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Study on Flow Characteristics of Compressible Laminar Flow Boundary Layers

机译：可压缩层流边界层流动特性研究

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The turbulence-transition model developed from the RANS methods has become the most used model in studying flow transition of hypersonic flight vehicles since it overcomes the flaw of stability analysis method that large-scale distributed parallel computing is not applicable. The current turbulence-transition model makes a numerical correction on compressibility in the prediction of transition within the hypersonic boundary layer, resulting in poor generality, so it needs to be further improved and perfected. Targeting this, the equations for the dimensionless boundary layer of compressible flow with pressure gradient are established, with the velocity profile and temperature profile similar outcome of the laminar plate obtained using the local similarity method. The influence of Mach number and pressure gradient on key parameters such as momentum loss thickness and shape factor are explored. The simulation results show that pressure gradient is the key factor forjudging transition of the boundary layer. A new fitting equation for determining hypersonic transition model is established.
机译：从Rans方法中开发的湍流转换模型已成为高超声速飞行车辆的流动过渡中最常用的模型，因为它克服了稳定性分析方法的缺陷，大规模分布式并行计算不适用。电流湍流转换模型对超声边界层的过渡预测中的可压缩性进行了数值校正，导致差的差，因此需要进一步改善和完善。靶向这一点，建立了具有压力梯度的可压缩流量的无量纲边界层的方程，具有使用局部相似性方法获得的层板的速度曲线和温度曲线类似的结果。探讨了马赫数和压力梯度对诸如动量损失厚度和形状因子的关键参数的影响。仿真结果表明，压力梯度是对边界层过渡的关键因素。建立了用于确定超声转换模型的新拟合方程。

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