Publishing type：Research paper (scientific journal)   Publisher：IOP Publishing  

Abstract

This study presents the development of a single-scale initial velocity field into steady homogeneous isotropic turbulence. Here, we verify a constructed Fourier spectral analysis, which is then used to approach the present issue. The present homogeneous isotropic turbulence is maintained steady by using a linear forcing scheme of Lundgren (2003). Here, two values of the primary coefficient are set here for this scheme. The initial velocity field is also set to maintain the flow globally isotropic. Temporal developments of turbulence statistics, visualization results and structure functions are examined in this study. As shown in a previous study, Rosales and Meneveau (2005), the turbulence length scale at a steady-state could remain independent of a value used for the linear forcing coefficient. The turbulence fields needed 10 to 15 times longer than the turbulence time scale to achieve this steady-state. A larger amount of time was needed for the structure functions to reach their steady-state profiles than for the turbulence statistics.

DOI： 10.1088/1742-6596/2313/1/012008

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Other Link： https://iopscience.iop.org/article/10.1088/1742-6596/2313/1/012008/pdf

Authorship：Last author   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2022-1972

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Authorship：Corresponding author   Language：Japanese   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

Self-oscillation of a shock wave, which is called a transonic shock buffet, occurs on a wing surface in a transonic flow. Lee and Crouch explained the self-sustained mechanism by different pressure wave propagations but there is no experimental evidence of which feedback loop is dominant. The principal aim of the present study is to identify the dominant feedback path of pressure waves related to the shock buffet using high-speed focusing schlieren movies. The experimental data showed that the flow fields include complicated phenomena such as a lot of pressure waves with a broad range of frequency. Therefore, we performed dynamic mode decomposition (DMD) analysis to extract dominant phenomena for the buffet. DMD modes whose frequencies corresponded to the shock oscillation frequencies were extracted and investigated. DMD modes revealed that pressure perturbations generated at the shock foot and the perturbation propagated to the height-wise direction along with the shock wave. This feedback pass is similar to that observed by Crouch using the global stability analysis for NACA 0012 airfoil.

DOI： 10.2322/jjsass.70.82

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

DOI： 10.1016/j.cjph.2021.07.002

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Aeronautics and Astronautics (AIAA)  

DOI： 10.2514/1.b37683

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

The electrodynamic flow control effect using an applied magnetic field has a potential of innovative technology for aerodynamic force control of atmospheric entry vehicle because of no requirement of mechanical devices. To control this effect and aerodynamic force, we proposed a new approach using exposed electrodes on the vehicle surface. If the induced electric current is bypassed via exposed electrodes, the electromagnetic field could be changed. In this study, the influence of the partial electrical conducting wall, which is consisted of insulating and conducting walls, is numerically investigated. As a result of the MHD simulation, flow field could be changed by the vehicle surface condition. Especially, in the frontal region of the vehicle where strong Lorentz force acts, if the conducting wall is installed to electrically short to the shoulder position, the electric field within this region and the electrodynamic effect is significantly changed. Additionally, stepwise drag variation could be made by selectable shorted electrodes.

DOI： 10.2322/tastj.19.68

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Authorship：Lead author   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2020-2040

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2020-1335

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Publishing type：Research paper (scientific journal)   Publisher：Scientific Research Publishing, Inc.  

DOI： 10.4236/ojfd.2020.101003

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

DOI： 10.1088/1755-1315/240/3/032001

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Other Link： http://stacks.iop.org/1755-1315/240/i=3/a=032001?key=crossref.08f776deeae50d29012044336078de89

Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：AIP Publishing  

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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Authorship：Lead author   Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

© 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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Language：Japanese   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

