Abstract The normal shock wave-boundary layer interaction (SBLI) phenomenon is known to constitute a main factor limiting the aerodynamic performance in many aeronautical applications (transonic wings, helicopter rotor blades, compressor and turbine cascades). The interaction process highly disturbs the boundary layer, often causing flow separation and onset of large scale unsteadiness (e.g. airfoil buffet or supersonic inlet buzz). In certain conditions it may also initiate a dramatic increase of acoustic emission levels (e.g. high-speed impulsive noise). To limit the negative impact of the phenomenon various flow control strategies are implemented, here in a form of a passive control system realised by placing a shallow cavity covered by a perforated plate just beneath the shock. Details of the flow structure obtained by this method are studied numerically. Three distinctive experimental set-ups are considered with the interaction taking place: on a flat wall (transonic nozzle, ONERA), on a convex wall (curved duct, University of Karlsruhe), and on an airfoil (NACA 0012, NASA Langley). Depending on the relative cavity length the ventilation process leads to a transformation of the normal shock topology into: a large λ-foot structure (classical, short cavity), a system of oblique waves (extended cavity), or a gradual compression (full-chord perforation). The reference and flow control cases are simulated with the SPARC code (RANS) with Spalart–Allmaras turbulence and Bohning–Doerffer transpiration models. The results are compared with the measurements, emphasizing the streamwise evolution of the boundary layer profiles and integral parameters during the interaction. The prediction capabilities of the solver in terms of the shock wave-boundary layer interaction control by wall ventilation are assessed and presented in details for the investigated range of flow configurations and conditions.

中文翻译：

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被动墙通风控制冲击波-边界层相互作用的数值模拟

摘要 正激波-边界层相互作用（SBLI）现象是许多航空应用（跨音速机翼、直升机旋翼叶片、压气机和涡轮叶栅）中限制气动性能的主要因素。相互作用过程高度扰动边界层，经常导致流动分离和大规模不稳定（例如翼型抖振或超音速进气嗡嗡声）的开始。在某些情况下，它也可能引发声发射水平的急剧增加（例如高速脉冲噪声）。为了限制这种现象的负面影响，实施了各种流量控制策略，这里采用被动控制系统的形式，通过在冲击下方放置一个由多孔板覆盖的浅腔来实现。通过这种方法获得的流动结构的细节进行了数值研究。三个独特的实验装置被认为是相互作用的：在平坦的壁上（跨音速喷嘴，ONERA），在凸壁上（弯曲的管道，卡尔斯鲁厄大学）和翼型（NACA 0012，NASA Langley）。根据相对空腔长度，通风过程导致正常激波拓扑结构转变为：大 λ 英尺结构（经典、短空腔）、斜波系统（扩展空腔）或逐渐压缩（全-弦穿孔）。参考和流量控制案例使用 SPARC 代码 (RANS) 与 Spalart-Allmaras 湍流和 Bohning-Doerffer 蒸腾模型进行模拟。将结果与测量值进行比较，强调相互作用过程中边界层剖面和积分参数的流向演化。评估求解器在冲击波-边界层相互作用控制方面的预测能力，并详细介绍了所研究的流动配置和条件范围。