前缘钝度对马赫数6平面压缩进气道流场的影响分析

Leading Edge Bluntness Influence Anaslysis on a Planar Compression Inlet Flowfield at Mach Number 6

高超声速飞行要求前缘钝化,钝化前缘将引起流场变化并影响进气道性能,进气道几何参数需要进行修正设计。为了解钝化前缘影响的程度及条件,为修正设计提供依据,采用S-A模型,研究了马赫数6条件下,固定外罩前缘钝度(Rc=2mm),前体前缘钝化(RN=1~6mm)对一种3+1波系平面压缩进气道流场特征和综合性能的影响。结果表明,前体前缘钝化引起外压缩激波偏离唇口、肩部分离区向上游传播,造成分离激波与其他激波形成复杂的相交结构,进而形成严重畸变的唇口弓形激波,在较大钝度(RN≥3mm)时激波干扰生成的滑移层进入内流道核心区,唇口弓形激波的畸变和大范围的滑移层吞入,是造成总压恢复系数严重下降的主要原因;外压缩激波的外移以及分离激波的介入是造成流量捕获急剧下降的主要原因。由此提示,应当通过合理的钝化设计或流场修正设计,使前体前缘钝化的进气道避免外压缩激波偏离唇口太远,并通过流动控制方法严格控制肩部分离区范围,勿使分离激波与其他激波干扰产生的滑移层进入内流道核心区范围,更要避免分离激波与唇口弓形激波在内流道核心区域相交。

Leading-edges for hypersonic systems might be blunted,and forbody-inlet flowfield might be distortion and performance is deteriorated. The inlet geometry needs to be redesigned. To know critical factors and conditions before inlet redesign,S-A model is used to study forebody leading-edge bluntness effect on flowfield and properties in a 3＋1 shock system planar inlet at Ma=6. Forebody leading-edge radius is from RN=1 to 6 mm with a cowl-lip radius of 2 mm. The results show that external compression shocks move away from cowl-lip and shoulder separated zone goes upstream with forebody leading-edge more blunted. The separation shock interacts with other shocks and cause complex flowfield. The cowl-lip shock is distorted seriously. When RN≥3 mm,shear layers from shock-shock interaction point are swallowed into the center zone of internal flow-path. The distorted cowl-lip shock and large scope of shear layers are critical factors to sharply decrease pressure recovery. The moving of external compression shocks and separation shock interaction lead to remarkable mass-flow decrease.These results suggest that reasonable bluntness and flowfield design is necessary to avoid external compression shocks moving far away cowl-lip. The separated zone size should be controlled strictly to avoid swallow of shear layers from the interaction point between separation shock and cowl-lip shock.