<p>The electrodynamic heat shield is new type of heat protection system for atmospheric-entry vehicle, which utilizes the Lorentz force acting on the weakly ionized plasma flow inside the shock layer. A huge amount of calculation is necessary to estimate overall effect of the electrodynamic heat shield, because its effect interacts with the atmospheric-entry trajectory. In the present study, the Viscous Shock Layer (VSL) analysis method for electrodynamic heat shield is proposed for the quick analysis, which can calculate much faster than Navier-Stokes (NS) simulation. For this purpose, the VSL equations for the electrodynamic heat shield analysis are introduced under the ideal gas assumption, including circumferential momentum equation and Maxwell equations. The new method is also proposed to solve the new VSL equations. By the comparison to the NS simulation, the new VSL method gives good estimations of drag force and wall heat flux for wide ranges of interaction parameter and Hall parameter although the slight difference of wall heat flux at the stagnation point is observed. Therefore, the present VSL method could be applicable to the estimation tool of the electrodynamic heat shield effect. </p>

DOI： 10.2322/jjsass.67.16

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (international conference proceedings)  

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Language：Japanese  

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (international conference proceedings)  

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

DOI： 10.1016/j.ijthermalsci.2017.05.026

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

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 δ (=0.1) over a wide range of the Reynolds number from 3000 to 21,000 for torsion parameter β (=torsion /2δ  = 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−ε 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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Authorship：Lead author   Language：English   Publishing type：Research paper (scientific journal)   Publisher：Springer Nature  

DOI： 10.1007/s00348-017-2377-z

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Other Link： http://link.springer.com/content/pdf/10.1007/s00348-017-2377-z.pdf

Language：English   Publishing type：Research paper (scientific journal)   Publisher：Physical Society of 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, <italic>Re</italic>(= 80, 300, and 1000), and two nondimensional curvatures, δ (= 0.1 and 0.05), over a wide range of torsion parameters, β (= nondimensional torsion/<inline-formula><tex-math version="MathJax"><![CDATA[\(\sqrt{2\delta } \)]]></tex-math></inline-formula>), 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. [<ext-link ext-link-type="uri" xlink:href="https://doi.org/10.1016/0169-5983(95)00022-6" xlink:type="simple">Fluid Dyn. Res. <bold>16</bold>, 237 (1995)</ext-link>]. It was found that the friction factor sharply increases as β increases from zero, then decreases after taking a maximum, and finally slowly approaches that of a straight pipe when β tends to infinity. It is interesting that a peak of the friction factor exists in the region 0.2 ≤ β ≤ 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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Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2017-0265

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Language：Japanese   Publisher：The Japan Society of Mechanical Engineers  

<p>Numerical prediction of air-entraining and submerged vortices in pump sumps is important for engineering applications. The validation of pump sump simulations, however, still is not enough, because the simulations is very complicated; for examples, treatment of gas-liquid interface, detection method of the vortices and selection of turbulence model etc. We conducted numerical simulations of the benchmark experiments of the pump sump conducted by Matsui et al. (2006, 2016) and compared the simulation with the experimental data to investigate the effects of turbulence model, grid density and detection method of the vortices. We determined a threshold of the gas-liquid fraction function of VOF method (α) and the second invariant of velocity gradient tensor (Q₂) to detect the air-entraining and submerged vortices by using vorticity, respectively. This method well detected the vortices and well reproduced the experiments for the RANS simulation using SST k-ω model. Large eddy simulation using Smagorinsky model, however, was sensitive to the grid system and difficult to reproduce the experimental vortex structures even for the finest grid system having 3.7 million cells.</p>

DOI： 10.1299/transjsme.17-00181

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

DOI： 10.2322/tjsass.60.56

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Authorship：Lead author   Language：Japanese   Publisher：Japan Society of Mechanical Engineers  

DOI： 10.1299/transjsme.16-00441

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：Scientific Research Publishing, Inc.  

