Numerical investigation on the shock wave transition in a three-dimensional scramjet isolator

Pseudo-shock waves and their interactions in high-speed intakes

A measurement campaign has been carried out to investigate the viscous flow effects in a fixed geometry two-dimensional scramjet inlet at Mach 7. The tests have been performed in a hypersonic blowdown wind-tunnel facility. The existing boundary layer separation due to shock interaction in the inlet throat can lead to inlet unstart. Therefore, an optional passive boundary layer bleed has been integrated at the throat. The passive bleed reduces the lip shock induced separation bubble in the throat significantly. The obtained experimental wall pressure distribution and pitot pressure profiles are discussed and compared to 2-D and 3-D computational fluid dynamics calculations. To investigate the compression behavior, different backpressure ratios have been applied to the isolator flow and the effects on the internal flow structure are analyzed by means of shadowgraph pictures, static pressure distribution, and pitot pressure profiles at the exit of the isolator. The heat transfer coefficient to the inner side wall of the isolator is calculated using the time-dependent surface temperature, measured with an infrared system. The resulting 2-D Stanton number distributions are presented for different configurations and different backpressure ratios. In addition, three different Reynolds numbers have been used to investigate the effect on the separation bubble located in the corner of the first and second ramp transitions and the resulting change in ramp heat flux.

Investigation of Two-Dimensional Scramjet Inlet Flowfield at Mach 7

To study the influence of viscous effects on the performance and, especially the internal flowfield, of scramjet engines, an experimental study has been carried out on a fixed geometry scramjet inlet operating at Mach 6. The tests have been performed using a hypersonic blowdown wind-tunnel facility. To reduce the risk of inlet unstart, a passive boundary-layer bleed has been integrated at the throat and successfully tested. The passive bleed reduces the lip shock-induced separation bubble on the ramp significantly. The obtained experimental results are discussed and compared with computational fluid dynamics calculations. To investigate the compression behavior, different backpressures have been applied and the effects on the internal flow structure are analyzed by means of schlieren change in pitot pressure and calculated Mach number profile base don the geometrical shape of the internal flowpath. The heat transfer coefrficient to the inner sidewall of the isolator is calculated using the time-dependent surface temperature, measured with an infrared system. The variation of the Stanton number based on different flow phenomena could be resolved successfully.

Internal Flowfield Investigation of a Hypersonic Inlet at Mach 6 with Bleed

The current experimental study at Mach 7 was carried out to investigate the viscous flow effects in a scramjet inlet with a fixed geometry operating at Mach 7.5-8. To reduce the risk of inlet unstart, an optional passive boundary-layer bleed was integrated at the throat and tested for different bleed dimensions. The passive bleed significantly reduces the lip-shock-induced separation bubble on the ramp. But the correct location of the bleed gap is important for the reduction of the separation bubble. To investigate the compression behavior, different backpressures were applied and the effects on the internal flow structure were analyzed by means of static pressure along the isolator and pitot pressure profiles at the exit of the isolator. Because the reference configuration has a very small internal contraction ratio to guarantee self-starting, the internal compression ratio was increased in three steps by means of additional sidewall compression along the internal flowpath. This includes internal contraction ratios above the starting limit. This measure led to a significant increase of the static pressure ratio at the end of the isolator. The complex flow structures are explained by means of detailed static and pitot pressure plots.

Experimental Investigation of a Two-Dimensional and a Three-Dimensional Scramjet Inlet at Mach 7

In this paper, the outline of the Mach 4 supersonic wind tunnel for the investigation of the supersonic internal flows in ducts was firstly described. Secondly, the location, structure and characteristics of the Mach 2 and Mach 4 pseudo-shock waves in a square duct were investigated by color schlieren photographs and duct wall pressure fluctuation measurements. Finally, the wall shear stress distributions on the side, top and bottom walls of the square duct with the Mach 4 pseudo-shock wave were investigated qualitatively by the shear stress-sensitive liquid crystal visualization method. The side wall boundary layer separation region under the first shock is narrow near the top wall, while the side wall boundary layer separation region under the first shock is very wide near the bottom wall.

Study on shock wave and turbulent boundary layer interactions in a square duct at Mach 2 and 4

The structures and characteristics of λ-shaped and X-shaped pseudo-shocks in a square duct are investigated through numerical simulations and experiments at Mach 2 and Mach 4, respectively. The experiments were carried out in a pressure-vacuum supersonic wind tunnel with a test cross section of 80×80 mm2. Numerical simulations were carried out using the Harten-Yee second-order TVD scheme and the Baldwin-Lomax turbulence model. The Reynolds numbers for the Mach 2 and 4 cases were Re∞=2.53×107 and Re∞=2.36×107, respectively, and the flow confinement was δ∞⁄h=0.35 for both cases. The computational results for the Mach 2 pseudo-shock wave are in good agreement with the experimental results. Based on this agreement, the flow quantities, which are very difficult to obtain experimentally, were analyzed by numerical simulation. Although several differences were found between the computational results and experiments in the case of Mach 4 due to asymmetric characteristics in experiment which could not be reproduced in numerical simulation, the computational results are valuable for understanding this complex asymmetric phenomenon.

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Numerical and Experimental Investigations on Mach 2 and 4 Pseudo-Shock Waves in a Square Duct

This paper presents an investigation on the structure and characteristics of the multiple shock wave/turbulent boundary layer interaction in a Mach 2 supersonic square duct by numerical simulation and experiment. The numerical simulation is carried out with the Harten-Yee’s second-order accuracy TVD scheme and the Baldwin- Lomax’s turbulence model. The flow conditions are: free stream Mach number M∞ = 2.0, unit Reynolds number Re∞/m = 2.5xl07m−1, and the flow confinement is δ∞/h = 0.25. Good agreements between the numerical analysis and the experiment for the shape of the shock train and wall pressure distribution along the duct are obtained. Based on these agreements, the flow quantities, which are difficult to obtain by experiment, are analyzed by numerical simulation.

Akira Tojo

Papers

Study on shock wave and turbulent boundary layer interactions in a square duct at Mach 2 and 4

Numerical and Experimental Investigations on Mach 2 and 4 Pseudo-Shock Waves in a Square Duct

Numerical and experimental study of the Mach 2 pseudo-shock wave in a supersonic duct

Study on Internal Supersonic Flows with Pseudo shock Wave Using Liquid Crystal Flow Visualization Method