高超声速飞行器鼻锥迎风凹腔结构防热效能研究  

Investigation on Thermal Protection Efficiency of Hypersonic Vehicle Nose with Forward-Facing Cavity

作　　者：陆海波[1] 刘伟强[1] 

机构地区：[1]国防科技大学航天与材料工程学院,长沙410073

出　　处：《宇航学报》2012年第8期1013-1018,共6页Journal of Astronautics

基　　金：国家自然科学基金(90916018);高等学校博士学科点专项科研基金(200899980006)

摘　　要：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。A numerical study on the effect of forward-facing cavity upon aerodynamic heating on the hypersonic vehicle nose is conducted. The numerical simulation resuh is validated by the experiment. Flow field parameters and heat flux distributions along the outer body surface and the cavity wall are obtained and the cooling effect of the forward-facing cavity with different dimensions is analyzed. The results show that the forward-facing cavity configuration works well in cooling the nose of hypersonic vehicles especially at the stagnation point area. The deeper the cavity, the smaller the heat flux. The maximal heat flux does not locates at the peak of the sharp lip, but at the upper wall of the cavity, and the cavity deep （L） has little effect on the location of the maximal heat flux. There is a very low surface heat flux along the base wall of the cavity except the cavity has a very shallow L （L/D 〈 O. 5 ）.

关 键 词：迎风凹腔 热防护 鼻锥 高超声速 

分 类 号：V211.3[航空宇航科学与技术—航空宇航推进理论与工程]