DOI： 10.4236/ojfd.2017.74043

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (international conference proceedings)  

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Authorship：Corresponding author   Language：English   Publishing type：Research paper (international conference proceedings)  

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

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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Other Link： http://link.springer.com/content/pdf/10.1007/s00348-016-2261-2.pdf

Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2016-1766

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Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

© 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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Language：English   Publishing type：Research paper (scientific journal)   Publisher：Society of Chemical Engineers, Japan  

DOI： 10.1252/jcej.15we099

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Authorship：Corresponding author   Language：Japanese   Publisher：Japan Society for Aeronautical and Space Sciences  

DOI： 10.2322/jjsass.64.244

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

DOI： 10.1007/s00348-015-2081-9

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Other Link： http://link.springer.com/article/10.1007/s00348-015-2081-9/fulltext.html

Authorship：Corresponding author   Language：Japanese   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

We visualized two-dimensional wing buffet using fast-framing focusing schlieren method. A supercritical airfoil of NASA SC(2)-0518 was used in this study. Mach number of the freestream was 0.7, and Reynolds number based on the wing code (c) was 5&times;10⁶. Angle of attack without corrections (&alpha;) was changed from 4&deg; to 6&deg;. The present focusing schlieren system had &plusmn;18mm depth of field, which was narrow than spanwise region where the two-dimensionality of the flow field was maintained. 8,345 successive images were captured by high speed CMOS camera with 7,000 frames per second with 20&mu;s exposure time. The present imaging system well captured the shock wave motion on the airfoil. At &alpha; = 4&deg;, the shock wave was positioned at x/c&sim;0.45. The shock wave was slightly fluctuated with a frequency of 585.5Hz. Above &alpha; = 5&deg;, the wing buffet occurred. The buffet frequency increased as increasing &alpha;. The focusing schlieren movies revealed that the shock wave was bifurcated on the wing surface, as the shock traveling upstream. The boundary layer largely separated from the leading oblique foot of the bifurcated shock wave, whereas, the bifurcated shock wave changed into a normal shock and the flow separation disappeared, as the shock traveling downstream. For both cases, many pressure waves propagated from the downstream of the shock wave to the upstream. These waves merged with the shock wave, and seem to drive the shock oscillation.

DOI： 10.2322/jjsass.63.166

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Other Link： http://ousar.lib.okayama-u.ac.jp/56129

Authorship：Lead author   Language：Japanese   Publisher：Japan Society of Mechanical Engineers  

DOI： 10.1299/transjsme.15-00273

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Aeronautics and Astronautics (AIAA)  

DOI： 10.2514/1.B35021

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

DOI： 10.2322/tastj.12.a85

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Language：Japanese   Publisher：Japan Society of Mechanical Engineers  

DOI： 10.1299/transjsme.2014fe0197

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Language：English   Publishing type：Research paper (scientific journal)   Publisher：American Institute of Aeronautics and Astronautics (AIAA)  

DOI： 10.2514/1.B34521

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DOI： 10.2514/1.B34524

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

DOI： 10.2514/1.J051550

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

DOI： 10.2514/1.B34371

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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.

DOI： 10.2514/1.J051588

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DOI： 10.2514/1.B34172

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Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

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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Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

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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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.

DOI： 10.2514/6.2012-5834

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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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Publishing type：Research paper (international conference proceedings)  

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 (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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Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

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

We investigated a supersonic flowfield with transverse injection of air or helium by stereoscopic particle image velocimetry and calculated single-time two-point spatial correlations of velocity fluctuations. For both injection gases, the bow shock wave formed by the injection and its reflection from the upper wall could be seen in the mean velocity distribution. In the velocity fluctuation intensity distribution, we observed vibrating counter-rotating vortex pair and the Reynolds stress was formed along the track of the maximum mean velocity gradient. Correlations showed the scale structure of turbulence and the shapes of the correlation region changed with the selected velocity component. Differences were observed between the results obtained for the two injectant species.

DOI： 10.1299/jfst.7.231

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DOI： 10.6108/KSPE.2012.16.4.050

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

DOI： 10.2514/1.J051067

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Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2011-2343

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

DOI： 10.1016/j.proci.2010.07.057

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Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

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.

