Research Of Supersonic Shock/Turbulent Boundary Layer Interaction Based On Large Eddy Simulation

Shock wave/boundary layer interaction（SWBLI）is one of the most noticeable physical phenomena in supersonic/hypersonic vehicles.Its inherent unsteadiness has an inestimable effect on aerodynamic performance and material life of aircraft.At present,the complex flow mechanism in SWBLI is still not fully understood.In order to have a deeper understanding of SWBLI,it is necessary to combine high-precision CFD simulation technology closely with advanced physical experiments.First and foremost,the governing equations of large eddy simulation in complete form is established through a series of transformations of Navier-Stokes equations in a general coordinate system.The unclosed terms generated after filtering are processed by dynamic subgrid-scale model.WENO_ZQ scheme with fifth order accuracy is adopted to discretize inviscid flux in order to maintain high accuracy in smooth region and stability near discontinuity.The 3^rd-order Runge-Kutta scheme is used for time integration.Recycling/Rescaling Method（RRM）was used as the inflow turbulence generation technique.By reviewing the theoretical basis of RRM for incompressible and compressible flows,a recycling method for supersonic flow is established and validated on a flat plate turbulent boundary layer.On the basis of the turbulent boundary layer generated above,several problems in reflected SWBLI flow field are studied emphatically.The results show that turbulent fluctuations generated by SWBLI are mainly distributed in the free shear layer and the reattached boundary layer.The distribution of wall pressure fluctuation versus streamwise location shows characteristic of single peak,which is different from the double peak structure found in the compression corner.The existence of bow-shaped reflected shock in unsteady transient flow field is confirmed,and its three-dimensional structure is analyzed in detail.Based on the inflow turbulence generation technique mentioned above,an in-depth study of supersonic compression corner flow is carried out.Structural changes of shock system and boundary layer in the interaction region were studied by flow visualization methods such as vortex recognition and numerical schlieren.The results show that turbulent kinetic energy is mainly placed in the outer of the reattached boundary layer after interaction with shock wave.A low turbulent kinetic energy region is formed near the corner.The large scale streamwise motion of the separation shock was analyzed by using intermittent factor.It is found that the main shock moves around the averaged separation position and the length scale is equal to 72%of inlet boundary layer thickness.The presence of G?rtler-like vortices has been confirmed.Detail investigation into the spanwise distribution and spatial evolution characteristics of these vortices was performed.In addition,it is found that the turbulence information in the upstream boundary layer is not directly related to the low-frequency instability in the downstream.