• 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.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2004년도 추계 학술대회논문집
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• pp.43-46
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• 2004

Numerical analysis for pressure drag on boattail afterbodies have been studied by Mach number, boattail angle and length ratio of body diameter and base diameter using CFD-FASTRAN that the commercial external flow CFD code. The numerical results have been compared with the experimental data that have been shown pressure drag reduction and supersonic turbulent flow characteristics for boattail afterbodies. And the prediction equation tot boattail base drag has been made by the numerical results about Mach number and boattail configuration parameters.

• 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.

• 대한기계학회논문집B
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• 제22권9호
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• pp.1229-1237
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• 1998

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 suction on normal shock-wave oscillations caused by shock wave/boundary layer interaction in a straight duct. Two-dimensional slits were installed on the top and bottom walls of the duct to bleed 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 below the range of 11 per cent. Time-mean and fluctuating wall pressures were measured, and Schlieren optical observations were made to investigate time-mean flow field. Time variations in the shock wave displacement were obtained by a high-speed camera system. The results show that boundary layer suction by slits considerably reduce shock-wave oscillations. For the design Mach number of 2.3, the maximum amplitude of the oscillating shock-wave reduces by about 75% compared with the case of no slit for boundary layer suction.

• 한국추진공학회지
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• 제6권1호
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• pp.63-70
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• 2002

본 연구에서는 초음속 및 음속 마이크로 제트 유동의 특성을 파악하기 위해서 2차원 축대칭 압축성 Navier-Stokes 방정식을 이용한 수치해석을 수행하였다. 수치해석에서는 여러 형태의 난류모델을 적용하여 실험결과와 비교하였으며, 부족팽창과 과팽창 상태를 얻기 위하여 노즐 출구의 압력비 $P_b/P_e$를 0.2~l.25로 변화시켰다. 또, 레이놀즈 수 Re를 약 600~40000까지의 범위에서 변화시켜, $P_b/P_e$와 Re가 마이크로 제트 유동장에 미치는 영향을 조사하였다. 본 연구로부터 층류 제트의 특성은 난류 제트와는 달리 레이놀즈 수의 의존성이 크다는 것을 알았다.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2010년 춘계학술대회논문집
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• pp.21-27
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• 2010

The flow characteristics on a supersonic inlet with bleeding system by changing angles of attack and channel length conditions are studied by computational 3D turbulent flow analysis. A compressible upwind flux difference splitting Navier-Stokes method with $k-{\omega}$ turbulence model is used to analysis the inlet flowfield. More non-uniform flowfields are shown at the AIP when angle of attack becomes bigger and bigger. These non-uniform flowfield works the performance aggravating factors of the supersonic engine. Non-uniform flowfield by changing channel length at the various angle of attack are investigated.

• 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.

• 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.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 1995년도 제5회 학술강연회논문집
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• pp.69-83
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• 1995

A great number of experimental data indicating shock-induced separation(SIS) in internal or external supersonic flows were reviewed to make clear the mechanism of SIS and to present the criterion of turbulent boundary layer separation. The interesting conclusions were obtained for the considerably wide range of flow geometries that the incipient separation is almost independent of the flow geometries, and that it is relatively unaffected by changes in gas specific heat, and boundary layer Reynolds number, Furthermore, the pressure rise necessary to separate boundary layer in external flows was found to be applicable to SIS in overexpanded propulsion nozzles. This is due to the fact that the SIS phenomenon caused by the interaction between shock waves and turbulent boundary layers is processed through a supersonic deceleration. This is, the SIS in almost all of interacting flow fields is governed by the concept of free interaction, and criterion of SIS is only a Function of upstream Mach number.

• Advances in aircraft and spacecraft science
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• 제6권1호
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• pp.19-30
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• 2019

Two- and three-dimensional turbulent airflows in a 9-degrees-bent channel are studied numerically. The inner surfaces of upper and lower walls are parallel to each other upstream and downstream of the bend section. The free stream is supersonic, whereas the flow at the channel exit is either supersonic or subsonic depending on the given backpressure. Solutions of the Reynolds-averaged Navier-Stokes equations are obtained with a finite-volume solver ANSYS CFX. The solutions reveal instability of formed shock waves and a flow hysteresis in considerable bands of the free-stream Mach number at zero and negative angles of attack. The instability is caused by an interaction of shocks with the expansion flow formed over the convex bend of lower wall.