• 대한기계학회논문집
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• 제19권9호
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• pp.2283-2296
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• 1995

Experiments of normal shock wave/turbulent boundary layer interaction were conducted in a supersonic diffuser. The flow Mach number just upstream of the normal shock wave was in the range of 1.10 to 1.70 and Reynolds number based upon the turbulent boundary layer thickness was varied in the range of 2.2*10$^{[-994]}$ -4.4*10$^{[-994]}$ . The wall pressures in streamwise and spanwise directions were measured for two test cases, in which the turbulent boundary layer thickness incoming into the supersonic diffuser was changed. The results show that the interactions of normal shock wave with turbulent boundary layer in the supersonic diffuser can be divided into three patterns, i.e., transonic interaction, weak interaction and strong interaction, depending on Mach number. The weak interactions generate the post-shock expansion which its strength is strong as the Mach number increases and the strong interactions form the pseudo-shock waves. From the spanwise measurements of wall pressure, it is known that if the flow Mach number is low, the interacting flow fields essentially appear two-dimensional, but they have an apparent 3-dimensionality for the higher Mach numbers.

• 한국추진공학회지
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• 제25권6호
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• pp.36-44
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• 2021

스크램제트 엔진 비행시험체의 굽어진 중앙동체 내부 유로에서 발생하는 충격파-경계층 상호작용에 대한 수치해석을 수행하였다. 수치해석에는 압축성 Raynolds Averaged Navier Stokes(RANS) 방정식에 난류모델 k-ω SST을 사용하였다. 대표적으로 노즐 윗 벽면의 박리기포, 오목한 충격파와 경계층의 상호작용, 모서리의 충격파-충격파 상호작용이 포착되었다. 해석 결과는 굽어진 내부 유로의 충격파-경계층 상호작용을 가시화하여 이해를 높이고 설계 유의점을 제시하였다.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2004년도 제22회 춘계학술대회논문집
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• pp.426-432
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• 2004

Large-Eddy Simulation (LES) is applied for the simulation of compressible flat plate boundary with Reynolds number up to 5 X 10$^{5}$ . Numerical examples include shock/boundary layer interaction and boundary layer transition, aiming future application to the analysis of transonic fan/compressor cascades. The present LES code uses hybrid com-pact/WENO scheme for the spatial discretization and compact diagonalized implicit scheme for the time integration. The present code successfully predicted the bypass transition of subsonic boundary layer. As for supersonic turbulent boundary layer, mean and fluctuation velocity of the attached boundary, as well as the evolution of the friction coefficient and the displacement thickness both upstream and downstream of the separation region are all in good agreement with experiment. The separation point also agreed with the experiment. In the simulation of the shock/laminar boundary layer interaction, the dependence of the transition upon the shock strength is reproduced qualitatively, but the extent of the separation region is overpredicted. These numerical examples show that LES can predict the behavior of boundary layer including transition and shock interaction, which are hardly managed by the conventional Reynolds-averaged Navier-Stokes approach, although there needs to be more effort before achieving quantitative agreement.

• 한국전산유체공학회지
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• 제18권2호
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• pp.56-65
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• 2013

Two compression ramp problems and an impinging shock problem are computed to investigate influence of turbulence models and eddy viscosity on the shock-wave / boundary layer interaction. A Navier-Stokes boundary layer generation code was applied to the generation of inflow boundary conditions. Computational results are validated well with the experimental data and effects of turbulence models are investigated. It is shown that the behavior of turbulence (eddy) viscosity directly affects both the extent of the separation and shock-wave positions over the separation.

• 한국산업융합학회 논문집
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• 제2권2호
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• pp.115-124
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• 1999

The interaction of weak normal shock wave with turbulent boundary layer in a supersonic nozzle was investigated experimentally by wall static pressure measurements and by schlieren optical observations. The lime-mean flow in the interaction region was classified into four patterns according to the ratio of the pressure $p_k$ at the first kink point in the pressure distribution of the interaction region to the pressure $p_1$ just upstream of the shock. It is shown for any flow pattern that the wall static pressure rise near the shock foot can be described by the "free interaction" which is defined by Chapman et al. The ratio of the triple point height $h_t$ of the bifurcated shock to the undisturbed boundary layer thickness ${\delta}_1$ upstream of the interaction increases with the upstream Mach number $M_1$, and for a fixed $M_1$, the normalized triple point height $h_t/{\delta}_1$ decreases with increasing ${\delta}_1/h$, where h is the duct half-height.

