly capturing such structures in long-duration supersonic combustion test facilities. This is because the severe flow temperatures locally change the refractive index of the window glass that is being used to provide optical access. On the other hand, focusing-schlieren imaging presents the potential of reduced sensitivity to thermal distortion of the windows and to clearly capture the flow structures even during a combustion test. This reduced sensitivity is due the technique's ability to achieve a narrow depth of focus. As part of this study, a focusing-schlieren system was developed with a depth of focus near +/- 5 mm and was applied to a direct-connect, continuous-flow type, supersonic combustion test facility with a stagnation temperature near 1200 K. The present system was used to successfully visualize the flowfield inside a dual-mode scramjet. The imaging system captured combustion-induced volumetric expansion of the fuel jet and an anchored bifurcated shock wave at the trailing edge of the ramp fuel injector. This is the first time successful focusing-schlieren measurements have been reported for a dual-mode scramjet.

    DOI: 10.1007/s00348-015-2081-9

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  • Collaborative Experimental and Computational Study of a Dual-Mode Scramjet Combustor

    Robert D. Rockwell, Christopher R. Goyne, Brian E. Rice, Toshinori Kouchi, James C. McDaniel, Jack R. Edwards

    JOURNAL OF PROPULSION AND POWER   30 ( 3 )   530 - 538   2014年5月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    Advanced computational models of hypersonic air-breathing combustion processes are being developed to better understand and predict the complex flows within a dual-mode scramjet combustor. However, the accuracy of these models can only be quantified through comparison to experimental databases. Moreover, the quality of computational results is dependent on accurate and detailed knowledge of the combustor inflow and boundary conditions. Toward these ends, this paper describes results from a collaboration of experimental and computational investigators. Detailed computational fluid dynamics and finite element analyses were performed throughout the design and implementation of experiments involving a direct-connect scramjet combustor operating at steady state during long duration testing. The test section hardware was designed to provide substantial access for optical laser diagnostics. Measurement locations included the inflow plane and several locations downstream of fuel injection. A suite of advanced in-stream diagnostics were applied, many of which are described in companion papers. Significant results in this paper include measured static wall pressures and temperatures, stereoscopic particle image velocimetry, and focused schlieren imaging. Validated thermal finite element calculations in the scramjet hardware and temperature maps of the flow path boundaries are also presented. Comparison of experimental results with computational fluid dynamics predictions are discussed in a separate paper.

    DOI: 10.2514/1.B35021

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  • Characteristics of Hydrogen Jets in Supersonic Crossflow: Large-Eddy Simulation Study 査読

    Junya Watanabe, Toshinori Kouchi, Kenichi Takita, Goro Masuya

    JOURNAL OF PROPULSION AND POWER   29 ( 3 )   661 - 674   2013年5月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    The characteristics of hydrogen jets transversely injected into a supersonic crossflow under four different injection and crossflow conditions were investigated by large-eddy simulation. The effects of the jet-to-crossflow momentum flux ratio and crossflow velocity were studied. The jet trajectory in the averaged field was controlled by the value of the square root of the jet-to-crossflow momentum flux ratio irrespective of the crossflow conditions. When the crossflow conditions were fixed many jet characteristics were similar in the space normalized using the square root of jet-to-crossflow momentum flux ratio. On the other hand, the crossflow conditions had a strong impact on the jet characteristics. Although the turbulent intensity around the jet was not affected to a great extent, the shape and convection velocity of the large-scale structures appearing on the windward side of the jet plume depended on the crossflow conditions. With a higher crossflow velocity the convection velocity was higher, the jet was more intermittent, and the progress of mixing was slower. The turbulent mixing state had different features depending on the crossflow conditions.

    DOI: 10.2514/1.B34521

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  • Supersonic Combustion Using a Stinger-Shaped Fuel Injector 査読

    Toshinori Kouchi, Goro Masuya, Kohshi Hirano, Akiko Matsuo, Sadatake Tomioka

    JOURNAL OF PROPULSION AND POWER   29 ( 3 )   639 - 647   2013年5月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    The authors developed a stinger-shaped injector (stinger injector) for supersonic combustors in cold-flow experiments. The stinger injector has a port geometry with a sharp leading edge in front of a streamwise slit. This injector produced higher jet penetration at a lower jet-to-crossflow momentum flux ratio J than a conventional circular injector. We applied the injector in a Mach 2.44 combustion test at a stagnation temperature of 2060 K. At a low fuel-equivalence ratio Phi regime (i.e., low J regime), the injector produced 10% higher pressure thrust than the circular injector because of high jet penetration as expected from the cold-flow experiments. Even at a moderate Phi regime, the stinger injector produced higher pressure thrust than the circular injector. At moderate Phi, the stinger injector held the flame around the injector and generated a precombustion shock wave in front of the injector. The presence of the precombustion shock wave decreased the momentum flux of the crossflow air and diminished the advantage of the injector for jet penetration. The injector, however, produced higher pressure thrust because better flame-holding produced higher pressure around the injector. At a higher Phi regime, the precombustion shock wave went upstream with both injectors. The far-upstream presence of a precombustion shock wave increased the turbulence in the crossflow and spread the fuel from both injectors. Thus, the difference in injector shape was insignificant for thrust performance.

    DOI: 10.2514/1.B34524

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  • Large-Eddy Simulation of Jet in Supersonic Crossflow with Different Injectant Species 査読

    Junya Watanabe, Toshinori Kouchi, Kenichi Takita, Goro Masuya

    AIAA JOURNAL   50 ( 12 )   2765 - 2778   2012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)   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    The effects of injectant species on the turbulent structure and mixing state of jets in supersonic crossflow were investigated using a large-eddy simulation. Hydrogen, helium, nitrogen, and ethylene were transversely injected into a Mach 1.9 airflow at a constant jet-to-crossflow momentum flux ratio. The time-averaged distribution of jet concentration was roughly the same for all the injectant species, but the large-scale structure of scalar fluctuation differed significantly. The probability density functions of injectant mass fraction revealed that the turbulent behavior of nitrogen and ethylene jets was highly intermittent. The velocity field was considerably different between the injectants, owing to the different injection velocities. The hydrogen and helium jets had a much higher velocity difference between jet and crossflow in the near field; thus, the observed turbulent intensities for these two injectants were much higher than those for nitrogen and ethylene. In addition, the spectral analysis of velocity fluctuations in the windward mixing layer showed that the scale of energetic eddies was larger in the hydrogen and helium jets than in the nitrogen and ethylene jets. These characteristics resulted in better mixing in the hydrogen jet than in the ethylene jet for the studied injection conditions.

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  • Flowfield characteristics of a transverse jet into supersonic flow with pseudo-shock wave 査読

    H. Yamauchi, B. Choi, K. Takae, T. Kouchi, G. Masuya

    Shock Waves   22 ( 6 )   533 - 545   2012年11月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)  

    We performed an experimental investigation of the flowfield of a transverse jet into supersonic flow with a pseudo-shock wave (PSW). In this study, we injected compressed air as the injectant, simulating hydrocarbon fuel. A back pressure control valve generated PSW into Mach 2. 5 supersonic flow and controlled its position. The positions of PSW were set at nondimensional distance from the injector by the duct height (x/H) of -1. 0, -2. 5, and -4. 0. Particle image velocimetry (PIV) gave us the velocity of the flowfield. Mie scattering of oil mist only with the jet was used to measure the spread of the injectant. Furthermore, gas sampling measurements at the exit of the test section were carried out to determine the injectant mole fraction distributions. Gas sampling data qualitatively matched the intensity of Mie scattering. PIV measurements indicated that far-upstream PSW decelerated the flow speed of the main stream and developed the boundary layer on the wall of the test section. The flow speed deceleration at the corner of the test section was remarkable. The PSW produced nonuniformity in the main stream and reduced the momentum flux of the main stream in front of the injector. The blowing ratio, defined as the square root of the momentum flux ratio, of the jet and the main stream considering the effect of the boundary layer thickness was shown to be a useful parameter to explain the jet behavior. © 2012 Springer-Verlag.

    DOI: 10.1007/s00193-012-0384-9

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  • Turbulent Characteristics for Jet Injected into Supersonic Flow with Pseudo Shock Wave 査読

    Byugil Choi, Koichi Takae, Toshinori Kouchi, Goro Masuya

    JOURNAL OF PROPULSION AND POWER   28 ( 5 )   971 - 981   2012年9月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    The mean and turbulent velocity field of a transverse jet in a supersonic flow with a pseudo shock wave by stereoscopic particle image velocimetry was measured. Both the mainstream and the injectant were unheated air. The pseudo shock wave was generated and controlled by means of a flow plug driven by a stepping motor. To study the influence of the generation and development of the pseudo shock wave, measurements are taken for four different cases of the pseudo shock wave location, including one without a pseudo shock wave. Sequences of up to 750 image pairs for each measurement plane were acquired. From these data, the mean and turbulent characteristics of the flowfield were examined, such as mean velocity and vorticity, turbulence intensities, Reynolds stresses, and turbulence kinetic energy. The mean velocity and vorticity indicate that the pseudo shock wave gradually lowered the flow speed of the mainstream and increased the size of the counter-rotating vortex pair. The pseudo shock wave strongly disturbed the flowfield, inducing complicated three-dimensional distortions. The pseudo shock wave disturbed the flowfields near the walls, and the turbulence kinetic energy in this area was kept at a high level over long streamwise distances. The streamwise fluctuation intensity was about 4 times stronger than that of the vertical and lateral components, and the streamwise velocity components fluctuated even more greatly than the cross-sectional velocity components. The transverse injection moderately affected not only turbulence around the jet, but also the turbulent characteristics caused by the pseudo shock wave. The injection not only produced turbulence around the jet, but also suppressed the velocity fluctuation caused by the pseudo shock wave.

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  • Flowfield Characteristics of a Hypermixer Interacting with Transverse Injection in Supersonic Flow 査読

    Chae-Hyoung Kim, In-Seuck Jeung, Byungil Choi, Toshinori Kouchi, Goro Masuya

    AIAA JOURNAL   50 ( 8 )   1742 - 1753   2012年8月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    Nonreacting experiments were conducted at Mach 2 to understand the relationship between a wall-mounted alternating-ramp-wedge-type hypermixer and transverse injection compared with a step mixer. Experimental techniques such as schlieren visualization and stereoscopic particle image velocimetry were employed to study the three-dimensional flowfield with vortex structures induced by the interaction between the mixer and transverse injection. Without injection, the hypermixer creates a region that contains vortical motions. When a transverse jet is injected in the vortical region, flow parameters on mixing performance, such as the penetration height, vorticity, and turbulent stress, are improved. Three turbulent-stress components show relatively distinct distributions, and strong turbulent stresses are observed at the upper side of the jet plume. Turbulent-stress distributions are significantly associated with strain rate, which peaks along the sonic line of the jet plume.

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  • Mechanism and Control of Combustion-Mode Transition in a Scramjet Engine 査読

    Toshinori Kouchi, Goro Masuya, Tohru Mitani, Sadatake Tomioka

    JOURNAL OF PROPULSION AND POWER   28 ( 1 )   106 - 112   2012年1月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    A sidewall compression scramjet engine operated in two combustion modes under Mach 6 flight condition, weak. and intensive-combustion modes. The weak mode occurred below the overall fuel equivalence ratio (Phi) of around 0.4. Transition from the weak mode to the intensive mode occurred at Phi similar to 0.4, accompanied by a sudden increase in thrust. Mechanisms of the transition were numerically investigated in this study. Simulations captured the sudden increase in thrust at the mode transition. In the weak mode, combustion occurred in only a region near the topwall where an igniter was installed. The combustion region expanded toward the cowl with boundary-layer separation at the mode transition. Simulations demonstrated that low ignition capability resulted in the weak mode. This study demonstrated that the presence of additional igniters on the sidewalls improved the ignition capability and achieved the intensive mode in the entire Phi range.

