Abstract
The flow of high-speed air in ducts may result in the occurrence of multiple shock-wave/boundary-layer interactions. Understanding the consequences of such interactions, which may include distortion of the velocity field, enhanced turbulence production, and flow separation, is of great importance in understanding the operating limits and performance of a number of systems, for example, the high-speed intake of an air-breathing missile. In this paper, the results of a computational study of multiple shock-wave/boundary-layer interactions occurring within a high-speed intake are presented. All of the results were obtained using the in-house computational fluid dynamics solver of Glasgow University, HMB3. First simulations of a Mach M=1.61 multiple shock-wave/boundary-layer interaction in a rectangular duct were performed. The M=1.61 case, for which experimental data is available, was used to establish a robust numerical approach, particularly with respect to initial and boundary conditions. A number of turbulence modelling strategies were also investigated. The results suggest that Reynolds-stress-based turbulence models are better suited than linear eddy-viscosity models. This is attributed to better handling of complex strain, in particular modelling of the corner separation. The corner separations affect the separation at the centre of the domain which in turn alters the structure of the initial shock and the subsequent interaction. Having established a robust numerical approach, the results of a parametric study investigating the effect of Mach number, Reynolds number, and confinement on the baseline solution are then presented. Performance metrics are defined to help characterize the effect of the interactions. The results suggest that reduced flow confinement is beneficial for higher-pressure recovery.
Highlights
The key to the design of supersonic intakes is compressing a large volume of air with minimal total pressure loss and flow distortion while avoiding the possibility of unstart overCommunicated by H
Multiple shock-wave/boundary-layer interactions (SWBLIs) are often referred to as shock trains or pseudo-shocks. These terms are used, at times, interchangeably, the term shock train refers to the series of shocks and the pseudo-shock term refers to the entire region of pressure rise [1]
The model slightly underpredicts the wall pressure at the beginning of the interaction which is attributed to the larger separation at the foot of the initial shock
Summary
The key to the design of supersonic intakes is compressing a large volume of air with minimal total pressure loss and flow distortion while avoiding the possibility of unstart overCommunicated by H. The key to the design of supersonic intakes is compressing a large volume of air with minimal total pressure loss and flow distortion while avoiding the possibility of unstart over. This design objective can be achieved in many different ways, for example, using an isentropic compression in a long duct, but the most reasonable, for a practical air vehicle, appears to be splitting the overall compression over a sequence of inclined shock waves. The overall efficiency of the intake depends fundamentally on the nature of the shocks, their interaction with the intake boundary layer, and the resultant flow. For most flight conditions, formation of multiple shock-wave/boundary-layer interactions (SWBLIs) inside the intake is expected. We have the pseudo-shock region, that part of the domain over which the (a)
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