Shigemi Shindo

The Effect of the Bleed Hole Arrangement on Suppression of Swept Shock Wave/Turbulent Boundary Layer Interactions

The paper describes the experiments on the swept shock/boundary layer interactions with and without boundary layer bleed, as the fundamental research issue in the development of a supersonic air intake. The experiments were conducted in a supersonic wind tunnel with a cross section of 1 m by 1 m of the National Aerospace Laboratory. The emphasis was placed on understanding the flow structures of the interaction, using a shock generator with a 15 deg. wedge at Mach numbers of 3.25, 3.11 and 2.63. Attention was also focused on suppression of the interaction by a bleed system of hole arrangement. Three cases of the bleed system, where the bleed portions were located upstream, downstream, and in both regions of the swept shock wave, were employed. It was found by both Pitot-pressure rake and surface pressure survey methods, in addition to the vapor screen visualization technique, that the interaction becomes stronger with increasing Mach number. The arrangement of the bleed portion covering both upstream and downstream of the shock position is effective to suppress swept shock wave/boundary layer interactions.

Experimental Study of Mixed Compression Air-Intake for Hypersonic Airbreathing Engines

Supersonic air-intake for Mach number higher than 2.5 is being investigated with experimental, analytical and computa- tional methods. The study is performed in a part of the joint research program led by National Aerospace Laboratory (NAL) on the hypersonic airbreathing turbo-engines with subsonic ram combustion. The wind tunnel models are de- signed in two-dimensional mixed compression type with multi- shock system and tested in NALs Mach 4 supersonic wind tunnel. Pressure measurements and flow visualization by schlieren method, oil-flow and vapor screen techniques are being done. Here, the test results of Mach 4 and Mach 5 models are discussed. The Mach 4 model is fixed geometry with 5-shock system and the Mach 5 one is variable geometry with 6-shock and an isentropic compression surface. An expansion fore- plate was installed at the Mach 5 model inlet to accelerate the air-speed at the entry. The bleed systems at throat, ramp and cowl are adopted and evaluated in terms of pressure recovery and stability. Importance of establishment of the internal shock wave system, reduction of upstream Mach number of terminal shock wave and suppression of flow separation at diffusers are found. It is also found that ramp bleed is effective to confirm intake start and to minimize shock/boundary layer interaction.

Flow visualization of superasonic air-intake flow with wind tunnel tests. 2. Boundary layrer visualization with vapor-screen method.

Supersonic air-intake is one of the most important components to realize the supersonic/hypersonic air-breathing engines. A mach 3 intake was designed and tested with the supersonic wind tunnel. In this study, several flow visualization methods such as schlieren photograph, oil-flow and vapor screen methods, were applied. Here, the results with vapor-screen method using acetone and pentane were presented. The boundary layer flow changes, associated with the back pressure controlled by the flow plug, were visualized clearly. The secondary flows and vortices were also caught.