Simulation of Hypersonic Shock-Boundary Layer Interaction

• Main Authors: Wei-Jen Lo, 羅偉仁
• Other Authors: Tzong-Hann Shieh
• Format: Others
• Language: zh-TW
• Published: 2011
• Online Access: http://ndltd.ncl.edu.tw/handle/65104448648249790865

碩士 === 逢甲大學 === 航太與系統工程所 === 99 === This paper describes using finite volume method to solve Navier-Stokes Equations about a hypersonic intake has a double-ramp compress section. Considering varied computing methods to process 2-D computing simulations of flow field. Computing simulations of the intake wall temperature rise and second intake face angle rise, we have further research about the flow field change result from the shock wave/boundary layer interaction. During the research, we test the computing flux type first, the final details show that the computing outcomes of Roe-FDS and AUSM flux type are similar, since the Roe-FDS has more stable computing process than AUSM, we choose Roe-FDS to be our computing flux type. About mesh test, we increase the mesh number at the leading edge of the intake, corner between two faces and wall. We also try to use mesh adaptation method to raise mesh number by the density contour surface gradient. About turbulence computing model test, simulation outcomes show only the SST k-ω computing model with the low Reynolds number correction could simulate the physics phenomenon of the shock wave/boundary layer interaction for the hypersonic air flow through the corner of two ramps in intake. The outcomes of the computing for the increasing intake wall temperature show that the range of the shock wave/boundary layer interaction region becomes bigger related with the increasing wall temperature, it has the biggest change when temperature rises from 300K to 600K. We also considering the effect of the variable specific heat ratio to the flow field characteristics in the increasing wall temperature computing, the effect fewer when the wall temperature is lower, contrariwise, the effect obvious when the wall temperature higher. We also compare the outcomes of increasing intake second ramp angle and increasing wall temperature, the range of the shock wave/boundary layer interaction region is smaller when the second face angle increases test case. The Stanton number pattern of the intake is mainly affected by the shock wave/boundary layer interaction, when the interaction between shock wave and boundary transits from laminar boundary to the turbulence boundary; it clearly shows the boundary layer separation forward with the increasing angle.