摘要
目的针对未来高超声速飞行器(飞行速度20 Ma),提出一种主动式气膜冷却防热技术,并计算验证其有效性。方法通过求解三维N-S方程组,采用PARK-1的5组分(N2,O2,N,O,NO)17方程有限速率化学反应模型,考虑了真实气体效应;针对典型的钝头体外形,在头部驻点处构造单个气膜冷却微孔,向外喷射冷却气体,计算了飞行马赫数为20、高度在30 km以下的气膜冷却效率。结果与无气膜冷却相比,有气膜冷却时,气膜孔附近等温壁面(300 K)热流密度的最高降幅约90%,冷却气体有效覆盖面积可达到约孔出口面积的10倍。结论气膜冷却在未来高超声速飞行器防热中具有广阔的应用前景。
Objective To put forward a new approach of active thermal protection system termed as film cooling for future hypersonic vehicles （Mach 20 flow） and validate its performance by CFD. Methods By solving Reynolds-averaged Navier - Stokes equations, the finite-rate chemical reaction model with Park-I 5 components （N2, 02, N, O, NO） 17 equations was used to investigate the real gas effect. For the typical blunt body, a numerical study of the effectiveness of film cooling was presented for Mach 20 flow at 30 kin, with the special designed micro inject hole located at the stagnation point. Results The reduction of heat flux on the isothermal wall （300 K） with film cooling at the vicinity of the hole could be as much as 90% compared to that without film cooling, and the covering area of the coolant flow could be l0 times larger than the out area of the hole. Conclusion The study showed promising prospect of film cooling for heat proof of future hypersonic vehicles.
出处
《装备环境工程》
CAS
2015年第3期1-7,共7页
Equipment Environmental Engineering
