高超声速飞行器及发动机是空天科技发展的核心装置。超声速燃烧室是吸气 式高超声速动力系统的关键部件，其面临着非常恶劣的热环境。热防护一直是高 超声速飞行器与发动机研究的重要领域。目前广泛使用的碳氢燃料再生冷却技术 基于管道对流传热机制，能够满足燃烧室绝大部分区域的结构冷却需求。但随着 飞行马赫数与发动机结构复杂度的不断提高，发动机燃烧室承受的热载荷也在不 断增大，其冷却性能已明显不足。因此需要发展更高效的冷却技术手段，以保证 相关部件及结构能够长时间安全运行。气膜冷却作为一种高效的主动冷却方式， 有望应用于超声速燃烧室的热防护。以往关于气膜冷却的研究主要针对亚声速来 流条件，讨论气膜冷却效率的影响因素，而关于超声速来流中气膜冷却特性的研 究较少，超声速来流的强可压缩效应以及复杂波系的干扰，使得超声速条件下气 膜冷却机理显著不同于亚声速条件，相关流动与传热机理尚不清楚。且目前关于 气膜冷却在超声速燃烧室的应用研究极少，因此，研究超声速燃烧室内气膜的流 动、传热以及减阻特性，揭示其流动机理是非常必要的。

本文首先通过数值仿真研究了超声速来流条件下气膜冷却流动、传热与减阻 特性。数值仿真采用雷诺平均方法（RANS）。数值结果表明，冷却剂通过离散孔 喷注在下游形成覆盖气膜，有效降低了高温主流向壁面传热；喷注射流会形成典 型的反向旋转对涡结构，该涡促进了射流的抬升及其与主流的掺混过程，深刻影 响气膜冷却性能；同时多排孔气膜相互作用，进一步改变了气膜的扩散特性。研 究发现，超声速来流条件下，气膜能够降低近壁面流动的速度梯度，显著降低壁 面摩阻，且摩阻分布与壁面热流密度分布具有相似性。

数值研究了几何参数（包括气膜孔直径、气膜孔排布方式以及气膜孔展向间 距）、冷却流参数（包括气膜冷却剂流量、气膜喷注角度）以及不同马赫数的超 声速来流等对超声速环境中气膜冷却效果、壁面减阻效果的影响规律。几何参数 的研究发现，气膜孔采用交错排布方式具有更高的冷却效率和减阻性能。相同流 量下，增大气膜孔径、减小气膜孔展向间距会导致气膜孔排布区域冷却效率、减 阻性能增大，但气膜孔下游区域均下降较快。冷却流参数的研究发现，气膜流量 越大，气膜冷却效率越大，且产生强度不同的激波作用于气膜后，出现了冷却效 率的迅速下降以及略微上升两种变化趋势。气膜喷注角度越大，在展向分布上， 冷却效率分布越均匀；在流向分布上，临近气膜孔下游气膜冷却效率越大，但下 游较远处冷却效率迅速下降。针对气膜的减阻性能，气膜流量、喷注角度增大均 会导致其在临近气膜孔下游一定距离内呈现增大趋势。不同马赫数的超声速来流 研究发现，主流马赫数增大，气膜冷却效率沿流向不断减小，沿展向分布越不均 匀，且马赫数越大，产生的激波越强，对气膜冷却效率的不利影响也越大。在临 近气膜孔下游一定距离内，壁面摩阻随着来流马赫数的增大而减小。同时激波强 度也会对壁面减阻效果产生影响，激波强度越大，对壁面减阻效果的不利影响越大。

本文通过实验研究获得了真实超声速燃烧室环境下气膜冷却效率的基本变化 规律。实验针对不同冷却测试件构型，研究了气膜流量对其冷却效率的影响，包 含燃烧条件下以及纯气动加热条件下的气膜冷却。通过直接测量有、无气膜作用 时燃烧室内壁面热流密度变化，以评估气膜冷却性能。实验结果表明，超声速条 件下气膜冷却具有较高的冷却性能，当气膜流量占主流流量 9.7%时，燃烧室壁 面气膜孔下游的测点热流密度最高下降了 53%；且气膜流量越大，其冷却效率越 高。同时实验中气膜孔下游 90mm 位置的冷却效率高于气膜孔下游 20mm 位置， 结合数值仿真可知冷却剂喷注后形成稳定气膜、获得较高冷却效率需要一定的流 向距离。