    DOI: 10.2514/1.B34172

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  • Quasi-one dimensional modeling on vitiation effects for a dual-mode combustor

    Junji Noda, Masaki Ohkoshi, Toshinori Kouchi, Goro Masuya, Sadatake Tomioka, Robert D. Rockwell, Christopher P. Goyne

    18th AIAA/3AF International Space Planes and Hypersonic Systems and Technologies Conference 2012   ( 2012-5862 )   2012年

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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 three-flow-tube, quasi-one dimentional calculation was carried out to estimate modetransition fuel equivalence ratios under clean and vitiated inflow conditions. A simple flame holding model was used in the calculation. In cases of vitiated inflows, both total temperature and total enthalpy matched conditions to that of clean air were tested. Calculated wall pressure distributions were well matched to those of experimental values, qualitatively. Additionally, the ratio of mode-transition equivalence ratio of clean inflow divided by that of vitiated inflow was well matched to that of experimental value. Matched total enthalpy was effective to eliminate the inflow vitiation effects on mode-transition equivalence ratios in a case that thermal choking was attained at constant area exit. However, in a case that thermal choking was attained in the diverging region, matched total enthalpy was not effective. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/6.2012-5862

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  • Quantification of convection velocity and dominant scale of large-scale structures by high-speed schlieren imaging

    Toshinori Kouchi, Goro Masuya

    48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2012   2012年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    We proposed a new method to evaluate simultaneously both convective velocity and scale of dominant turbulent structures captured in a high speed framing schlieren movie. We took 90 time-series schlieren images of a transverse jet into a Mach 2 supersonic flow with 1 MHz sampling. We focused on periodical image gradations due to the turbulent structures within the jet and boundary layer. We remapped the image intensities on the structures along their trajectory as a time-space map. The time-space map showed swept stripe patterns. The presence of the stripe patterns indicates stable structures were periodically generated and convected downstream. Slope of the stripes corresponds to the convective velocities of the structures. Cycles in the space- and time-directions of the stripes correspond to the scale and formation cycle of the dominant structure. Two-dimensional Fourier transform efficiently extracted the dominant ones from the time-space map. We demonstrated this newlyproposed method using two-dimensional Fourier transform accurately evaluated the convective velocity and scale of the dominant turbulent structures within the jet and boundary layer. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/6.2012-4148

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  • Focusing-schlieren visualization in direct-connect dual-mode scramjet

    Toshinori Kouchi, Christopher P. Goyne, Robert D. Rockwell, Roger Reynolds, Roland Krauss, James C. McDaniel

    18th AIAA/3AF International Space Planes and Hypersonic Systems and Technologies Conference 2012   2012年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    Schlieren imaging has high sensitivity for density gradient in a flow field, and easily captures not only shock and expansion waves but also turbulent structures in the flow field. The conventional schlieren imaging, however, cannot capture clearly them in directconnect supersonic combustor tests with long duration, because the severe temperature condition in the combustor locally changes the refractive index of the window glasses for optical access. The conventional schlieren technique is essentially sensitive to the entire region of the light path including the glasses. On the other hand, focusing-schlieren technique has narrow depth of focus. Therefore, the imaging is expected to avoid the thermal distortion of the glasses. In this work, we applied focusing-schlieren imaging to the supersonic combustion tests, and tried to clearly capture the flow features inside a dualmode scramjet. The present system had about ± 5 mm depth of focus, and successfully visualized the flow field in the combustor, though the system sensitivity declined with the tunnel heating up. Combustion drastically changed the flow field inside the combustor, and induced volumetric expansion of fuel jet, resulting in pressure rise. Combustiongenerated pressure rise pushed up the shock train up to the trilling edge of the ramp fuel injector. The system recorded two image pairs by using PIV system. These image pairs yielded the convection velocity of turbulent structures in the combustor. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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  • Large-eddy /reynolds-averaged navier-stokes simulations of a dual-mode scramjet combustor

    Jesse A. Fulton, Jack R. Edwards, Hassan A. Hassan, Robert Rockwell, Christopher Goyne, Jim McDaniel, Chad Smith, Andrew Cutler, Craig Johansen, Paul M. Danehy, Toshinori Kouchi

    50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition   2012年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    Numerical simulations of reacting and non-reacting flows within a scramjet combustor configuration experimentally mapped at the University of Virginia's Scramjet Combustion Facility (operating with Configuration "A") are described in this paper. Reynolds-Averaged Navier-Stokes (RANS) and hybrid Large Eddy Simulation /Reynolds-Averaged Navier-Stokes (LES /RANS) methods are utilized, with the intent of comparing essentially 'blind' predictions with results from non-intrusive flow-field measurement methods including coherent anti-Stokes Raman spectroscopy (CARS), hydroxyl radical planar laser-induced fluorescence (OH-PLIF), stereoscopic particle image velocimetry (SPIV), wavelength modulation spectroscopy (WMS), and focusing Schlieren. NC State's REACTMB solver was used both for RANS and LES /RANS, along with a 9-species, 19-reaction H2-air kinetics mechanism by Jachimowski. Inviscid fluxes were evaluated using Edwards' LDFSS flux-splitting scheme, and the Menter BSL turbulence model was utilized in both full-domain RANS simulations and as the unsteady RANS portion of the LES /RANS closure. Simulations were executed and compared with experiment at two equivalence ratios, pdbl= 0.17 and pdbl= 0.34. Results show that the pdbl = 0.17 flame is hotter near the injector while the pdbl = 0.34 flame is displaced further downstream in the combustor, though it is still anchored to the injector. Reactant mixing was predicted to be much better at the lower equivalence ratio. The LES /RANS model appears to predict lower overall heat release compared to RANS (at least for pdbl = 0.17), and its capability to capture the direct effects of larger turbulent eddies leads to much better predictions of reactant mixing and combustion in the flame stabilization region downstream of the fuel injector. Numerical results from the LES/RANS model also show very good agreement with OH-PLIF and SPIV measurements. An un-damped long-wave oscillation of the pre-combustion shock train, which caused convergence problems in some RANS simulations, was also captured in LES /RANS simulations, which were able to accommodate its effects accurately. Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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  • Effect of vent mixer to mixing characteristics in supersonic flow

    Chae Hyoung Kim, In Seuck Jeung, Byungil Choi, Toshinori Kouchi, Goro Masuya

    28th Congress of the International Council of the Aeronautical Sciences 2012, ICAS 2012   4   2864 - 2872   2012年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    A Mach 2 hydrogen - air supersonic combustor model with no diffuser was designed and operated in atmospheric inflow condition. A new supersonic mixer, which was vent slot mixer (VSM), was developed and fabricated to use as the device of the fuel-air mixing. At low enthalpy inflow condition, a plasma jet torch was used as the igniter and the flame-holder. An isolator was located between the nozzle and the combustor. In supersonic combustion, the VSM had an effect to increase the mixing efficiency compared with the step mixer. While the combustion pressure was increased, the combustion mode was changed from unstable supersonic combustion to dual-mode transition. When the unstable flow of shock train affected the combustor inlet, the VSM was showed that the combustion stability was slightly sustained when compared with the case ofthe step mixer.

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  • Large-eddy simulation of pseudo-shock wave in a square duct

    S. Lee, J. Watanabe, T. Kouchi, K. Takita, G. Masuya

    18th AIAA/3AF International Space Planes and Hypersonic Systems and Technologies Conference 2012   2012年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    Large-eddy simulation (LES) of Pseudo-Shock Wave (PSW) in a square duct was perfomed to study the three-dimensional turbulent structure. A flow with the PSW in the square section duct produced by the high back pressure was reproduced using LES. The inflow Mach number was 2.5 and the inflow boundary layer of 4.2mm thick. The incoming boundary layer for a square duct was established using the multi rescaling method. The Mach number distribution, schlieren image, wall pressure, and velocity data obtained from the LES were compared with the past experimental data. The shock train structure observed in the experiment was well reproduced by the LES. The averaged wall pressure and streamwise velocity distributions agreed well between the LES and the experiments. Also, the distributions of intensity and spatial correlation of velocity fluctuation reasonably agreed between the LES and the experiment, although some differences were observed in the velocity fluctuation intensity distribution at the position of front shock wave. © 2012 by the American Institute of Aeronautics and Astronautics, Inc.

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  • Velocity field in secondary combustor of ejector-jet

    Kiyoshi Nojima, Toshinori Kouchi, Goro Masuya, Sadatake Tomioka

    18th AIAA/3AF International Space Planes and Hypersonic Systems and Technologies Conference 2012   ( 2012-5838 )   2012年

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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-component velocity distributions in a rocket-ramjet combined cycle engine model operating under a sea-level static condition were measured with a stereoscopic particle image velocimetry. The engine model had two rocket nozzles on top wall side of a flow-path and fuel injectors for the secondary combustion. The rocket exhaust was simulated by the cold compressed air injected at operation condition of 2.4 MPa in chamber pressure. The secondary fuel was simulated by the helium gas. The pressure rise due to combustion was simulated by contracting the exit area of flow path by using a flow plug. The separation region appeared in the cowl side in the ensemble-averaged velocity distribution in the case of combustion simulation. The flow around the separation region was strongly disturbed. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/6.2012-5838

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  • Turbulent Structure of Supersonic Flowfield 査読

    Shohei Uramoto, Toshinori Kouchi, Goro Masuya

    Journal of Fluid Science and Technology   7 ( 2 )   231 - 241   2012年

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    記述言語:英語   掲載種別:研究論文(学術雑誌)  

    DOI: 10.1299/jfst.7.231

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  • Turbulence Induced by Transverse Injection and Pseudo-Shock Wave in Supersonic Flow

    Goro Masuya, Byung-Il Choi, Koichi Takae, Toshinori Kouchi

    20th International Symposium on Air Breathing Engines   ( ISABE Paper 2011-1525 )   2011年9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)  

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  • Numerical Study on the Turbulent Structure of Transverse Jet into Supersonic Flow 査読

    Junya Watanabe, Toshinori Kouchi, Kenichi Takita, Goro Masuya

    AIAA JOURNAL   49 ( 9 )   2057 - 2067   2011年9月

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:AMER INST AERONAUT ASTRONAUT  

    The three-dimensional turbulent structure and mixing state of a transverse air jet injected into a Mach 1.9 supersonic crossflow were investigated using a large-eddy simulation. Both a grid sensitivity study and detailed comparisons with acetone planar laser-induced fluorescence data were conducted. The large-eddy simulation results well reproduced both the distribution of averaged injectant concentration and the large-scale turbulent features in the windward region of the jet plume. Large-scale vortices on the windward side of the jet plume caused large protrusions of injectant and were the main cause of the turbulent diffusion of injectant toward the crossflow in the near field. The windward large-scale vortex structure consisted of a string of hairpinlike vortices meandering along the jet trajectory. The head of these hairpin vortices tilted toward the upstream and upward directions. The instantaneous combustible injectant mass flux was evaluated by assuming the injectant air to be ethylene. A large amount of combustible injectant existed inside the large-scale protrusions induced by the windward large-scale vortex structure. The combustible injectant mass flow rate passing through the cross section bad substantial fluctuations over time, mainly due to the intermittent appearance of the windward large-scale vortices.

