高超声速风洞是高超声速飞行器与发动机研究的重要地面实验设备。对于高超声速风洞系统而言，其核心部件之一是能将气流由亚声速状态转变成超声速状态的超声速喷管。由于长时间受到高速、高温、有氧气流的冲刷，超声速喷管喉道处壁面热流密度很高，这为高超声速风洞的正常运行带来了极大的挑战，诸多针对喷管的热防护方法也应运而生。
研究表明，气膜冷却具有结构简单、冷却效率高、可重复使用的优点，在高超声速风洞喷管中有很大的应用前景。但气膜在超声速环境与大曲率壁面上的流动与传热特性尚不明确，气膜冷却效率以及引入气膜对喷管出口气流品质的影响
也不清楚。因此，急需系统研究超声速喷管气膜冷却特性及其流动机理。本文将
基于雷诺时均方法与 SST 𝑘−ω 湍流模型，对气膜作用下马赫数 6高超声速喷管
流动、壁面热载荷以及出口气流品质进行数值研究。主要内容与结论有：第一部分对超声速喷管中狭缝气膜冷却的二维流动与传热特性进行了数值研究。使用狭缝平行喷射空气的方式对喷管壁面进行冷却。数值结果表明，气膜可以有效降低喷管壁面，尤其是喉道处的热流密度。当气膜流量仅占主流 3.53%时，喉道热流密度降低了 30.1%。引入气膜不会对喷管出口马赫数分布造成显著影响，仅对喷管出口总温分布造成较小的影响，使得温度边界层变厚。同时，相同气膜流量条件下，狭缝高度、台阶厚度、射流距离的改变对喷管喉道的冷却效率影响很小。
第二部分对超声速喷管中壁面多孔气膜冷却的三维流动与传热特性进行了数值研究。采用多孔喷射空气的方式对喷管壁面进行冷却，研究了射流角度、气膜流、气膜孔径等参数对冷却效率及喷管出口气流品质的影响。结果表明，气膜可以有效降低喷管整体热流密度。当射流角为 30°，气膜流量仅占主流 3.07%时，喉道平均热流密度下降 31.1%。由于气膜孔存在间距，喷管喉道热流密度分布呈周期性分布，在正对气膜孔展向位置，喉道热流密度出现最低值；在相邻气膜孔的中心位置，喉道热流密度出现峰值。同时研究发现，射流角过大时气膜易与壁面发生分离，使冷却效率急剧下降，同时对喷管出口气流品质影响较大。而射流角度较小时，气膜贴壁发展程度更高，对喷管出口气流品质影响较小。

Supersonic wind tunnel is an important ground experimental equipment for the research of hypersonic aircraft and engines. For a hypersonic wind tunnel system, one of its core components is the supersonic nozzle which can transform the airflow from subsonic state to a supersonic state. Due to the long-term scouring of high-speed, high-temperature and oxygenated flow, the wall heat flux at the throat of supersonic nozzle is very high, which brings great challenges to the normal operation of the hypersonic wind tunnel, and many thermal protections method for the nozzle also came into being.
Studies have shown that film cooling has the advantages of simple structure, high cooling efficiency, and reusability. Thus, film cooling has great application prospects in hypersonic wind tunnel nozzles. However, the flow and heat transfer characteristics of the film in supersonic and large curvature wall environment are not clear. The cooling efficiency of the film and the influence of the introduction of the film on the quality of the airflow at nozzle outlet are also unclear. Therefore, it is urgent to systematically study the film cooling characteristics and flow mechanism of supersonic nozzles. In this paper, based on the Reynolds-Average method and the SST k-ω turbulence model, the flow characteristics, wall thermal load characteristics and outlet airflow quality of the Mach 6 hypersonic nozzle with the gas film are numerical studied. The main contents and conclusions are:
In the first part, the two-dimensional flow and heat transfer characteristics of film cooling in supersonic nozzle are numerically studied. The nozzle wall was cooled by the way of parallel air injection through a slot. The numerical results show that the air film can effectively reduce the heat flux at the nozzle wall, especially at the throat. When the air film flow rate only accounts for 3.53% of the mainstream, the heat flux at the throat decreases by 30.1%. The introduction of gas film will not significantly affect the distribution of Mach number at the nozzle outlet, but only has a small effect on the total temperature distribution at the nozzle outlet, which makes the temperature boundary layer thicker. At the same time, under the same air film flow condition, the change of slot height, step thickness, and jet distance has little effect on the cooling efficiency of the nozzle throat.
In the second part, the three-dimensional flow and heat transfer characteristics of film cooling in supersonic nozzle are numerically studied. The nozzle wall is cooled by a porous air injection, and the effects of parameters such as jet angle, film flow rate, and film aperture on the cooling efficiency and the airflow quality of the nozzle outlet are studied. The results show that the gas film can effectively reduce the overall heat flux of the nozzle. When the jet angle is 30° and the film flow only accounts for 3.07% of the main flow, the average heat flux of the throat drops by 31.1%. Due to the film hole spacing, the nozzle throat heat flow distribution is periodically distributed. The heat flux is the lowest in the opposite position of the film hole. And has the peak value in the center of the adjacent film hole. Moreover, studies have found that when the jet angle is too large, the air film is easier to separate from the wall, which makes the cooling efficiency drop sharply, having a greater impact on the quality of the airflow at the nozzle outlet. When the jet angle is small, the air film adheres to the wall and has less impact on the quality of the airflow at the nozzle outlet.