• International Journal of Aeronautical and Space Sciences
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• v.18 no.1
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• pp.1-7
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• 2017

To predict the transonic buffet onset for a supercritical airfoil with shock-boundary layer interactions, a practical steady approach has been proposed. In this study, it is assumed that the airfoil flow is steady even when buffet onset occurs. Steady Navier-Stokes computations are performed on the supercritical airfoil. Using the aerodynamic parameters calculated from Navier-Stokes solver, various steady approaches for predicting buffet onset are discussed. Among the various steady approaches considered in this study, Thomas' criterion based on Navier-Stokes computation has shown to be the most appropriate indicator of identifying the buffet onset for a supercritical airfoil with shock-boundary layer interactions. Good agreements have been obtained compared with the results of unsteady transonic wind tunnel tests. The present method is shown to be reliable and useful for transonic buffet onset for a supercritical airfoil with shock-boundary layer interactions in terms of practical engineering viewpoint.

• Proceedings of the KSME Conference
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• 2002.08a
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• pp.159-160
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• 2002

New S-shaped aeroelastic mesoflaps are utilized to control normal shock/boundary-layer interactions. New generation of the mesoflaps is designed f3r a better rigidness and a good flow uniformity across the ulteractions. ,Major advantages of the mesoflap system can be a better total pressure recovery downstream of the interactions due to the lambda shock structure over the flap system, and a rehabilitation of the thickened boundary layer due to bleeding through a cavity underneath the flap system. Skin friction has been measured downstream of the interactions, using the laser interferometer skin friction (LISF) meter, which optically detects the rate of thinning of an oil film applied to the test surface. Various flap-thicknesses of the S-shaped mesoflap arrays are tested, and the results are compared to the solid-wall reference case. Overall, not much difference in the level of skin friction is noticed for the S-shaped flap arrays of various thicknesses, and its level is lower than the skin friction downstream of the solid-wall interaction

• Journal of Mechanical Science and Technology
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• v.17 no.12
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• pp.2116-2124
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• 2003

An experimental study providing additional knowledge of quasi-conical symmetry in swept shock wave/turbulent boundary-layer interactions is described. When a turbulent boundary layer on the flat plate is subjected to interact with a swept planar shock wave, the interaction flowfield far from fin leading edge has a nature of conical symmetry, which topological features of the interaction flow appear to emanate from a virtual conical origin. Surface streakline patterns obtained from the kerosene-lampblack tracings have been utilized to obtain representative surface features of the flow, including the location of the virtual conical origin. The scaling law for the sharp-fin interactions suggested by previous investigators has been reexamined for different freestream Mach numbers. It is noticed that the scaling law reasonably agrees with the present experimental data, however, that the law is not appropriate to estimate the location of the virtual conical origin. Further knowledge of the correlation for the virtual conical origin has thus been proposed.

• Transactions of the Korean Society of Mechanical Engineers
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• v.19 no.9
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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.

• International Journal of Fluid Machinery and Systems
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• v.1 no.1
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• pp.24-32
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• 2008

Bleeding away a part of the boundary layer next to the wall is an effective method for controlling boundary-layer distortions from incident shock waves or curvature in geometry. When the boundary-layer flow is supersonic, the physics of bleeding with and without an incident shock wave is more complicated than just the removal of lower momentum fluid next to the wall. This paper reviews CFD studies of shock-wave/boundary-layer interactions on a flat plate with bleed into a plenum through a single hole, three holes in tandem, and four rows of staggered holes in which the simulation resolves not just the flow above the plate, but also the flow through each bleed hole and the plenum. The focus is on understanding the nature of the bleed process.

• 한국전산유체공학회:학술대회논문집
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• 2000.05a
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• pp.51-59
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• 2000

The Interactions of shock wave with turbulent boundary layers in high-speed flows cause complex flowfields which result in increased adverse pressure gradients, skin friction and temperature. Accurate and reliable prediction of such phenomena is needed in designing high-speed propulsion systems. Such analyses of the complex flowfields require sophisticated numerical scheme that can resolve interactions between shock wave and boundary layers accurately. Therefore the purpose of the present. article is to introduce an accurate and efficient mixed explicit-implicit generalized Galerkin finite element method. To demonstrate the validity of the theory and numerical procedure, several benchmark cases are investigated.

• Bulletin of the Korean Space Science Society
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• 2004.10b
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• pp.328-331
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• 2004

Shock wave boundary layer interactions can make flows around a vehicle be very high pressure and temperature due to pass shock waves in small areas of the hypersonic vehicle. These phenomena can affect a critical problem in the design of hypersonic vehicles. To research the effect of shock wave boundary layer interactions, comer flows were studied in this paper using numerical studies with the NSMB (Navier-Stokes Multi Block) solver and then comparing corresponding numerical results with experimental data of the Huston High Speed Flow Field Workshop II. The mach number of flows is 12.3 in comer flows. The comparison with the computational result is presented based on diverse numerical schemes. Good agreement is obtained.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.25 no.3
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• pp.322-329
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• 2001

A study of the shock wave turbulent boundary layer interaction is presented. The focus of the study is the interactions of the shock waves with the turbulent boundary layer on the falt plate. Three examples are investigated. The computations are performed, using mixed explicit-implicit generalized Galerkin finite element method. The linear equations at each time step are solved by a preconditioned GMRES algorithm. Numerical results indicate that the implicit scheme converges to the asymptotic steady state much faster than the explicit counterpart. The computed surface pressures and skin friction coefficients display good agreement with experimental data. The flowfield manifests a complex shock wave system and a pair of counter-rotating vortices.

• Journal of Mechanical Science and Technology
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• v.15 no.3
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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.

• Journal of computational fluids engineering
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• v.10 no.3 s.30
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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.