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  • Mixing Characteristics of Secondary Fuel in RBCC Ejector-Jet Mode Operation

    Kiyoshi Nojima, Toshinori Kouchi, Goro Masuya, Sadatake Tomioka

    28th International Symposium on Space Technology and Science   ( ISTS 2011-a-65 )   2011年6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)  

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  • Effect of fuel injection locations with a hyper mixer in supersonic combustion

    Chae Hyoung Kim, In Seuck Jeung, Byungil Choi, Toshinori Kouchi, Goro Masuya

    47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011   ( 2011-5830 )   2011年

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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 reacting-experiment was conducted to understand the relation between a wallmounted alternating-ramp-wedge type hyper mixer and various injection locations in unheated supersonic flow. The hydrogen gas and a plasma jet torch were used as the fuel and the igniter, respectively. The parallel injection method shows the maximum combustion efficiency for the forced ignition. This is because the fuel-air mixture can be ignited in the heat source (plasma jet) and providing the fuel-air mixture to the plasma jet, i.e., the parallel injection, is very effective to enhance the combustion efficiency. In the cold main flow, the combustion region is restrained and affected by the location of the plasma jet. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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  • Effect of fuel injection location on a plasma jet assisted combustion with a backward-facing step 査読

    Chae-Hyoung Kim, In-Seuck Jeung, Byungil Choi, Toshinori Kouchi, Kenichi Takita, Goro Masuya

    PROCEEDINGS OF THE COMBUSTION INSTITUTE   33 ( 2 )   2375 - 2382   2011年

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    記述言語:英語   掲載種別:研究論文(学術雑誌)   出版者・発行元:ELSEVIER SCIENCE INC  

    Non-reacting and reacting experiments on the ignition by a plasma jet (PJ) torch were performed to understand the correlation between fuel injection location and combustion characteristics in unheated Mach 2 airflow. Fuel was injected through three sonic injectors in the recirculation region behind a backward-facing step: a parallel injector at 2 mm from the bottom wall and two normal injectors at 2 and 9 mm from the step wall. In order to mitigate the combustion pressure interaction with nozzle, an isolator was installed between the nozzle and combustor. The combustion performance of normal injection was little affected by the difference of fuel injection locations. Moreover, normally injected fuel was escaped not to be held in the recirculation region despite of low fuel injection rates. This led to lower combustion performance relative to the parallel injection which provided fuel not to leave the recirculation region. In this case, the role of the recirculation region was to fully hold fuel, and the PJ torch provided hot gases as a heat source and acted as a flame-holder to ignite fuel-air mixtures. In a low temperature inflow condition, combustible regions were constrained around the bottom wall where embedded with the PJ torch. When thermal choking occurred in the combustor, it induced shock train both in the combustor and isolator. Under this unstable condition, the combustion performance of the normal injection was lower than that of the parallel injection. This is because the normal injection led most fuel into low temperature incoming air-stream. (C) 2010 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

    DOI: 10.1016/j.proci.2010.07.057

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  • Correlation between hypermixer and fuel injection locations

    Chae Hyoung Kim, In Seuck Jeung, Byungil Choi, Toshinori Kouchi, Goro Masuya

    17th AIAA International Space Planes and Hypersonic Systems and Technologies Conference 2011   ( 2011-2343 )   2011年

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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)  

    An experiment was conducted to understand the relation between a wall-mounted alternating-ramp-wedge type hyper mixer and transverse injectors with two different injection locations in a supersonic flow. Three experimental techniques, such as a schlieren visualization, a gas-sampling method, and a stereoscopic particle image velocimetry, were employed to study flowfield having velocity components and vortex structures induced by the interaction between the hyper mixer and the transverse injections, and also to compare the difference of mixing performance of the hyper mixer driven by the different injection location. For normal 1 injection case, an injection hole is located under the compression wedge, so injected helium is immediately impinged on the wedge and widely spread downstream, leading to enhancing mixing performance and holding helium in the mixing layer. On the other hand, for normal 2 injection case, helium is injected downstream of the hyper mixer, thus two flow structures created from the hyper mixer and the transverse injection interact with each other; as momentum flux ratio is increased, the flow structure from the transverse injection plays a significant role on the mixing region, and plenty of helium is penetrated into the supersonic flow. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/6.2011-2343

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  • Large-eddy simulations of hydrogen and ethylene injections into a supersonic crossflow

    Junya Watanabe, Toshinori Kouchi, Kenichi Takita, Goro Masuya

    47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit 2011   ( 2011-5764 )   2011年

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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)  

    Turbulent structure and mixing state of hydrogen and ethylene jets transversely injected into Mach 1.9 supersonic crossflow was numerically investigated using large-eddy simulation. The feature of windward large-scale structure was largely different between the hydrogen and ethylene jets. The ethylene jet had a highly intermittent feature with intensive geometric change of jet plume. The hydrogen jet had much larger velocity difference between the jet and the crossflow in the near-field, thus much higher turbulent intensity was obtained than in the ethylene jet. In addition, the scale of energetic eddy of hydrogen jet was larger in the windward mixing layer. These probably result in the better mixing in the hydrogen jet than in the ethylene jet. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/6.2011-5764

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  • Evaluation of heat-flux on scramjet engine wall in Mach 6 flight condition

    Shuichi Ueda, Masao Takegoshi, Toshinori Kouchi, Fumiei Ono, Toshihito Saito, Muneo Izumikawa

    AIAA 57th International Astronautical Congress, IAC 2006   9   6344 - 6353   2006年

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    掲載種別:研究論文(国際会議プロシーディングス)  

    Space transportation system uses the liquid hydrogen of the fuel as a coolant of the engine. Therefore, it is necessary to predict the heat flux to the engine wall in high accuracy for efficient cooling with a minimum coolant. However, the prediction of the heat flux to the scramjet engine combustor walls, is quite difficult due to complex interaction of flow and combustion. In this study, a sidewall-compression-type scramjet engine was tested under Mach 6 flight conditions using Ramjet Engine Test Facility. The heat flux was measured using water-cooled heat-flux meters installed on the sidewall within the combustor sections. The experimental results were compared with CFD analysis of the whole engine to predict heat flux distribution on the engine wall. The numerical results of wall pressure distribution showed good agreement with the experimental data. However, the experimental results of heat flux were extremely higher than the numerical results due to the radiation heating which was not considered in the CFD analysis. Based on the results, the average heat flux of the sub-scale scramjet engine with 300 K isothermal walls was 1 MW/m2 in the stoichiometric fuel rate at Mach 6 flight conditions.

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  • Heat Flux Prediction for Scramjet Engines

    T. Kouchi, T. Mitani, T. Hiraiwa, M. Kodera, G. Masuya

    Proceedings of 1st International Conference of Flow Dynamics   2004年11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)  

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MISC

  • Flowfield characteristics on a vent slot mixer in supersonic flow

    C. Kim, K. Sung, I. -S. Jeung, B. Choi, T. Kouchi, G. Masuya

    SHOCK WAVES   20 ( 6 )   559 - 569   2010年12月

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    記述言語:英語   出版者・発行元:SPRINGER  

    A research was conducted on a new mixing device referred as a "vent slot mixer", using experimental and computational methods. The experiment was conducted in a laboratory-scale supersonic wind-tunnel of Mach number 2. Inflow air was under atmospheric air condition, and hydrogen gas was used as fuel. In addition, the computational simulation approach was performed to support the experimental result. The vent slot mixer can directly entrain the main airflow into the recirculation region, inducing complex flow structures in the recirculation region. This also leads to gradual development of the shear layer to reduce the total pressure loss mainly induced by a recompression shock. Contrary to typical shear layers of step mixer, for the vent slot mixer, two-dimensional large-scale structures and weak shocks were clearly identified around the shear layer through experimental and computational methods. When the fuel was injected from one circular injector in the recirculation region, the high fuel concentration of the vent slot mixer was evenly distributed along the spanwise direction, but with the step mixer the fuel was highly concentrated along the region downstream of the injector. Therefore, the vent slot mixer is effective to uniformly spread the fuel toward the spanwise direction in the recirculation region. As the fuel injection rate increased, the shear layer downstream of the vent slot mixer grew uniformly along the spanwise direction; consequently, shock structures such as a recompression shock and weak shocks on the shear layer were significantly mitigated at J = 3.2.

    DOI: 10.1007/s00193-010-0280-0

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  • Flowfield characteristics on a vent slot mixer in supersonic flow

    C. Kim, K. Sung, I. S. Jeung, B. Choi, T. Kouchi, G. Masuya

    Shock Waves   20 ( 6 )   559 - 569   2010年12月

     詳細を見る

    記述言語:英語   出版者・発行元:SPRINGER  

    A research was conducted on a new mixing device referred as a "vent slot mixer", using experimental and computational methods. The experiment was conducted in a laboratory-scale supersonic wind-tunnel of Mach number 2. Inflow air was under atmospheric air condition, and hydrogen gas was used as fuel. In addition, the computational simulation approach was performed to support the experimental result. The vent slot mixer can directly entrain the main airflow into the recirculation region, inducing complex flow structures in the recirculation region. This also leads to gradual development of the shear layer to reduce the total pressure loss mainly induced by a recompression shock. Contrary to typical shear layers of step mixer, for the vent slot mixer, two-dimensional large-scale structures and weak shocks were clearly identified around the shear layer through experimental and computational methods. When the fuel was injected from one circular injector in the recirculation region, the high fuel concentration of the vent slot mixer was evenly distributed along the spanwise direction, but with the step mixer the fuel was highly concentrated along the region downstream of the injector. Therefore, the vent slot mixer is effective to uniformly spread the fuel toward the spanwise direction in the recirculation region. As the fuel injection rate increased, the shear layer downstream of the vent slot mixer grew uniformly along the spanwise direction; consequently, shock structures such as a recompression shock and weak shocks on the shear layer were significantly mitigated at J = 3.2. © 2010 Springer-Verlag.

    DOI: 10.1007/s00193-010-0280-0

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  • Scalar Spatial Correlations in a Supersonic Mixing Flowfield

    Hidemi Takahashi, Hiroki Oso, Toshinori Kouchi, Goro Masuya, Mitsutomo Hirota

    AIAA JOURNAL   48 ( 2 )   443 - 452   2010年2月

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    記述言語:英語   出版者・発行元:AMER INST AERONAUT ASTRONAUT  

    Single-time two-point spatial correlations of injectant concentrations in the supersonic mixing flowfield produced by a sonic transverse injection of air into a Mach 2.0 supersonic airstream were investigated using acetone planar laser-induced fluorescence data. Side-view and end-view, contour maps were obtained in several planes to characterize the turbulent structure and three-dimensionality of the mixing flowfield. The correlation maps indicated an organized large-scale structure in the upper region of the jet. The side-view correlation maps revealed that the shape of the large dominating structure was elliptic and that its major axis turned from backward-leaning to forward-leaning as the reference point of correlation moved downstream. The end-view correlation maps showed that the instantaneous jet plume appeared by turns either in the top or in the lower sides of the time-averaged injectant plume in each cross section.

