超声速压气机叶栅前缘通道激波损失的鼓包控制研究

为了有效减小超声速压气机叶栅变进气马赫数条件下的前缘通道激波损失及由激波诱导的边界层分离,提出了一种带有平直过渡区的新型鼓包结构,并采用数值方法详细分析了新型鼓包结构对激波与激波/边界层相互作用机理以及鼓包几何尺寸与位置对控制效果的影响机制。研究结果表明:新型鼓包在迎风侧凹面产生的压缩波系有效削弱了前缘通道激波的强度,鼓包过渡区产生的膨胀波系使边界层流体加速,明显抑制了局部流动分离,并使分离提前再附。当某一超声速压气机叶栅的前缘通道激波入射在鼓包的过渡区范围内,鼓包高度为0.35倍的边界层厚度且鼓包迎风侧与背风侧长度分别为过渡区长度4倍与5倍时,可以实现较好的控制效果。此外,与无鼓包方案相比,新型鼓包结构可使超声压气机叶栅在设计工况下的总压损失减少4.6%,同时超声速压气机叶栅进气马赫数在1.65~1.8范围内仍能取得较好的气动减损效果。 

Investigation on the first passage shock loss with bump control inside a supersonic compressor cascade

In order to effectively reduce both losses induced by the first passage shock and resulted from boundary layer separation inside a supersonic compressor cascade at variable inlet Mach numbers, a type of method employing a new kind of 2-D bump structure with straight transition region on the suction side of the supersonic compressor cascade was presented to control the shock-boundary layer interaction. Numerical simulations and then detailed analyses to demonstrate the control mechanism of the 2-D bump on shocks and shock-boundary layer nteractions were conducted as well to investigate the influences of geometric dimensions and positions of the 2-D bump on control results. The investigations show that the compression wave system induced from concave windward surface of the 2-D bump effectively weakens the intensity of the first passage shock and that the expansion wave system produced by the transition region of the bump accelerates the flow in boundary layer, both of which suppress the local flow separation significantly and make the separation reattach in advance. While the first passage shock inside a supersonic cascade impinged on the straight section of the bump, optimal control effect was achieved at dimensionless height of the bump which retained 0.35 times of boundary layer thickness, and also at dimensionless length of windward and leeward side which remained 4 and 5 times of transition region respectively. Moreover, compared with the cascade without 2-D bump, the level for total pressure loss reduced by the bump reached 4.6% with the free stream Mach number 1.75, while the aerodynamic loss of the supersonic compressor cascade with bump was still better than that without bump when the inlet Mach number varied from 1.65 to 1.80.