DOI： 10.2514/6.2011-5830

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Language：English   Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

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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DOI： 10.2514/1.B34164

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

DOI： 10.1007/s00193-010-0280-0

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Other Link： http://link.springer.com/article/10.1007/s00193-010-0280-0/fulltext.html

Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2010-6644

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DOI： 10.2514/6.2010-6645

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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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DOI： 10.2514/6.2009-7257

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DOI： 10.2514/6.2009-7316

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DOI： 10.2514/6.2009-25

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DOI： 10.1007/978-3-540-85181-3_54

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A small model aerodynamic tests of the combined cycle engine were carried out to evaluate its performance in subsonic and supersonic conditions. In this regime of the flow speed, the combined cycle engine operates as an ejector-jet or ramjet. The nitrogen gas was exhausted as the substitution for the actual rocket gas. In a subsonic condition, there appeared local pressure rise at the kink point of the ramp, increasing the pressure drag. Both wall pressure and the pitot pressure distribution at the exit of the model suggested that the flow structure is "two-layered" ; one is subsonic induced air flow, and the other is the supersonic rocket exhaust. A slit was carved on the topwall inside the isolator section, expecting a better suction performance in the ejector-jet mode. The modification actually had an effect to enhance the lower limit of the rocket pressure at which the choking of the induced air is achieved.

DOI： 10.2322/tstj.7.pa_59

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DOI： 10.2514/6.2008-4503

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DOI： 10.2514/1.35177

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DOI： 10.2322/jjsass.56.163

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DOI： 10.2322/jjsass.56.110

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Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2007-5028

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DOI： 10.2514/6.2007-5405

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DOI： 10.2322/jjsass.55.43

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DOI： 10.2514/6.2006-4867

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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.

DOI： 10.2514/6.iac-06-c4.5.04

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DOI： 10.1016/j.combustflame.2004.10.004

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

DOI： 10.2514/1.7967

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DOI： 10.2514/6.2005-3216

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DOI： 10.2322/jjsass.52.1

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Publishing type：Research paper (international conference proceedings)   Publisher：American Institute of Aeronautics and Astronautics  

DOI： 10.2514/6.2003-7038

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DOI： 10.2322/jjsass.51.403

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

令和3年度宇宙航行の力学シンポジウム（2021年12月20日-21日. オンライン開催)
Symposium on Flight Mechanics and Astrodynamics: 2021 (December 20-21, 2021. Online Meeting)
資料番号: SA6000174007

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Language：English   Publisher：Japan Aerospace Exploration Agency (JAXA)  

Fluid Dynamics Conference / Aerospace Numerical Simulation Symposium 2020 Online (September 28-30, 2020. The Japan Society for Aeronautical and Space Sciences : Japan Aerospace Exploration Agency (JAXA), Online meeting)A fully three-dimensional (3D) RANS-based global stability analysis is performed on wings in the presence of side-walls to study the complex interaction between buffet shock-oscillations, buffet cells and corner separations. A validation of both the nonlinear and linear stability analysis solvers is given for a two-dimensional (2D) incompressible laminar cylinder and for turbulent transonic buffet on a 2D OAT15A airfoil. The numerical setup is based on the experiments conducted at JAXA 0.8 m × 0.45 m high Reynolds number transonic wind tunnel on a 2D Common Research Model (CRM) profile. The CRM profile is extruded in the spanwise direction and flash mounted on lateral walls at its extremities. An unswept and a swept configuration at sweep angle of 10 deg are considered. The effect of the angle of attack (AoA) is studied for both configurations by selecting AoA = 4 and 7 deg. The RANS solutions are compared against the experimental oil-flow visualizations. Despite some differences in terms of size of the corner separation and shock locations, the main flow features are captured. For the unswept case, the results show that the separation in the middle wing section and corners increases with the AoA. Since the flow slows down near the side-walls, the shock is weakened and moves upstream towards the leading edge. When a sweep angle is applied, the flow is distorted by a crossflow velocity component that causes the corner separations to increase or decrease depending on the boundary-layer thickness of the secondary flow created in the wing spanwise direction. Linear global stability calculations carried out on the unswept wing at the lowest AoA show the presence of a 2D oscillatory mode at St approx. = 0.06 that is spatially localized on the shock. Another unstable mode at higher frequencies (St approx. = 0.1) is located near the corner separations and perturbation packets travel downstream along the shear-layer. For the highest AoA, the 2D mode no longer exists and two 3D modes at higher frequencies appear (St approx. = 0.07 and St approx. = 0.2). These modes are organized in spanwise perturbation wavepackets generated from the corner separation and convected downstream towards the wing middle section. In the separated region in the wing mid-section, the perturbations travel upstream from the trailing-edge towards the shock. Future investigations will focus on the effect of the sweep angle and detection of buffet cells.Physical characteristics: Original contains color illustrations