    DOI: 10.2514/1.44684

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  • Scalar spatial correlations in a supersonic mixing flowfield

    Hidemi Takahashi, Hiroki Oso, Toshinori Kouchi, Goro Masuya, Mitsutomo Hirota

    AIAA Journal   48 ( 2 )   443 - 452   2010年2月

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    記述言語:英語   出版者・発行元:AMER INST AERONAUT ASTRONAUT  

    Single-time two-point spatial correlations of injectant concentrations in the supersonic mixing flowfield produced by a sonic transverse injection of air into a Mach 2.0 supersonic airstream were investigated using acetone planar laser-induced fluorescence data. Side-view and end-view contour maps were obtained in several planes to characterize the turbulent structure and three-dimensionality of the mixing flowfield. The correlation maps indicated an organized large-scale structure in the upper region of the jet. The side-view correlation maps revealed that the shape of the large dominating structure was elliptic and that its major axis turned from backward-leaning to forward-leaning as the reference point of correlation moved downstream. The end-view correlation maps showed that the instantaneous jet plume appeared by turns either in the top or in the lower sides of the time-averaged injectant plume in each cross section. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

    DOI: 10.2514/1.44684

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  • A plasma jet assisted combustion with a backward- facing step by fuel injection locations

    Chae-Hyoung Kim, In-Seuck Jeung, Byungil Choi, Toshinori Kouchi, Kenichi Takita, Goro Masuya

    The 33rd International Symposium on Combustion   2010年

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  • Numerical Simulation of Pulsed Injection into a Supersonic Crossflow

    K. Sasaya, J. Watanabe, T. Kouchi, G. Masuya

    Asian Joint Conference on Propulsion and Power   AJCPP2010-058   2010年

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  • Effects of Injectant Species on Compressible Tubulent Scalar Mixing Structure

    Y. Horikoshi, H. Takahashi, T. Kouchi, G. Masuya

    Asian Joint Conference on Propulsion and Power   AJCPP2010-106   2010年

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  • Penetration Characteristics of Pulsed Injection into Supersonic Crossflow

    T. Kouchi, K. Sasaya, J. Watanabe, H. Shibayama, G. Masuya

    46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit   2010年

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  • The Effect of a Vent Slot Mixer with a Plasma Jet Torch in the Supersonic Combustion

    C. Kim, I. Jeung, B. Choi, T. Kouchi, G. Masuya

    46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit   2010年

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  • Stereoscopic PIV Measurement of Supersonic Injection Flowfield and Its Error Analysis

    S. Uramoto, S. Tsuru, T. Kouchi, G. Masuya

    Asian Joint Conference on Propulsion and Power   AJCPP2010-059   2010年

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  • Development of Ignition Enhance Method by Using Non-Equilibrium Plasma in High Speed Flow

    T. Yamamoto, Y. Matsubara, K. Takita, T. Kouchi

    Asian Joint Conference on Propulsion and Power   AJCPP2010-037   2010年

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  • Effects of Injectant Species on Compressible Tubulent Scalar Mixing Structure

    Y. Horikoshi, H. Takahashi, T. Kouchi, G. Masuya

    Asian Joint Conference on Propulsion and Power   AJCPP2010-106   2010年

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  • Stereoscopic PIV Measurement of Supersonic Injection Flowfield and Its Error Analysis

    S. Uramoto, S. Tsuru, T. Kouchi, G. Masuya

    Asian Joint Conference on Propulsion and Power   AJCPP2010-059   2010年

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  • Development of Ignition Enhance Method by Using Non-Equilibrium Plasma in High Speed Flow

    T. Yamamoto, Y. Matsubara, K. Takita, T. Kouchi

    Asian Joint Conference on Propulsion and Power   AJCPP2010-037   2010年

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  • Penetration Characteristics of Pulsed Injection into Supersonic Crossflow

    T. Kouchi, K. Sasaya, J. Watanabe, H. Shibayama, G. Masuya

    46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit   2010年

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  • The Effect of a Vent Slot Mixer with a Plasma Jet Torch in the Supersonic Combustion

    C. Kim, I. Jeung, B. Choi, T. Kouchi, G. Masuya

    46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit   2010年

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  • A plasma jet assisted combustion with a backward- facing step by fuel injection locations

    Chae-Hyoung Kim, In-Seuck Jeung, Byungil Choi, Toshinori Kouchi, Kenichi Takita, Goro Masuya

    The 33rd International Symposium on Combustion   2010年

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  • Numerical Simulation of Pulsed Injection into a Supersonic Crossflow

    K. Sasaya, J. Watanabe, T. Kouchi, G. Masuya

    Asian Joint Conference on Propulsion and Power   AJCPP2010-058   2010年

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  • Supersonic Combustion with Supersonic Injection Through Diamond-Shaped Orifices

    Sadatake Tomioka, Toshinori Kouchi, Ryo Masumoto, Muneo Izumikawa, Akiko Matsuo

    45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit   AIAA-2009-5229   2009年

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  • Performance of a scramjet engine model in Mach 6 flight condition

    S. Ueda, T. Kouchi, M. Takegoshi, S. Tomioka, K. Tani

    SHOCK WAVES, VOL 2, PROCEEDINGS   2   1129 - 1134   2009年

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    記述言語:英語   出版者・発行元:SPRINGER-VERLAG BERLIN  

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  • Mechanism of Mixing Enhanced by Pseudo-Shock Wave

    Hideki Yamauchi, Byongil Choi, Toshinori Kouchi, Goro Masuya

    47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace   AIAA-2009-25   2009年

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  • Influence of Velocity Fields on Mixing in Pseudo-Shock Wave

    Goro Masuya, Byongil Choi, Hideki Yamauchi, Toshinori Kouchi

    19th ISABE conference   ISABE-2009-1355   2009年

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  • Time- Space Trajectory of Unsteady Jet into Supersonic Crossflow Using High- Speed Framing Schlieren Images

    T. Kouchi, T. Hoshino, K. Sasaya, G. Masuya

    16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference   AIAA-2009-7316   2009年

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  • Mechanism of mixing enhanced by pseudo-shock wave

    Hideki Yamauchi, Byongil Choi, Toshinori Kouchi, Goro Masuya

    47th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition   AIAA-2009-25   2009年

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    An experiment was conducted to investigate the mixing mechanism of a jet injected perpendicularly into a mainstream with a pseudo-shock wave (PSW). The velocity fields in the cross-sectional plane were obtained by particle-image velocimetry (PIV). Mixing of the injectant was observed by Mie scattering light of oil mist seeded in the injectant. The PSW thickened the boundary layer and reduced the momentum flux of the mainstream so the injectant penetrate higher from the wall. This affected the injectant mixing. The rd-scale model was applicable to estimate the injectant mixing in PSW by considering the flow condition just upstream of the injection jet. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

    DOI: 10.2514/6.2009-25

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  • Scalar spatial correlations in a supersonic mixing flowfield

    Hidemi Takahashi, Hiroki Oso, Toshinori Kouchi, Goro Masuya, Mitsutomo Hirota

    47th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition   AIAA-2009-23   2009年

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    Scalar spatial correlations in the supersonic mixing flowfield produced by a sonic transverse injection of air into a Mach 2.0 supersonic airstream were investigated using acetone planar laser-induced fluorescence (acetone PLIF) data. Side-view and end-view images were obtained in several planes to characterize the turbulent structure and three-dimensionality of the mixing flowfield. Single-time two-point spatial correlations of injectant concentration were obtained from the acetone PLIF data. Contour maps of the correlations indicated an organized large-scale structure in the upper region of the jet. Side-view correlation maps revealed that the shape of the large dominating structure was elliptic, and that its major axis turned from backward-leaning to forward-leaning as the reference point of correlation moved downstream. End-view correlation maps revealed that the instantaneous jet plume appeared by turns either in the top or in the lower sides of the timeaveraged injectant plume in each cross-section. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

    DOI: 10.2514/6.2009-23

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  • Supersonic Combustion with Supersonic Injection Through Diamond-Shaped Orifices

    Sadatake Tomioka, Toshinori Kouchi, Ryo Masumoto, Muneo Izumikawa, Akiko Matsuo

    45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit   AIAA-2009-5229   2009年

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  • Performance of a scramjet engine model in Mach 6 flight condition

    S. Ueda, T. Kouchi, M. Takegoshi, S. Tomioka, K. Tani

    SHOCK WAVES, VOL 2, PROCEEDINGS   2   1129 - 1134   2009年

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    記述言語:英語   出版者・発行元:SPRINGER-VERLAG BERLIN  

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  • 이차원 비대칭형 초음속 노즐 설계와 실험적 검증

    김채형, 성근민, 정인석, 최병일, Toshinori Kouchi, Goro Masuya

    韓國航空宇宙學會誌   37 ( 9 )   2009年

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  • Time-space trajectory of unsteady jet into supersonic crossflow using high-speed framing schlieren images

    Toshinori Kouchi, Takahisa Hoshino, Keita Sasaya, Goro Masuya

    16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference   AIAA-2009-7316   2009年

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    Supersonic transverse jet flow-field has an important role for supersonic combustor in scramjet engines. Supersonic speed induces the difficulty in unsteady measurements, because characteristics time is very short in the flow. Lacks of time-series data make the combustor difficult to develop. To investigate the jet unsteady motion, high-speed framing camera captured 100 successive schlieren images of the supersonic transverse jet flow-field with 250 kHz sampling rate. The mainstream Mach number was 2.0 and the injectant was helium gas, simulating fuel hydrogen. This work revealed the effects of density-weighted jetto-crossflow velocity ratio (blowing ratio: r) and the injector diameter (D) on the motion of the jet boundary. Newly suggested image processing, based on the Sobel spatial filter, efficiently tracked the jet boundary from the time-series schlieren images. 2700 sample data gave the probability density function (PDF) of the jet boundary. The PDF distributions indicated the jet penetration was insensitive to both r and D in rD space. The jet spreading, however, was sensitive to both r and D even in rD space. This was due the compressibility effect generated at the near field of injector. Remapping the time-series jet boundary as the time-space jet trajectory indicated the convective velocity of the large eddies in the jet was insensitive to r and D. It was constant of 460 m/s. The map gave us another important information on the jet motion: cycle of high-penetrated eddy formation in the jet. Spectrum analysis of the time-space jet trajectory investigated the shedding frequencies of highpenetrated eddy in the transverse jet. As the results, we find out that the high-penetrated eddy occurred at comparatively various frequencies. The specific amplified frequency of the jet instability probably did not exist at the far field of supersonic transverse injection. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

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  • Influence of Velocity Fields on Mixing in Pseudo-Shock Wave

    Goro Masuya, Byongil Choi, Hideki Yamauchi, Toshinori Kouchi

    19th ISABE conference   ISABE-2009-1355   2009年

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  • Development of a new supersonic mixer for dual-mode scramjet engine combustor

    Chae Hyoung Kim, In Seuck Jeung, Byungil Choi, Toshinori Kouchi, Goro Masuya

    16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference   AIAA-2009-7257   2009年

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    The Mach 2 hydrogen-air supersonic combustor model without diffuser was designed and tested in atmospheric inflow condition. A new supersonic mixer, which was vent slot mixer (VSM), was developed and confirmed its performance as a mixing device. At low enthalpy inflow condition, plasma jet torch was used as the igniter and the flame-holder. The VSM showed that the interaction between the fuel and the air through the vent spread the fuel to the spanwise direction in the recirculation region. The shock train structure generally depends on the combustion pressure, and the combustor condition becomes the unstart condition with combustion pressure increase. Although the mixer loses its mixing efficiency under the shock train, the VSM shows that it stabilizes the combustion around the vent. Therefore, the VSM has distinct characteristics for the mixing efficiency and the combustion stabilization. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.

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  • Scalar Spatial Correlations in a Supersonic Mixing Flowfield

    Hidemi Takahashi, Hiroki Oso, Toshinori Kouchi, Goro Masuya

    47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace   AIAA-2009-23   2009年

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  • Design and Experimental Verification of Two Dimensional Asymmetric Supersonic Nozzle

    Chae-hyoung Kim, Kunmin Sung, In-Seuck Jeung, Byoungil Choi, Toshinori Kouchi, Goro Masuya

    journal of KSAS   37 ( 9 )   2009年

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  • Development of a New Supersonic Mixier for Dual- Mode Scramjet Engine Combustor

    C. Kim, I. Jeung, B. Choi, Y. Matsubara, T. Kouchi, G. Masuya

    16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference   AIAA-2009-7257   2009年

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  • Matched pressure injections into a supersonic crossflow through diamond-shaped orifices

    Sadatake Tomioka, Muneo Izumikawa, Toshinori Kouchi, Goro Masuya, Kohshi Hirano, Akiko Matsuo

    JOURNAL OF PROPULSION AND POWER   24 ( 3 )   471 - 478   2008年5月

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    記述言語:英語   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    Matched pressure injections through diamond-shaped injectors were applied to a Mach 2.5 supersonic crossflow, and penetration and mixing characteristics of the injected plume were experimentally investigated. In determining injection conditions, the effective backpressure to the injectant plume was assumed to be equal to pressure on a solid-wedge surface with the identical wedge angle to the injector orifice at a designed flow rate. Both subsonic and supersonic injections were introduced to attain the required low plume pressure at a high supply pressure, ensuring a stable injectant flow rate in reacting flows with high backpressures. The matched pressure injections through the diamond-shaped orifices resulted in little jet-airflow interaction. With the supersonic injection, the plume floated from the injection wall, and the best penetration height was attained, whereas the benefit of matched pressure supersonic injection over the matched pressure sonic injection was not as remarkable as the circular injector case. The penetration height increased at an overexpanded condition, while the maximum mass fraction decay was insensitive to the injection pressure. In the case with the subsonic injection, the plume shape was similar to a pillar, and a certain fraction of the injectant was left within the boundary layer region. The penetration height as well as the maximum mass fraction decay was found to be insensitive to the injection pressure.

