摘要
为模拟高超声速湍流问题,对剪切应力输运(SST)湍流模型系数进行了修正。数值格式采用改进的总变差递减(TVD)格式,并对湍流模型的负值强制项进行了隐式处理。在此基础上计算了绕平板以及具有分离、再附、激波/边界层干扰等复杂流动结构的压缩拐角的高超声速流动。计算结果与试验数据及半经验公式的对比表明:SST湍流模型引入的雷诺剪切应力与湍动能之比为常数(Bradshaw数)在高超声速绕流中并不成立。Bradshaw数修正后的SST湍流模型与原模型相比,所计算的壁面压力、摩擦阻力和壁面热流分布更接近试验结果。
An improved shear stress transport （SST） turbulence model is proposed for hypersonic flows in this study which allows for corrected coefficients. A numerical scheme is established by making use of the improved total variation diminishing （TVD） scheme and by applying an implicit scheme to the negative source terms of the turbulence model. Hypersonic flat- plate boundary-layer flows and hypersonic compression ramp flows marked with separation, reattachment and shock/bound- ary layer interactions are then computed. A comparison between the computational results, the experimental results and the semi-empirical formulations shows that the ratio of Reynolds shear stress and turbulent kinetic energy is not a constant for hy- personic flows. For flows with adverse pressure gradients, shock-wave/turbulent-boundary interaction and separation, cal- culation results of wall pressures, friction coefficients and wall heat transfer rate distributions using the improved model and established scheme agree better with the experimental results than those using the original SST model.
出处
《航空学报》
EI
CAS
CSCD
北大核心
2012年第12期2192-2201,共10页
Acta Aeronautica et Astronautica Sinica
基金
国家自然科学基金(11102079)
航空科学基金(20111456005)~~
关键词
湍流模型
数值模拟
高超声速
激波
边界层
热流
turbulence model
numerical simulation
hypersonic
shock wave
boundary layer
heat flux