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Language：Japanese   Publisher：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

Self-oscillation of a shock wave, which is called a transonic shock buffet, occurs on a wing surface in a transonic flow. Lee and Crouch explained the self-sustained mechanism by different pressure wave propagations but there is no experimental evidence of which feedback loop is dominant. The principal aim of the present study is to identify the dominant feedback path of pressure waves related to the shock buffet using high-speed focusing schlieren movies. The experimental data showed that the flow fields include complicated phenomena such as a lot of pressure waves with a broad range of frequency. Therefore, we performed dynamic mode decomposition (DMD) analysis to extract dominant phenomena for the buffet. DMD modes whose frequencies corresponded to the shock oscillation frequencies were extracted and investigated. DMD modes revealed that pressure perturbations generated at the shock foot and the perturbation propagated to the height-wise direction along with the shock wave. This feedback pass is similar to that observed by Crouch using the global stability analysis for NACA 0012 airfoil.

DOI： 10.2322/jjsass.70.82

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Language：Japanese   Publisher：The Japan Society of Mechanical Engineers  

DOI： 10.1299/jsmecs.2021.59.07b1

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Authorship：Lead author   Language：Japanese   Publishing type：Article, review, commentary, editorial, etc. (scientific journal)  

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Language：Japanese   Publisher：The Japan Society of Mechanical Engineers  

DOI： 10.1299/jsmecs.2020.58.05b5

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DOI： 10.1299/jsmecs.2020.58.05c2

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DOI： 10.1299/jsmecs.2020.58.05c3

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DOI： 10.1299/jsmecs.2020.58.05a5

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The electrodynamic heat shield is a new type of heat protection system for the atmospheric-entry vehicle, which utilizes the Lorentz force acting on the weakly ionized plasma flow inside the shock layer. We are developing the Viscous Shock Layer (VSL) analysis method for an electrodynamic heat shield for quick estimation of performance, including the effect on the reentry trajectory. In this study, we propose the chemical non-equilibrium VSL equations for the electrodynamic heat shield analysis and compare its result to CFD calculation and theoretical relationship. In proposed VSL equations, momentum and energy equations have additional terms, and species conservation equations have no change. Inside the shock layer, ionization reaction is occurred due to high temperature, and electrical conductivity is changed by the flow condition, such as altitude and flow velocity. VSL results in the case of several altitudes of reentry trajectory agree with CFD results on electrical conductivity estimation. And, the stagnation heat flux diminish rates also agree with the theoretical relationship based on the interaction parameter Q. Therefore, our VSL analysis method can estimate the effect of electrodynamic heat shield depending on the flow condition.

DOI： 10.1299/jsmemecj.2020.s19104

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Language：English   Publisher：American Institute of Aeronautics and Astronautics  

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 were addressed.

