高速飞行器表面或喷气发动机内部通常会出现湍流边界层。当飞行器外部或发动机内部的激波与湍流边界层发生相互作用时，就会产生激波/湍流边界层干扰现象。激波/湍流边界层干扰会引起流动分离、强压力脉动和局部峰值热流等现象，并严重破坏飞行器的气动性能。本文针对高超声速条件下的激波/湍流边界层干扰现象进行了数值模拟研究，对其中的流场结构、湍流边界层演化特性、壁面摩擦阻力和壁面热流的生成机制等进行了研究。主要内容如下：

(1)采用大涡模拟方法研究了高超声速压缩拐角构型的流动，对比三个不同拐角角度（14°、24°和34°）对激波/湍流边界层干扰的影响规律。研究表明，增大拐角角度会引起分离激波强度及分离区尺度的增大。小角度情况下干扰区内的拟序结构表现为茎状或藤条状的流向涡，而大角度工况则表现为大量随机排列的发卡涡包。观察边界层内Görtler涡结构的分布发现，Görtler涡结构会随着拐角角度的增大而增强。通过对湍流边界层的演化特性进行详细分析，结果发现拐角角度的增大会引起湍流脉动放大的增强，并且拐角角度的变化对干扰区的湍动能输运机制也产生了显著的影响。

(2)采用直接数值模拟方法研究了高超声速凹曲壁的摩阻生成机制。考虑了不同弧角半径（10mm、20mm、40mm）的凹曲壁及压缩拐角工况。结果表明，凹曲壁可以减小激波/湍流边界层干扰流场的分离特征。研究发现，压缩拐角和凹曲壁工况都会显著的增强干扰区的雷诺应力。然而拐角工况的湍流放大主要由分离激波引起，而弧角工况则主要是由凹曲壁产生的逆压梯度所导致。通过对壁面摩擦阻力进行分解，发现与湍流运动相关的贡献项始终保持较大的值。采用二维经验模态分解技术定量分析不同尺度湍流运动对壁面摩阻生成的贡献。结果表明，上游无干扰区主要是由近壁小尺度运动主导。在下游再附区，拐角工况主要由外区大尺度结构的贡献占据主导；而对于凹曲壁工况，大尺度和小尺度结构的贡献相当。

(3)采用直接数值模拟方法进行了高超声速压缩拐角马赫数效应（Ma_∞=3、5、7）的研究。结果表明，来流马赫数的提升会引起分离区的减小，主激波会更加贴近壁面并导致下游边界层的运动受到抑制。此外，增大马赫数还会减小湍流放大现象。摩阻分解结果表明下游再附区的粘性耗散与马赫数呈正相关关系。针对来流马赫数为7的高马赫数壁面热流分析表明，雷诺应力作功项对壁面热流的生成提供主要的正贡献，而湍流热输运项则提供主要的负贡献。对壁面热流的尺度分析表明，上游无干扰区的雷诺应力作功项主要由小尺度结构主导，而湍流热输运项则由小尺度结构和大尺度结构共同提供贡献；在下游再附区，雷诺应力作功项和湍流热输运项的生成则都是由大尺度结构提供主要贡献。

Turbulent boundary layers often exist on the surfaces of high-speed aircraft or inside jet engines. Shock wave/turbulent boundary layer interactions occur when shock waves outside the aircraft or inside the engine interact with the turbulent boundary layer. Shock wave/turbulent boundary layer interactions can lead to flow separation, strong pressure fluctuations, and local peak heat flow, significantly affecting the aircraft’s aerodynamic performance. In this paper, numerical simulations of shock wave/turbulent boundary layer interactions under hypersonic conditions are conducted to investigate flow structures, turbulent boundary layer evolution characteristics, mechanisms for the generation of skin friction drag, and wall heat flux. The main contents are as follows:

(1) The flow law of the hypersonic compression corner configuration is investigated using large eddy simulation to compare the effects of three different corner angles (14°, 24°, and 34°) on the shock wave/turbulent boundary layer interaction. It is shown that increasing the angle of the corner causes an increase in the intensity of the separated shock and the scale of the separation zone. Coherent structure within the interaction zone in the small angle case shows stem-like or cane-like flow vortices, while the large angle case shows a large number of randomly arranged hairpin vortex packets. Observation of the distribution of Görtler structures within the boundary layer reveals that the Görtler structures are enhanced with increasing corner angle. By analyzing the evolutionary characteristics of the turbulent boundary layer in detail, it is found that the increase of the angle of compression corner causes the enhancement of the turbulent fluctuation amplification, and the change of the compression corner also has a significant effect on the turbulent kinetic energy transport mechanism in the interaction region.

(2) The drag generation mechanism of hypersonic concave surface is investigated by direct numerical simulation. Different radii (10mm, 20mm, 40mm) of the concave surface and compression corner conditions are considered. The results show that the concave surface can reduce the separation characteristics of the shock wave/turbulent boundary layer interaction flow field. It is found that both compression corner and concave surface conditions significantly enhance the reynolds stress in the interaction zone. However, the turbulence amplification in the corner case is mainly caused by the separation shock, while the concave surface case is mainly caused by the reverse pressure gradient generated by the concave surface. By decomposing the skin friction drag, it is found that the contribution term associated with turbulent motion consistently maintains a large value. The Bidimensional Empirical Mode Decomposition technique is used to quantitatively analyze the contribution of different scales of turbulent motion to the generation of skin friction drag. The results show that the upstream undisturbed region is mainly dominated by near-wall small-scale motions. In the downstream reattachment zone, the corner case is mainly dominated by the contribution of large-scale structures in the outer zone; while for the concave surface case, the contributions of large-scale and small-scale structures are comparable.

(3) Direct numerical simulation are carried out to investigate the Mach number effects (Ma_∞=3, 5, and 7) in hypersonic compression corner. The results show that an increase in the Mach number of the incoming flow induces a decrease in the separation zone, and the main shock wave comes closer to the wall and leads to the suppression of the motion of the downstream boundary layer. In addition, increasing the Mach number reduces turbulence amplification. The results of the drag decomposition show that the viscous dissipation in the downstream reattachment zone is positively correlated with the Mach number.

The analysis of the high Mach number wall heat flux for an incoming Mach number of 7 shows that the reynolds stress work term provides the main positive contribution to the generation of wall heat flux, while the turbulent heat transport term provides the main negative contribution. The scaling analysis of the wall heat flux shows that the reynolds stress work term in the upstream undisturbed region is dominated by small-scale structures, while the turbulent heat transport term is contributed by both small and large scale structures; in the downstream reattachment region, the generation of both the reynolds stress work term and the turbulent heat transport term are dominated by large scale structures.