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  • 複合エンジンの静止大気中における吸い込み性能(第1報数:値計算による吸い込み性能予測)

    河内俊憲, 富岡定毅, 苅田丈士

    日本航空宇宙学会論文集   56 ( 650 )   110 - 115   2008年

  • Pumping Performance of RBCC Engine under Sea Level Static Condition (2nd Report: Two-Stream Flow Analysis of Ejector and Improvement of Pumping Performance)

    河内俊憲, 富岡定毅, 苅田丈士

    日本航空宇宙学会論文集   56 ( 651 )   163 - 168   2008年

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  • Pumping Performance of RBCC Engine under Sea Level Static Condition (1st Report: Numerical Prediction of Pumping Performance))

    河内俊憲, 富岡定毅, 苅田丈士

    日本航空宇宙学会論文集   56 ( 650 )   110 - 115   2008年

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  • Combustion performance of supersonic combustor with stinger-shaped fuel injector

    T. Kouchi, K. Hirano, A. Matsuo, K. Kobayashi, S. Tomioka, M. Izumikawa

    44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit   AIAA-2008-4503   2008年

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    Supersonic combustor with single "stinger-shaped" injector or multi ones arranged side-by-side in the spanwise direction were tested in a direct-connect wind-tunnel facility at a Mach 2.5 with a stagnation condition of 1.05 MPa and 2060 K. The stinger-shaped injector, having sharp leading edge followed by a streamwise slit, produced higher jet penetration in the height-wise direction as compared with the circular injector case at low jet dynamic pressure regime in our previous cold flow tests. In the combustion test, the single stinger-shaped injector produced higher jet penetration, resulting in 10 % higher thrust than the circular case when combustion region was limited in the diverging-area combustor. This injector also provided higher thrust even when flame was anchored around the injector with increasing fuel equivalence ratio (i.e. jet dynamic pressure), though the penetration reduced. This was because the stinger-shaped injector had better flame-holding performance near the injector. In the multi injector case, even with the stinger-shaped injector, the plume growing in the vertical direction, the adjacent fuel jets merged and the large plume was formed. The advantage of the stinger shape was diminished. The thrust performance in the stinger case, however, remained higher than that in the circular case. © 2008 by the American Institute of Aeronautics and Astronautics, Inc.

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  • 複合エンジンの静止大気中における吸い込み性能(第2報:エジェクタ二層流解析と吸い込み性能改善)

    河内俊憲, 富岡定毅, 苅田丈士

    日本航空宇宙学会論文集   56 ( 651 )   163 - 168   2008年

  • Aerodynamic Experiments of Small Scale Combined Cycle Engine in Various Mach Numbers

    Kouichiro Tani, Toshinori Kouchi, Kanenori Kato, Noboru Sakuranaka, Syuuichi Watanabe

    Proceedings of the 26th International Symposium on Space Technology and Science   2008年

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  • Aerodynamic Experiments of Small Scale Combined Cycle Engine in Various Mach Numbers

    Kouichiro Tani, Toshinori Kouchi, Kanenori Kato, Noboru Sakuranaka, Syuuichi Watanabe

    Proceedings of the 26th International Symposium on Space Technology and Science   2008年

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  • Matched pressure injections into a supersonic crossflow through diamond-shaped orifices

    Sadatake Tomioka, Muneo Izumikawa, Toshinori Kouchi, Goro Masuya, Kohshi Hirano, Akiko Matsuo

    Journal of Propulsion and Power   24 ( 3 )   471 - 478   2008年

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    記述言語:英語   出版者・発行元:AMER INST AERONAUTICS ASTRONAUTICS  

    Matched pressure injections through diamond-shaped injectors were applied to a Mach 2.5 supersonic crossflow, and penetration and mixing characteristics of the injected plume were experimentally investigated. In determining injection conditions, the effective backpressure to the injectant plume was assumed to be equal to pressure on a solid-wedge surface with the identical wedge angle to the injector orifice at a designed flow rate. Both subsonic and supersonic injections were introduced to attain the required low plume pressure at a high supply pressure, ensuring a stable injectant flow rate in reacting flows with high backpressures. The matched pressure injections through the diamond-shaped orifices resulted in little jet-airflow interaction. With the supersonic injection, the plume floated from the injection wall, and the best penetration height was attained, whereas the benefit of matched pressure supersonic injection over the matched pressure sonic injection was not as remarkable as the circular injector case. The penetration height increased at an overexpanded condition, while the maximum mass fraction decay was insensitive to the injection pressure. In the case with the subsonic injection, the plume shape was similar to a pillar, and a certain fraction of the injectant was left within the boundary layer region. The penetration height as well as the maximum mass fraction decay was found to be insensitive to the injection pressure. Copyright © 2007 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/1.35177

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  • Combustion Performance of Supersonic Combustor with Stinger-Shaped Fuel Injector

    T. Kouchi, K, Hirano, A. Matsuo, S. Tomioka, K. Kudo, M. Izumikawa

    44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit   AIAA-2008-4503   2008年

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  • New Injector Geometry for Penetration Enhancement of Perpendicular Jet into Supersonic Flow

    K. Hirano, A. Matsuo, T. Kouchi, S. Tomioka, M. Izumikawa

    43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit   AIAA-2007-5028   2007年

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  • Pulsed transverse injection applied to a supersonic flow

    Toshinori Kouchi, Noboru Sakuranaka, Muneo Izumikawa, Sadatake Tomioka

    Collection of Technical Papers - 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference   4   3947 - 3958   2007年

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    An experimental investigation was conducted to reveal jet penetration and mixing performance of pulsed injection in a Mach 2.5 crossflow. Helium and nitrogen gas were injected perpendicularly through flush-mounted circular sonic orifice. Probing techniques including species composition sampling and high speed framing schlieren were employed to determine the penetration and mixing performance at several downstream locations. Our investigation consisted essentially of two parts. The first part was an investigation of the continuous jet. The performance of the continuous jet was mainly controlled by effective velocity ratio (r) as the square root of the momentum-flux ratio and the orifice diameter (d). The centerline trajectories of the jets and the maximum concentration decay were collapsed by the rd scale and the ratio of oncoming boundary-layer thickness to the injector diameter. The second part was an comparison of the performance between the pulse and continuous jets. The penetration of the pulse jet was adjustable by changing the pulse duty cycle at a condition of fixed injectant mass flow rate. Even at a condition of fixed injection pressure, the pulsed injection showed better mixing performance and the higher penetration, due to the fluctuation of the large-scale eddies in the jet associated with the fluctuation of the bow shock in front of the jet.

    DOI: 10.2514/6.2007-5405

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  • New injector geometry for penetration enhancement of perpendicular jet into supersonic flow

    K. Hirano, A. Matsuo, T. Kouchi, M. Izumikawa, S. Tomioka

    Collection of Technical Papers - 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference   1   238 - 252   2007年

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    Both the wind tunnel experiments and the numerical analyses on helium injections with various injector geometry configurations were performed in order to achieve high penetration and effective mixing between supersonic crossflow and injectant. The plume from the multiple-type injector, which is drawn upon the genetic algorithm, prevented plume diffusion because of aiming to reduce total pressure loss. The injectant from the orifice, which has sharp leading edge, mixed more effectively than the case of blunt leading edge injector and the smaller the half angle of the wedge shaped injector was, the larger the penetration height was obtained. In order to reduce the disturbance to the mainstream, the stinger shaped injector was proposed using the numerical analyses. When the stinger shaped injector was used, the penetration height increased by approximately 60% in comparison with the circular injector case and showed the most effective penetration performance. It is considerable that as the disturbance became smaller, the penetration height became larger. However, the stinger shaped injector has effective performance for the penetration in the condition of low dynamic pressure ratio only unlike the circular injector.

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  • Matched pressure injections into a supersonic crossflow through diamond-shaped orifices

    Sadatake Tomioka, Toshinori Kouchi, Muneo Izumikawa, Shunsuke Koike, Goro Masuya, Kohshi Hirano, Akiko Matsuo

    Collection of Technical Papers - 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference   4   3909 - 3918   2007年

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    Matched pressure injection through diamond-shaped injectors was applied to a Mach 2.5 supersonic cross-flow, and penetration and mixing characteristics of the injected plume were experimentally investigated. Both subsonic and supersonic injections were investigated at various injection pressure. The subsonic injection case was utilized to evaluate effective back-pressure to the plume, which was almost the same value to that on a solid wedge surface with identical wedge angle to the injector orifice at designed flow rate for matched pressure condition. With the supersonic injection, the plume floated from the injection wall, and best penetration height was attained. The penetration height increased at over-expanded condition, while maximum mass fraction decay was insensitive to the injection condition. With the subsonic injection, the plume shaped like a pillar, leaving certain fraction of the injectant within the boundary layer region. The penetration height as well as the maximum mass fraction decay was found to be insensitive to the injection pressure.

    DOI: 10.2514/6.2007-5402

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  • エンジン風洞における超音速流の閉塞

    三谷 徹, 宮島 博, 谷 香一郎, 河内俊憲, 櫻中 登, 渡邊修一

    日本航空宇宙学会論文集   55 ( 636 )   43 - 50   2007年

  • Matched Pressure Injections into a Supersonic Crossflow through Diamond-Shaped Orifices

    Sadatake, T, Kouchi, T, Izumikawa, M, Koike, S, Masuya., G, Hirano, K, Matsuo, A

    43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit   AIAA-2007-5402   2007年

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  • エンジン風洞における超音速流の閉塞

    三谷 徹, 宮島 博, 谷 香一郎, 河内俊憲, 櫻中 登, 渡邊修一

    日本航空宇宙学会論文集   55 ( 636 )   43 - 50   2007年

  • Pulsed Transverse Injection Applied to a Supersonic Flow

    T. Kouchi, N. Sakuranaka, M. Izumikawa, S. Tomioka

    43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit   AIAA-2007-5405   2007年

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  • Performance of a RBCC combustor operating in ramjet mode

    Toshinori Kouchi, Kan Kobayashi, Kenji Kudo, Atsuo Murakami, Kanenori Kato, Sadatake Tomioka

    Collection of Technical Papers - AIAA/ASME/SAE/ASEE 42nd Joint Propulsion Conference   7   5325 - 5338   2006年

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    The direct-connect combustion tests and the numerical simulations of the rocket-ramjet combined cycle combustor, requiring large bases to install rocket chambers, were conducted to determine whether the combustor could be operated in ramjet mode. The rocket engines in the combustor were operated at fuel rich condition as fuel injectors for the freestream. For the rocket operation alone, the combustor was operated in scramjet mode. By the secondary fuel injection from the cowl, the supersonic freestream was decelerated to subsonic speed by the combustion-generated shock train in front of the secondary injection. The subsonic combustion of the secondary injected fuel resulted in higher combustion and thrust performances. The further upstream secondary injection was suitable for the ramjet operation. By the increase in fuel flow rate of the secondary injection, the thermal load on both the rocket chamber and throat was reduced without decrease in thrust. Gas sampling measurements at the exit of the combustor revealed that the merger of the secondary fuel with the rocket plumes resulted in the decrease of the combustion and thrust performances.

