National Aerospace Laboratory Structural Mechanics Division
National Aerospace Laboratory Aeroengine Division
National Aerospace Laboratory Aeroengine Division
National Aerospace Laboratory Aeroengine Division
National Aerospace Laboratory Aeroengine Division
Science University of Tokyo Department of Mechanical Engineering, Faculty of Engineering
An experimental investigation of a three-dimensional flow field with a swept shock wave and a turbulent boundary layer interaction was made in NAL's 1 m x 1 m Mach 4 supersonic wind tunnel. The emphasis of this study is placed on understanding the behaviors of the flow interaction with Mach number variation and the effects of bleed on the surface flow. A shock generator of 15 deg wedge placed normally on the flat plate was applied to generate an oblique shock wave. The experiments were conducted at Mach numbers of 3.25, 3.11, and 2.63. The conical coordinate system was adopted to analyze the flow. Air-bleed was also tested utilizing rows of bleed holes located at the foot print of the swept shock wave on the flat plate surface. It is generally accepted that the size and the strength of the separation vortex generated along the shock wave on the surface become wider and stronger as the Mach number increases. The corner flow around the leading edge of the shock generator was also clarified. As for the air-bleed, the suppression effects of the bleed on the separation vortex were also clarified at several bleed locations. The bleed upstream of the shock wave foot print was effectively reduced the size of the vortex, while the downstream bleed suppressed the vortex strength. Therefore, it is concluded that the combination of the upstream and downstream bleed effectively suppresses the interaction between the shock wave and the boundary layer.
斜め衝撃波/乱流境界層の3次元干渉場の研究:流れ構造と抽気位置の影響
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