• 대한기계학회논문집B
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• 제21권10호
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• pp.1273-1283
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• 1997

Experiments of shock wave/turbulent boundary layer interaction were conducted by using a supersonic wind tunnel. Nominal Mach number was varied in the range of 1.6 to 3.0 by means of different nozzles. The objective of the present study is to investigate the effects of boundary layer flow bleed on the interaction flow field in a straight tube. Two-dimensional slits were installed on the tube walls to bleed the turbulent boundary layer flows. The bleed flows were measured by an orifice. The ratio of the bleed mass flow to main mass flow was controlled within the range of 11 per cent. The wall pressures were measured by the flush mounted transducers and Schlieren optical observations were made for almost all of the experiments. The results show that the boundary layer flow bleed reduces the multiple shock waves to a strong normal shock wave. For the design Mach number of 1.6, it was found that the normal shock wave at the position of the silt was resulted from the main flow choking due to the suction of the boundary layer flow.

• 한국연소학회지
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• 제1권1호
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• pp.83-91
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• 1996

A numerical study is carried out to examine the ignition and propagation process of detonation wave in SCRam-accelerator operating in superdetonative mode. The time accurate solution of Reynolds averaged Navier-Stokes equations for chemically reacting flow is obtained by using the fully implicit numerical method and the higher order upwind scheme. As a result, it is clarified that the ignition process has its origin to the hot temperature region caused by shock-boundary layer interaction at the shoulder of projectile. After the ignition, the oblique detonation wave is generated and propagates toward the inlet while constructing complex shock-shock interaction and shock-boundary layer interaction. Finally, a standing oblique detonation wave is formed at the conical ramp.

• 한국전산유체공학회지
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• 제10권3호
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• pp.19-26
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• 2005

The purpose of this paper is the study on the characteristics of an inlet system with shock/boundary layer interactions by using various types of bumps which are substituted for the conventional bleeding system in supersonic inlet. in this study a comprehensive numerical analysis has been performed to understand the three-dimensional flow field including shock/boundary layer interaction and growth of turbulent boundary layer that might occur around a three-dimensional bump in a supersonic inlet. The characteristics of boundary layer seen in the current numerical simulations indicate the potential capability of a three-dimensional bump to control shock/boundary layer interaction in supersonic inlets.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2005년도 춘계 학술대회논문집
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• pp.213-218
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• 2005

The purpose of this study is the characteristics of an innovative inlet system with shock/boundary layer interactions by using various types of bumps which are substituted for the conventional bleeding system in supersonic inlet. This study performs a comprehensive numerical effort that be directed at better understanding the three-dimensional flowfield includes shock/boundary layer interaction and growth of turbulent boundary layer that occur around a three-dimensional bump in a supersonic inlet. The characteristics of boundary layer seen in the current numerical simulations indicates the potential capability of the three-dimensional bump to control shock/boundary layer interaction in supersonic inlets.

• Journal of Mechanical Science and Technology
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• 제15권3호
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• pp.366-373
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• 2001

A numerical analysis of shock wave/boundary layer interaction in transonic/supersonic axial flow compressor cascade has been performed by using a characteristics upwind Navier-Stokes method with various turbulence models. Two equation turbulence models were applied to transonic/supersonic flows over a NACA 0012 airfoil. The results are superion to those from an algebraic turbulence model. High order TVD schemes predicted shock wave/boundary layer interactions reasonably well. However, the prediction of SWBLI depends more on turbulence models than high order schemes. In a supersonic axial flow cascade at M=1.59 and exit/inlet static pressure ratio of 2.21, k-$\omega$ and Shear Stress Transport (SST) models were numerically stables. However, the k-$\omega$ model predicted thicker shock waves in the flow passage. Losses due to shock/shock and shock/boundary layer interactions in transonic/supersonic compressor flowfields can be higher losses than viscous losses due to flow separation and viscous dissipation.