DOI： 10.2514/6.2003-7009

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Event date： 2018

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<p>We conducted numerical simulations of the benchmark experiments of the pump sump conducted by Matsui et al. (2006 & 2015) and compared the simulation with the experimental data to investigate the effects of turbulence model, grid density and detection method of the vortices. We determined a threshold of the gas-liquid fraction function of VOF method (α) and the second invariant of velocity gradient tensor (Q2) to detect the air-entraining and submerged vortices by using vorticity, respectively. This method well detected the vortices and well reproduced the experiments for the RANS simulation using SST k-ω model. Large eddy simulation using Smagorinsky model, however, was sensitive to the grid system and difficult to reproduce the experimental vortex structures even for the finest grid system having 3.7 mil-lion cells.</p>

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Event date： 2018

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<p>In semiconductor production, efficiency and precision of the cleaning process are very important. In this study, we handle "Single Wafer Cleaner" that is one of the washing methods in the way of blowing air to a wafer with a turning disk at a high speed. However, the flow in the device becomes turbulent, which causes the problem that water drops and dusts re-attach to the wafer surface. Therefore, it is necessary to understand the flow in the device exactly to solve this problem. In previous studies, we experimented to establish the method of understanding the flow in the semiconductor wafer cleaner without a covering cup that is used for liquid splash preventing. In this study, numerical calculations were performed to observe the flow in the device to study influence of covering cup and find a suitable turbulence model to simulate the experiment.</p>

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Event date： 2016

Language：English  

© 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.

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Event date： 2015.2.27

Language：Japanese  

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Event date： 2015

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Event date： 2015

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Event date： 2015

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Event date： 2014.9.7

Language：Japanese  

In a single wafer spin cleaner, the wafer is rotated at high speed to spin off ultrapure water and clean air is blown perpendicular onto the surface to dry the wafer. Interaction between the rotating wafer and blown air generates vortices and the recirculation flow due to the vortices may cause reattachment of contaminants on the wafer surface. Therefore, we experimentally studied vortical structures and the conditions of their formation in a modeled cleaner by using particle image velocimetry (PIV). The model was a 330 mm diameter rotating disk in a 520 mm diameter cylindrical stationary housing, and laminar blown air came from the housing inlet onto the rotating disk. The angular speed of the rotating disk was changed up to 500 rpm and the blown air rate was 0.5 m3/min. Our PIV data show that turbulent kinetic energy near the housing wall with a vortex is as large as that around the rotating disk around at SOOrpm. This vortex may cause reattachment of particles onto the wafer. The vortices generated in the experimental apparatus have three-dimensional structures regardless of their size.

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Event date： 2014

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Event date： 2014

Language：Japanese  

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Event date： 2013.9.8

Language：Japanese  

The cleaning process of a silicon wafer is important in the manufacturing of semiconductors, which determines the quality of the final products crucially. In the so-called single wafer cleaning, a wafer washed with cleaning solution and rinse liquid is dried up by rotating at a high speed. We made an experimental apparatus which has a similar structure to the single wafer cleaning chamber, and measured the velocity field in the chamber by the particle image velocimetry(PI V) method.

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Event date： 2012.3.13

Language：Japanese  

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Event date： 2010

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Event date： 2010

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Event date： 2010

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Event date： 2010

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Event date： 2010

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Event date： 2010

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Event date： 2009

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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Event date： 2009

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Event date： 2009

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.

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Event date： 2009

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Event date： 2009

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.

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Event date： 2009

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Event date： 2009

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Event date： 2009

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Event date： 2008

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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Event date： 2007

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Event date： 2007

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.

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Event date： 2007

Language：Japanese  

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 stinger shaped injector showed the best penetration performance among all injectors. There was a difference between the experiment and the numerical analysis in evaluation of the injectant concentration. However, the penetration height resulted from the numerical analysis showed good agreement with that from the experiment. It is believed that the reduction of the disturbance of the plume to the crossflow leads to the increment in the penetration.

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Event date： 2007

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Event date： 2007

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Event date： 2007

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Event date： 2006

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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Event date： 2004.11

Language：English  

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Event date： 2003

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.

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