    DOI: 10.2514/6.2006-4867

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  • Performance Comparison of Combined Fuel Injection Methods in a Staged-Combustion Scramjet Engine

    Ueda, S, Tomioka, S, Kouchi, T, Kobayashi, K, Kudo, K, Kato, K

    Proceedings of the 25th International Symposium on Space Technology and Science   121 - 126   2006年

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  • Transition of Combustion Modes in a Scramjet Engine

    Kouchi, T, Mitani, T, Tomioka, S, Ueda, S

    Proceedings of the 25th International Symposium on Space Technology and Science   107 - 114   2006年

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  • Transition of Combustion Modes in a Scramjet Engine

    Kouchi, T, Mitani, T, Tomioka, S, Ueda, S

    Proceedings of the 25th International Symposium on Space Technology and Science   107 - 114   2006年

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  • Performance of a RBCC Combustor Operating in Ramjet Mode

    T. Kouchi, K. Kobayashi, K. Kudo, A. Murakami, K. Kato, T. Sadatake

    42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit   AIAA-2006-4867   2006年

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  • Performance Comparison of Combined Fuel Injection Methods in a Staged-Combustion Scramjet Engine

    Ueda, S, Tomioka, S, Kouchi, T, Kobayashi, K, Kudo, K, Kato, K

    Proceedings of the 25th International Symposium on Space Technology and Science   121 - 126   2006年

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  • Flame structures and combustion efficiency computed for a Mach 6 scramjet engine

    Tohru Mitani, Toshinori Kouchi

    Combustion and Flame   142 ( 3 )   187 - 196   2005年8月

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    記述言語:英語   出版者・発行元:ELSEVIER SCIENCE INC  

    Our hydrogen-fueled scramjet engines with a length of 2.1 m delivered net thrusts exceeding the engine drags and exhibited fuel specific impulses of about 10 km/s under Mach 4 to 8 flight conditions. A three-dimensional, reactive CFD code using unstructured hybrid grids was developed to accelerate the engine studies. Combustion in the scramjet engine under the Mach 6 condition was simulated by using this code. In this paper, the engine testing and the CFD code were outlined first. Timewise progress of hydroxyl radicals was investigated to understand autoignition and upstream-wise developments of combustion in the engine. Autoignition occurred from the cowl section at 0.1 ms after fuel mixing was completed. The reaction zones propagated upstream at speeds of about 500 m/s and reached the backward-facing steps in the combustor at 1 ms after the autoignition. Steady-state solutions showed small flames around individual fuel jets in the combustor and a large-scale diffusion flame downstream in the engine. Sonic combustion was autonomously realized in the combustor, resulting in delivery of a maximum thrust of 2250 N in the stoichiometric condition. Variations of combustion efficiency indicated that combustion performance was determined in a narrow region with a length of 0.15 m in the combustor and that the combustion downstream of the engine was rate-controlled by a large diffusion flame. The results found by the CFD computations enable us to not only improve engine performances but also to optimize computations for scramjet engines. © 2004 Published by Elsevier Inc. on behalf of The Combustion Institute.

    DOI: 10.1016/j.combustflame.2004.10.004

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  • Flame structures and combustion efficiency computed for a Mach 6 scramjet engine

    T Mitani, T Kouchi

    COMBUSTION AND FLAME   142 ( 3 )   187 - 196   2005年8月

     詳細を見る

    記述言語:英語   出版者・発行元:ELSEVIER SCIENCE INC  

    Our hydrogen-fueled scramjet engines with a length of 2.1 m delivered net thrusts exceeding the engine drags and exhibited fuel specific impulses of about 10 km/s under Mach 4 to 8 flight conditions. A three-dimensional, reactive CFD code using unstructured hybrid grids was developed to accelerate the engine studies. Combustion in the scramjet engine under the Mach 6 condition was simulated by using this code. In this paper, the engine testing and the CFD code were outlined first. Timewise progress of hydroxyl radicals was investigated to understand autoignition and upstream-wise developments of combustion in the engine. Autoignition occurred from the cowl section at 0.1 ms after fuel mixing was completed. The reaction zones propagated upstream at speeds of about 500 m/s and reached the backward-facing steps in the combustor at I ms after the autoignition. Steady-state solutions showed small flames around individual fuel jets in the combustor and a large-scale diffusion flame downstream in the engine. Sonic combustion was autonomously realized in the combustor, resulting in delivery of a maximum thrust of 2250 N in the stoichiometric condition. Variations of combustion efficiency indicated that combustion performance was determined in a narrow region with a length of 0.15 m in the combustor and that the combustion downstream of the engine was rate-controlled by a large diffusion flame. The results found by the CFD computations enable us to not only improve engine performances but also to optimize computations for scrarnjet engines. (c) 2004 Published by Elsevier Inc. on behalf of The Combustion Institute.

    DOI: 10.1016/j.combustflame.2004.10.004

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  • Numerical simulations in scramjet combustion with boundary-layer bleeding

    T Kouchi, T Mitani, G Masuya

    JOURNAL OF PROPULSION AND POWER   21 ( 4 )   642 - 649   2005年7月

     詳細を見る

    記述言語:英語   出版者・発行元:AMER INST AERONAUT ASTRONAUT  

    Airframe-integrated scramjet engines swallow the boundary layers that develop on the forebody of space planes. The scramjet engine easily falls into engine stall (engine unstart) as a result of the boundary-layer separation resulting from combustion. In this study, to investigate the unstart characteristics, numerical simulations of a whole scramjet engine with boundary-layer bleeding are performed by using a reacting How code, and the physics determining the engine performance is examined. Our computations well reproduce the engine combustion tests results with bleeding. Bleeding of 0.65% in a captured airflow suppresses the separation of the ingested boundary layer and extends the start limit from the fuel equivalence ratio of 0.5 to 1.0. The numerical results predict small discrete circular flames anchored around individual fuel jets near the injectors. These discrete flames merge to form a large envelope diffusion flame in the downstream portion of the combustor, as a result of the secondary flow produced by high pressure of the cowl shock and intensive combustion. This merged structure causes a large mass of unburned fuel and restricts the combustion efficiency and the thrust performance.

    DOI: 10.2514/1.7967

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  • Problems of Numerical Diffusion Found in Scramjets

    Kouchi, T, Mitani, T, Masuya, G

    AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference   AIAA-2005-3216   2005年

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  • Problems of numerical diffusion found in scramjets

    Toshinori Kouchi, Tohru Mitani, Goro Masuya

    A Collection of Technical Papers - 13th AIAA/CIRA International Space Planes and Hypersonic Systems and Technologies Conference   1   111 - 121   2005年

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    Numerical simulations of a scramjet combustor were conducted to determine the effects of design changes on mixing and combustion characteristics and to investigate the physics of the supersonic combustion flowfield. The mixing and combustion performances might be greatly affected by numerical diffusion due to a low grid density in the computational domains. Two test cases were examined to check the effects of numerical diffusion on the flowfields relating to the scramjet combustor. One was a complex shape closely approximating a scramjet and the other was simple transverse injection through a circular hole in a flat plate. In the two cases, the contamination levels of the flowfields by the numerical diffusion were very different. In the scramjet simulations, comparisons between the physical diffusion and the numerical diffusion indicated that the effects of the numerical diffusion are acceptable for finer solutions. In fact, the combustion efficiency converged within ±5 % at the grid with 2-mm spacing. On the other hand, the numerical diffusion of the flat plate injection using a 0.5-mm grid was comparable to the physical diffusion. The grid convergence was not proved in the current grid systems. These differences were caused by the high eddy viscosity due to the shock train of the scramjet combustor. Copyright © 2005 by the American Institute of Aeronautics and Astronautics. Inc. All rights reserved.

    DOI: 10.2514/6.2005-3216

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  • Numerical simulations in scramjet combustion with boundary-layer bleeding

    Toshinori Kouchi, Tohru Mitani, Goro Masuya

    Journal of Propulsion and Power   21 ( 4 )   642 - 649   2005年

     詳細を見る

    記述言語:英語   出版者・発行元:AMER INST AERONAUT ASTRONAUT  

    Airframe-integrated scramjet engines swallow the boundary layers that develop on the forebody of space planes. The scramjet engine easily falls into engine stall (engine unstart) as a result of the boundary-layer separation resulting from combustion. In this study, to investigate the unstart characteristics, numerical simulations of a whole scramjet engine with boundary-layer bleeding are performed by using a reacting flow code, and the physics determining the engine performance is examined. Our computations well reproduce the engine combustion tests results with bleeding. Bleeding of 0.65% in a captured airflow suppresses the separation of the ingested boundary layer and extends the start limit from the fuel equivalence ratio of 0.5 to 1.0. The numerical results predict small discrete circular flames anchored around individual fuel jets near the injectors. These discrete flames merge to form a large envelope diffusion flame in the downstream portion of the combustor, as a result of the secondary flow produced by high pressure of the cowl shock and intensive combustion. This merged structure causes a large mass of unburned fuel and restricts the combustion efficiency and the thrust performance. Copyright © 2004 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/1.7967

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  • RJTF試験におけるスクラムジェット性能達成度

    三谷 徹, 富岡定毅, 苅田丈士, 谷 香一郎, 鎮西信夫, 河内俊憲

    日本航空宇宙学会論文集   52 ( 600 )   1 - 9   2004年

  • Scramjet Performance Achieved in Engine Tests from M4 to M8 Flight Conditions

    Mitani, T, Tomioka, S, Kanda, T, Chinzei, N, Kouchi, T

    JAXA Research and Development Report   2004年

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  • Heat Flux Prediction for Scramjet Engines, - Accuracy of Reynolds Analogy on Scramjet Internal Walls

    Kouchi, T, Mitani, T, Hiraiwa, T, Kodera, M, Masuya, G

    Proceedings of the 24th International Symposium on Space Technology and Science   20 - 26   2004年

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  • Heat Flux Prediction for Scramjet Engines, - Accuracy of Reynolds Analogy on Scramjet Internal Walls

    Kouchi, T, Mitani, T, Hiraiwa, T, Kodera, M, Masuya, G

    Proceedings of the 24th International Symposium on Space Technology and Science   20 - 26   2004年

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  • RJTF試験におけるスクラムジェット性能達成度

    三谷 徹, 富岡定毅, 苅田丈士, 谷 香一郎, 鎮西信夫, 河内俊憲

    日本航空宇宙学会論文集   52 ( 600 )   1 - 9   2004年

  • Scramjet Performance Achieved in Engine Tests from M4 to M8 Flight Conditions

    Mitani, T, Tomioka, S, Kanda, T, Chinzei, N, Kouchi, T

    JAXA Research and Development Report   2004年

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  • 抗力測定によるスクラムジェットの推力性能見積り

    河内俊憲, 三谷 徹, 平岩徹夫, 富岡定毅, 升谷五郎

    日本航空宇宙学会論文集   51 ( 595 )   403 - 411   2003年

  • 抗力測定によるスクラムジェットの推力性能見積り

    河内俊憲, 三谷 徹, 平岩徹夫, 富岡定毅, 升谷五郎

    日本航空宇宙学会論文集   51 ( 595 )   403 - 411   2003年

  • Numerical Experiments of Scramjet Combustion with Boundary-Layer Bleeding

    Kouchi, T, Mitani, T, Kodera, M, Masuya, G

    AIAA 12th International Space Planes and Hypersonic Systems and Technologies Conferece   AIAA paper 2003-7038   2003年

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  • Scramjet performance achieved in engine tests from M4 to M8 flight conditions

    Tohru Mitani, Sadatake Tomioka, Takeshi Kanda, Nobuo Chinzei, Toshinori Kouchi

    12th AIAA International Space Planes and Hypersonic Systems and Technologies   AIAA paper 2003-7009   2003年

