飞行器鼻锥凹腔-发散组合冷却数值模拟

尖锐鼻锥冷却方案是可复用式航天飞行器研究领域一个十分重要的课题。传统发散冷却虽然可以有效降低鼻锥结构温度，但是由于驻点外极高的热流、压力，会出现驻点冷却效果差的问题。迎风凹腔结构是一种针对鼻锥驻点区域的减阻防热方案，尖锐唇口的分流作用可以使附近压力、热流降低。因此，提出一种新型冷却结构——凹腔-发散组合冷却，利用迎风凹腔结构对驻点的强化冷却解决发散冷却中驻点难以冷却的问题。以楔形鼻锥为物理模型，对发散冷却、迎风凹腔结构和凹腔-发散冷却3种冷却结构进行数值模拟，并和无冷却的纯鼻锥结构进行对比。结果表明，与传统发散冷却相比，使用凹腔-发散组合冷却可以使结构温度峰值下降16.8%；与没有冷却的纯鼻锥模型相比，鼻锥头部圆弧段表面平均温度降幅可达64%，证实了这种新型冷却结构的可行性和高效性。

Transpiration cooling of nose-cone with forward-facing cavity: Numerical simulation

The thermal protection system of sharp nose-cones is an important subject in the investigation of reusable spacecraft. Transpiration cooling, with an ability to effectively reduce the temperature of the nose-cone, has been widely recognized as a promising thermal protection approach. However, the cooling effect at the stagnation point remains poor due to the extremely high heat flux and pressure there. The nose-cone with forward-facing cavity is a drag and heat reduction structure specifically designed for the stagnation region. It can reduce the heat flux and pressure there with the shutting effect of the sharp lip. Therefore, a new combined cooling structure is proposed in this paper: transpiration cooling with forward-facing cavity. It utilizes the cooling enhancement effect of the forward-facing cavity to solve the problem of low cooling efficiency at the stagnation region. Using a wedge-shaped nose cone as the physical model, the cooling performance of three structures, i.e., transpiration cooling, forward-facing cavity and transpiration cooling with the cavity, is studied and compared. The simulation results indicate that, using the new combined cooling structure, the maximum temperature of the nose-cone can be reduced by 16.8% compared with that of the traditional transpiration cooling, and the average circular surface temperature can drop by 64% compared with that on the pure nose-cone, confirming the feasibility and high efficiency of the new cooling structure.