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对高超声速飞行器鼻锥使用迎风凹腔结构作为热防护系统时,凹腔结构的防热效能进行了数值研究。通过与相关实验对比,验证了本文数值方法的可靠性,获得了鼻锥的流场参数,外表面、凹腔内壁面的热流分布,分析了不同的凹腔尺寸参数选择对鼻锥冷却效果的影响。结果表明迎风凹腔结构能够有效的对高超声速飞行器的鼻锥尤其是驻点区域进行冷却,凹腔越深,其冷却效果越好。鼻锥气动加热的最大热流并不在尖锐唇缘的顶点,而是位于凹腔内的侧壁面上,凹腔的深度(L)变化对最大热流的出现位置影响很小。除非凹腔很浅(L/D<0.5),凹腔底面的热流值都非常小,基本可以忽略。
The numerical results of the thermal performance of the cavity structure are studied numerically when the upwind cavity structure is used as the thermal protection system for the hypersonic nose cone. The reliability of the numerical method is verified by comparison with the related experiments. The flow field parameters of the nose cone, the outer surface and the heat flow distribution in the inner wall of the cavity are obtained. The effects of different cavity size parameters on the cooling effect of the nose cone influences. The results show that the windward cavity structure can effectively cool the nose cone, especially the stagnant area, of the hypersonic vehicle. The deeper the cavity, the better the cooling effect. The maximum heat flux of aerodynamic heating of nose cone is not at the apex of the sharp lip, but at the side wall in the cavity. The change of depth (L) of the cavity has little effect on the location of maximum heat flux. Unless the cavity is very shallow (L / D <0.5), the heat flow at the bottom of the cavity is very small and can be neglected.