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    Thrust performances of scramjet engines were compared with theoretical values to quantify the progress in engine performances (defined as "achievement factors", or "factors") from Mach (termed as "M") 4 to M8 flight conditions. An engine with a ramp produced a net thrust of 215 N under the M8 tests and a comparison of a theoretical thrust yielded a thrust achievement factor of 51%. By excluding boundary layer, an engine with a thick strut delivered a net thrust of 560 N and showed a thrust factor of 92% and a net thrust factor of 45%. The thrusts were limited by flow separation caused by engine combustion (termed as "engine unstart"). The starting characteristics was improved by boundary layer controls in M6 and M4 conditions. An engine with a thin strut doubled the thrust from 1620 N to 2460 N by the boundary layer bleeding in the M6 tests. The improved thrust factor was 60% at the stoichiometric H2 condition. Under M4 tests, the net thrust was tripled by the bleed and a two-staged injection of H2. As results, the thrust factor was raised from 53% to 70%, the net thrust factor was increased from 32% to 55%. Studies required for improving the net performance was addressed. © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/6.2003-7009

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  • 小型風洞とエンジン補完実験

    三谷 徹, 平岩徹夫, 苅田丈士, 志村 隆, 富岡定毅, 小林 完, 泉川宗男, 櫻中登, 渡邊修一, 樽川雄一, 河内俊憲, 北村英二郎, 八並知実

    独立行政法人 航空宇宙技術研究所報告   2003年

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  • Numerical experiments of scramjet combustion with boundary-layer bleeding

    Toshinori Kouchi, Tohru Mitani, Masatoshi Kodera, Goro Masuya

    12th AIAA International Space Planes and Hypersonic Systems and Technologies   AIAA paper 2003-7038   2003年

     詳細を見る

    Airframe-integrated scramjet engines swallow the boundary-layer which develops on the airframe of space planes. The scramjet engine easily falls into engine stall (engine unstart) due to the boundary-layer separation resulting from combustion. In this study, to investigate the unstart characteristics, numerical simulations of a scramjet engine with boundary-layer bleeding were performed using a reacting flow code which includes a one-equation turbulence model. This computation is calibrated by the experimental wall pressure distributions, heat fluxes and thrusts. The computation reproduces the prevention of engine unstart in the combustion tests with bleeding. Bleeding of 0.6% in a captured flow rate suppresses the flow separation and extends the start limit from the fuel equivalence ratio (φ) of 0.5 to 1.0. The computation at φ=1.0 shows that small-scale circular diffusion flames are anchored around individual fuel jets near the injectors. These structures disappear to form a large-scale envelope diffusion flame downstream of the combustor. The circular flames near the injectors account for 80% of the combustion efficiency and control the thrust performance. © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

    DOI: 10.2514/6.2003-7038

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  • 小型風洞とエンジン補完実験

    三谷 徹, 平岩徹夫, 苅田丈士, 志村 隆, 富岡定毅, 小林 完, 泉川宗男, 櫻中登, 渡邊修一, 樽川雄一, 河内俊憲, 北村英二郎, 八並知実

    独立行政法人 航空宇宙技術研究所報告   2003年

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  • Scramjet Performance Achieved in Engine Tests from M4 to M8 Flight Conditions

    Mitani, T, Tomioka, S, Kanda, T, Chinzei, N, Kouchi, T

    AIAA 12th International Space Planes and Hypersonic Systems and Technologies Conferece   AIAA paper 2003-7009   2003年

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  • Effects of Reynolds Number on Scramjet Inlet Performance

    Kouchi, T, Mitani, T, Kodera, M, Masuya, G

    Proceedings of the 23rd International Symposium on Space Technology and Science   21   2412 - 2417   2002年

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  • Effects of Reynolds Number on Scramjet Inlet Performance

    Kouchi, T, Mitani, T, Kodera, M, Masuya, G

    Proceedings of the 23rd International Symposium on Space Technology and Science   21   2412 - 2417   2002年

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▼全件表示

講演・口頭発表等

  • ステレオPIVを用いた超音速流れ場の境界層速度計測

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • Penetration Characteristics of Pulsed Injection into Supersonic Crossflow

    46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit  2010年 

     詳細を見る

  • 超音速流中における噴流の乱流混合のLES解析

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • エジェクタ・ラムジェットの二次燃料分布

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • Penetration Characteristics of Pulsed Injection into Supersonic Crossflow

    46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit  2010年 

     詳細を見る

  • 高速シュリーレンを用いた超音速噴流場のスペクトル解 析

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • ステレオPIVを用いた超音速流れ場の境界層速度計測

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • 超音速流中における噴流の乱流混合のLES解析

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • エジェクタ・ラムジェットの二次燃料分布

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • 高速シュリーレンを用いた超音速噴流場のスペクトル解 析

    第42回流体力学講演会/ 航空宇宙数値シミュレーション技術シンポジウム2010  2010年 

     詳細を見る

  • 超音速横風中のパルス噴射において観測された噴流貫通のヒステリシス

    日本航空宇宙学会北部支部2009年講演会ならびに第10回再使用型宇宙推進系シンポジウム  2009年 

     詳細を見る

  • Time- Space Trajectory of Unsteady Jet into Supersonic Crossflow Using High- Speed Framing Schlieren Images

    16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference  2009年 

     詳細を見る

  • 超音速横風中のパルス噴射において観測された噴流貫通のヒステリシス

    日本航空宇宙学会北部支部2009年講演会ならびに第10回再使用型宇宙推進系シンポジウム  2009年 

     詳細を見る

  • Time- Space Trajectory of Unsteady Jet into Supersonic Crossflow Using High- Speed Framing Schlieren Images

    16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conference  2009年 

     詳細を見る

  • 超音速流中におけるパルス噴流の貫通特性

    第40回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム 2008  2008年 

     詳細を見る

  • 高速シュリーレン撮影を用いた超音速流中における垂直噴流の貫通の評価

    第40回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム 2008  2008年 

     詳細を見る

  • Combustion Performance of Supersonic Combustor with Stinger-Shaped Fuel Injector

    44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit  2008年 

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  • 超音速流中におけるパルス噴流の貫通特性

    第40回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム 2008  2008年 

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  • 高速シュリーレン撮影を用いた超音速流中における垂直噴流の貫通の評価

    第40回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム 2008  2008年 

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  • Combustion Performance of Supersonic Combustor with Stinger-Shaped Fuel Injector

    44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit  2008年 

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  • Pulsed Transverse Injection Applied to a Supersonic Flow

    43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit  2007年 

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  • 数値計算による複合エンジンの静止大気中における吸い込み性能予測

    第47 回 航空原動機・宇宙推進講演会  2007年 

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  • Pulsed Transverse Injection Applied to a Supersonic Flow

    43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit  2007年 

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  • 超音速流中でのパルス噴射の貫通と混合

    日本航空宇宙学会北部支部20周年記念講演会ならびに第8回再使用型宇宙推進系シンポジウム  2007年 

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  • 数値計算による複合エンジンの静止大気中における吸い込み性能予測

    第47 回 航空原動機・宇宙推進講演会  2007年 

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  • 超音速流中でのパルス噴射の貫通と混合

    日本航空宇宙学会北部支部20周年記念講演会ならびに第8回再使用型宇宙推進系シンポジウム  2007年 

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  • ロケット複合エンジンのエジェクタモードにおける始動特性

    日本航空宇宙学会北部支部2006年講演会ならびに第7回再使用型宇宙推進系シンポジウム  2006年 

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  • Performance of a RBCC Combustor Operating in Ramjet Mode

    42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit  2006年 

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  • Transition of Combustion Modes in a Scramjet Engine

    25th International Symposium on Space Technology and Science  2006年 

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  • ロケット複合エンジンのエジェクタモードにおける始動特性

    日本航空宇宙学会北部支部2006年講演会ならびに第7回再使用型宇宙推進系シンポジウム  2006年 

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  • Performance of a RBCC Combustor Operating in Ramjet Mode

    42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit  2006年 

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  • Transition of Combustion Modes in a Scramjet Engine

    25th International Symposium on Space Technology and Science  2006年 

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  • Problems of Numerical Diffusion Found in Scramjets

    AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference  2005年 

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  • Problems of Numerical Diffusion Found in Scramjets

    AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference  2005年 

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  • Heat Flux Prediction for Scramjet Engines, - Accuracy of Reynolds Analogy on Scramjet Internal Walls

    24th International Symposium on Space Technology and Science  2004年 

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  • Heat Flux Prediction for Scramjet Engines, - Accuracy of Reynolds Analogy on Scramjet Internal Walls

    24th International Symposium on Space Technology and Science  2004年 

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  • 境界層抽気がスクラムジェット性能に及ぼす影響

    日本機械学会流体工学部門講演会  2003年 

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  • 境界層抽気がスクラムジェット性能に及ぼす影響

    日本機械学会流体工学部門講演会  2003年 

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  • Numerical Experiments of Scramjet Combustion with Boundary-Layer Bleeding

    12th AIAA International Space Planes and Hypersonic Systems and Technologies  2003年 

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  • Numerical Experiments of Scramjet Combustion with Boundary-Layer Bleeding

    12th AIAA International Space Planes and Hypersonic Systems and Technologies  2003年 

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  • Effects of Reynolds Number on Scramjet Inlet Performance

    23rd International Symposium on Space Technology and Science  2002年 

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  • Effects of Reynolds Number on Scramjet Inlet Performance

    23rd International Symposium on Space Technology and Science  2002年 

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  • レイノルズ数のスクラムジェットインレット性能に及ぼす影響

    日本航空宇宙学会北部支部2002年講演会ならびに第3回再使用型宇宙推進系シンポジウム  2002年 

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  • レイノルズ数のスクラムジェットインレット性能に及ぼす影響

    日本航空宇宙学会北部支部2002年講演会ならびに第3回再使用型宇宙推進系シンポジウム  2002年 

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▼全件表示

Works(作品等)

  • 複合エンジンにおける二次燃焼効率向上

    2008年
    -
    2010年

     詳細を見る

  • 複合エンジンにおける二次燃焼効率向上

    2008年
    -
    2010年

     詳細を見る

  • 推力向上のための混合促進手法の検証

    2007年
    -
    2008年

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  • 推力向上のための混合促進手法の検証

    2007年
    -
    2008年

     詳細を見る

受賞

  • 第42回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム2010 最優秀賞

    2010年  

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  • 第42回流体力学講演会/航空宇宙数値シミュレーション技術シンポジウム2010 最優秀賞

    2010年  

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    受賞国:日本国

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  • 東北大学総長賞

    2005年  

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  • 東北大学総長賞

    2005年  

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    受賞国:日本国

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  • 日本航空宇宙学会北部支部 2002 年講演会 Best Presentation Award for Student

    2002年  

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  • 日本航空宇宙学会北部支部 2002 年講演会 Best Presentation Award for Student

    2002年  

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    受賞国:日本国

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▼全件表示

共同研究・競争的資金等の研究

  • 先進流体計測が解き明かす後退翼における遷音速バフェットのメカニズム

    研究課題/領域番号:18H03814  2018年04月 - 2022年03月

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

    河内 俊憲, 小池 俊輔, 杉岡 洋介, 橋本 敦, 石田 崇

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    配分額:44200000円 ( 直接経費:34000000円 、 間接経費:10200000円 )

    2018年度に,風洞試験で用いる後退翼模型の製作,およびこれを用いた風洞試験による模型の安全性確認と基本的な空力特性の取得を行った.またこれら風洞試験に加えて,2断面同時可視化装置の開発を開始した.
    これら結果を踏まえ,本年度はまず研究のキーとなる計測システム,2断面同時断層シュリーレン可視化システムの確立を行った.2018年度の検討から,計測領域が広いと風洞設備と計測装置の干渉が不可避であることが分かった.他方で2018年度の実験から,計測領域を絞っても,衝撃波の横揺れを検出可能であることが分かった.これらを加味し,視野100 mm,奥行き(基準断面)±50 mm,被写界深度±10 mmを目標に,2断面同時可視化システムの設計・開発を行った.その結果,これら仕様を満たすシステムの確立に成功し,「風洞を用いない実験」においてその性能の確認が出来た.
    これら可視化システムの確立に加え,既存の1断面のみを可視化できる断層シュリーレン装置を風洞実験に適用し,後退翼の各スパン位置におけるバフェット特性の取得を行った.試験条件は2018年度の試験を元に,マッハ0.72,迎角7°,およびこれの比較対象迎角4°を選定した.これら試験では,計測システムを翼スパン方向にスイープさせ,翼スパンの3/4の領域において,後退翼における衝撃波振動のデータベースを世界で初めて作ることに成功した.
    またこれらの実験に加え,2020年度に行う非定常PSP計測や2断面同時断層シュリーレン計測を行うための治具の製作,また両計測の共存性をその設置方法や光源の干渉を含め検討を行い,両立可能であることを明らかにした.
    また風洞実験に対応する数値計算を行い,風洞壁が流れ場に及ぼす影響を明らかにし,翼根のみならず,翼端でも衝撃波の横揺れが生じる可能性があることを明らかにした.

