Supersonic air-intake for Mach number higher than 2.5 is being investigated with experimental, analytical and computational 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 designed in two-dimensional mixed compression type with multi-shock system and tested in NAL’s 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.
Experimental Study on Axisymmetric Air Intake for the Supersonic Transport.
