Research on Shockwave/Boundary Layer Interactions Induced by Double Compression Corners Under Hypersonic Quiet and Noise Inflow Conditions

The problem of shock wave/boundary layer interaction induced by compression corners widely exists in the external and internal flows of various supersonic/hypersonic aircraft. In practical engineering applications, multistage continuous compression is often used in the fin/rudder structure, while in internal flow, multistage compression schemes are usually employed at the inlet to enhance total pressure recovery; therefore, it is necessary to investigate the characteristics of multistage compression corner shockwave/boundary layer interactions. In basic research, it is usually simplified as the double compression corner shockwave/boundary layer interaction issue. In this paper, an experimental study of hypersonic shock/boundary layer interaction characteristics is conducted under quiet and noise inflow conditions, respectively, for the double compression corner model. Using high-speed Schlieren, the typical structure of shockwave/shockwave interaction and shockwave/boundary layer interaction above the corner is explored under both quiet and noisy incoming flow conditions. Then, based on gray average, root-mean-square analysis, Fast Fourier transform, proper orthogonal decomposition, and dynamic mode decomposition methods, the time-average and unsteady characteristics of the double compression corner configuration-induced separation were studied, and a comparative analysis was conducted. The difference law between wind tunnel noise level and interaction characteristics was summarized. Finally, the characteristic length and spectral characteristics of unstable waves that dominated the stability of the plate boundary layer were studied. The formation mechanism of separation is discussed, which provides technical support for the internal and external aerodynamic design and targeted optimization of hypersonic vehicles.