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  • イメージインテンシファイア残光を利用した全く新しい非定常流速測定法の確立

    研究課題/領域番号:17K18938  2017年06月 - 2019年03月

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

    河内 俊憲

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    配分額:6370000円 ( 直接経費:4900000円 、 間接経費:1470000円 )

    本研究では,これまで困難とされてきた超音速流における流速の非定常計測を可能にする新しい手法の確立を目指した.まず低速流で,研究の根幹となるアイデアの確認を行い,提案するアイデアで流速が計測できることを確認できた.しかしながら,超音速流れにこの手法を適用する際,低速流で問題となった点がさらに顕著になり,残念ながらこれを解決するには至らなかった.その一方で,計測に使用する粒子の新たな生成手法の開発に成功した.新手法で生成された粒子を用いれば,これまでより詳細な流速場の計測が可能になると期待される.

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  • 二色蛍光比法による超音速燃焼器内の瞬時噴流モル分率の計測

    研究課題/領域番号:15H04199  2015年04月 - 2018年03月

    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)

    河内 俊憲, 柳瀬 眞一郎, 永田 靖典

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    配分額:16380000円 ( 直接経費:12600000円 、 間接経費:3780000円 )

    次世代の宇宙輸送システムとしてスペースプレーンの研究が行われています.スペースプレーン用のエンジンでは超音速燃焼が行われ,燃料の混合を適切に評価することが重要となります.本研究では,二種類の蛍光トレーサをそれぞれ空気流と燃料噴流に添加し,同時にレーザで励起して各トレーサからの蛍光発光を測定することで,超音速燃焼器において「瞬時」モル分率を計測できる新しい計測法(二色蛍光比法)の開発を目指しました.その結果,低温環境で蛍光トレーサの蛍光特性を計測する装置や二色蛍光比法計測装置の開発に成功しました.また計測精度をさらに向上するには,流れ場の温度推定が課題であることが分かりました.

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  • 2色蛍光比法による圧縮性混合場の瞬時モル分率及び密度測定

    研究課題/領域番号:24656512  2012年

    日本学術振興会  科学研究費助成事業 挑戦的萌芽研究  挑戦的萌芽研究

    升谷 五郎, 滝田 謙一, 河内 俊憲

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    配分額:4030000円 ( 直接経費:3100000円 、 間接経費:930000円 )

    燃料と空気の迅速な混合は極超音速空気吸込みエンジンを実現するための最も重要な技術課題の一つである。瞬時の燃料モル分率分布を測定するために,2種類のトレーサー物質のレーザー誘起蛍光を用いる新しい手法を開発しようとしている。数種類の候補となる物質の蛍光特性を調査し,アセトンとトルエンをトレーサー物質として選定した。これらのガスの混合気が,空気やその他の気体中でどのような蛍光特性を示すかを実験的に調べた。

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  • PIV・LIFの同時計測によるスクラムジェット燃焼器内の混合メカニズムの解明

    研究課題/領域番号:23360377  2011年04月 - 2014年03月

    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)

    河内 俊憲, 升谷 五郎, 滝田 謙一, 富岡 定毅

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    配分額:19110000円 ( 直接経費:14700000円 、 間接経費:4410000円 )

    本研究では,現在,世界各国で研究開発が進められている超音速旅客機やスペースプレーン用の推進機関として期待されているスクラムジェットエンジン内の燃料の混合状態を計測する新しい手法の開発を世界に先駆けて行った.この計測法により,これまで超音速流では計測が困難であった渦拡散流束と呼ばれる乱流渦に起因する燃料拡散を計測できるようになった.この計測装置をエンジンを模擬した流れ場に適用し,渦拡散流束を計測した結果,これまで数値シミュレーション等で用いられてきた勾配拡散モデルが,エンジン内の燃料混合に対しても十分適用可能であることを明らかにした.これらにより今後エンジンの研究・開発が促進されると期待される.

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  • 平衡/非平衡プラズマの併用による超音速流中での着火・燃焼促進

    研究課題/領域番号:23360375  2011年04月 - 2014年03月

    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)

    滝田 謙一, 升谷 五郎, 河内 俊憲

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    配分額:18460000円 ( 直接経費:14200000円 、 間接経費:4260000円 )

    本研究は,平衡(熱)プラズマと非平衡(低温)プラズマの併用により極めてエネルギ効率の高い着火・燃焼促進技術の新規開発を目的とする。平成24年度は ①平衡/非平衡プラズマを併用した着火システムによる着火・燃焼実験,②プラズマ及び火炎の光学計測,③プラズマ反応のモデリング,④3次元燃焼解析コードを用いたプラズマ着火の数値シミュレーション の実施を計画し,以下の成果を得た。
    ①については,プラズマジェット(PJ)トーチと誘電体バリア放電(DBD)を併用した着火システムを用いて,水素及び炭化水素を燃料とする着火・燃焼実験を行った。その結果,水素燃料についてはDBD装置の併用による着火促進効果が見られた。しかしスクラムジェットの代表的な炭化水素燃料であるエチレンに対しては,目立った着火促進効果を捉えることができなかった。
    ②については今年度は実施しなかった。
    ③については,空気中でDBD装置を作動させた場合に,多く生成されることが知られているオゾンを燃焼機構に加えて着火遅れ時間の解析を行い,上記①で観測されたエチレン燃料に対してDBD装置を併用しても着火促進効果がほとんど見られない原因を探った。
    ④については水素燃料について,DBD装置の作動を模擬するオゾン添加を行う領域とオゾン添加濃度を変化させた計算を行った。その結果,断面全域に平均的にオゾンを加えた場合と同等の着火促進効果を,噴射孔側の壁面付近に集中的にオゾンを供給することができ,より効率的にオゾンの着火促進効果を得られることを見出した。

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  • 擬似衝撃波による混合促進機構の解明

    研究課題/領域番号:20360081  2008年 - 2011年

    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)

    升谷 五郎, 滝田 謙一, 河内 俊憲

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    資金種別:競争的資金

    配分額:19890000円 ( 直接経費:15300000円 、 間接経費:4590000円 )

    デュアルモードラムジェット燃焼器などの擬似衝撃波が存在する超音速内部流れにおいて, 主流中に噴射した気体燃料の混合が著しくかつ非等方的に促進される機構を解明するために, 粒子画像速度計(PIV)及び平面レーザー誘起蛍光法(PLIF)による速度場及び噴射気体濃度場の計測を行うと共に, 数値シミュレーションを行い実験結果と比較した。その結果, 擬似衝撃波が壁面付近の乱れを著しく強化し, それにより噴射気体の混合が促進されることが分かった。

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  • 超音速流中への燃料のパルス噴射による混合・貫通の促進

    研究課題/領域番号:20760105  2008年 - 2010年

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

    河内 俊憲

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    資金種別:競争的資金

    配分額:4290000円 ( 直接経費:3300000円 、 間接経費:990000円 )

    超音速流中のパルス噴流の非定常挙動を,ハイスピードシュリーレン法を中心とした風洞実験と非定常CFDによるパラメータ計算により調べた.その結果,パルス噴流の貫通特性にはヒステリシスが存在し,有効噴射速度比が上昇しているときの噴流貫通は下降時のそれと比べて大きくなること,またその原因が噴射時に生じる渦によって生じることを明らかにした.パルス周波数と噴流の貫通高さおよび混合には密接な関係があり,周波数を変化することで,噴流の貫通,および混合をアクティブにコントロールできることが明らかになった.

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  • 非平衡プラズマを用いた超音速流における着火・燃焼促進

    研究課題/領域番号:20360378  2008年 - 2010年

    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)

    滝田 謙一, 升谷 五郎, 河内 俊憲

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    配分額:18720000円 ( 直接経費:14400000円 、 間接経費:4320000円 )

    超音速流中において、誘電体バリア放電による非平衡プラズマの生成に成功し、その分光計測により着火・燃焼促進効果を有する励起窒素分子や酸素ラジカルが存在することを示した。さらに非平衡プラズマの生成に及ぼす印加電圧や印加周波数の影響を明らかにし、プラズマ生成に要する電力量の算出を行った。また、理論的研究により、高効率なラジカル生成を達成する高電圧印加法を明らかにした。

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  • 非平衡プラズマを用いた超音速流れにおける着火・燃焼促進

    2008年

    科学研究費補助金 

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    資金種別:競争的資金

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  • スクラムジェットエンジン内部形状の最適化

    研究課題/領域番号:03J07299  2003年 - 2005年

    日本学術振興会  科学研究費助成事業 特別研究員奨励費  特別研究員奨励費

    河内 俊憲

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    配分額:1700000円 ( 直接経費:1700000円 )

    極超音速飛行するスペースブレーン用のエンジンとして有望視されている,スクラムジェットエンジン内壁におけるレイノルズアナロジーの妥当性を,数値計算を用いて定量的に議論し,その適用限界を調べた.その結果±10%程度の範囲で,エンジン内壁のおよそ90%でレイノルズアナロジーが適用できることが分かった.そこで,このレイノルズアナロジーを組み込んだ一次元解析により,いくつかのエンジン形状と飛行条件について熱負荷や冷却要求を調べた.例えばストラット付きエンジンでは,マッハ8飛行条件において,エンジン平均熱流束は燃焼時に0.5MW/m2となり,冷却に必要な水素流量は,燃焼用の燃料流量と同程度になることが分かった.従って,今後エンジンの作動マッハ数域を広げるには,冷却要請のおよそ80%を占める燃焼器においてフィルム冷却を行ったり,形状を最適化していく必要がある.
    マッハ6飛行条件のストラット付きスクラムジェットエンジンと,そのエンジンの燃焼器の一部を切り出したセクター燃焼器モデルに対して反応流計算を行い,エンジン内の火炎形状が,燃焼性能をどのように決定しているのか,またインレットの偏流が燃焼性能に対してどのような影響を与えているかを考察した.その結果,インレットの偏流により,燃料の大規模な合体が生じ,燃焼効率の増加が著しく抑制されていることが分かった.従って,今後エンジン性能を改善していくには,燃料噴流の合体を防ぐために,燃料噴射孔間隔の最適化を図ること,特にインレットを最適化し,カウル衝撃波による境界層内の二次流を低減していく必要がある.また数値計算で得られた,エンジン内の火炎構造と燃焼効率分布を比較することで,供試エンジンでも,ストラット壁側からの垂直噴射,ストラットベース面・カウル側からの平行噴射を行えば、燃焼性能を改善しうることが分かった.

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