理论研究方面，基于 Goldstein 的二维气膜冷却效率理论模型，并考虑离散 孔的三维效应、多排孔叠加效应等，首次建立了超声速燃烧室内多排均匀离散孔 气膜冷却效率的理论预测模型，预测结果与实验数据吻合良好，最大差异不超过 5%。

Hypersonic vehicle and engine are the core devices for the development of space science and technology. Supersonic combustor is the key component of aspirated hypersonic power system, which is faced with very bad thermal environment. Thermal protection is always an important research field of hypersonic vehicle and engine. The widely used hydrocarbon fuel regenerative cooling technology is based on the convective heat transfer mechanism in the pipeline and can meet the structural cooling requirements in most areas of the combustor. However, with the increasing of flight Mach number and the complexity of engine structure, the thermal load of engine combustor is also increasing, and its cooling performance is obviously insufficient. Therefore, more efficient cooling techniques need to be developed to ensure that the relevant components and structures can operate safely for a long time. As an efficient active cooling method, film cooling is expected to be applied to the thermal protection of supersonic combustor. Previous research on the film cooling is aimed at subsonic flow conditions, discuss the influence factors of the film cooling efficiency, and on supersonic flow gas film cooling characteristics of study is less, the strong compressible effect of supersonic flow and complex wave interference, makes the gas film cooling mechanism under the condition of supersonic significantly different from subsonic condition, related flow and heat transfer mechanism is unclear. At present, there are few studies on the application of gas film cooling in supersonic combustor. Therefore, it is very necessary to study the flow, heat transfer and friction reduction characteristics of gas film in supersonic combustor and reveal its flow mechanism.

Firstly, the flow, heat transfer and drag reduction characteristics of film cooling under supersonic flow condition are studied by numerical simulation. Reynolds average method (RANS) was used for numerical simulation. The numerical results show that the coolant is sprayed through discrete holes to form a covering film downstream, which effectively reduces the heat transfer from the high temperature mainstream to the wall surface. The injection flow forms a typical antirotation vortex structure, which dominates the invasion depth and mixing process of the film in the supersonic flow, and profoundly affects the cooling performance of the film. At the same time, the multi-row film interaction further changes the diffusion characteristics of the film. It is found that the gas film can reduce the velocity gradient of the flow near the wall and significantly reduce the wall friction under the condition of supersonic flow, and the friction distribution is similar to the wall heat flux distribution.

The effects of geometrical parameters (including film hole diameter, film hole arrangement and film hole spacing), cooling flow parameters (including film coolant flow rate, film injection Angle) and supersonic inlet Mach number on film cooling effect and wall friction reduction effect in supersonic environment were numerically studied. The study of geometric parameters shows that staggered arrangement of film holes has higher cooling efficiency and friction reduction performance. At the same flow rate, increasing the aperture of film and decreasing the spanwise spacing of film holes will lead to the increase of cooling efficiency and friction reduction performance in the distribution area of film holes, but the downstream area of film holes will decrease rapidly. The study of cooling flow parameters shows that the higher the film flow rate is, the higher the cooling efficiency is. After the shock wave with different intensity acts on the film, the cooling efficiency decreases rapidly and rises slightly. The larger the injection Angle, the more uniform the distribution of cooling efficiency. In terms of flow distribution, the cooling efficiency of the downstream film near the film hole is larger, but the mixing of coolant and the main shear is stronger, leading to a rapid decrease of the cooling efficiency at a distance downstream. For the friction reduction performance of the film, the increase of the film flow rate and the increase of the injection Angle will lead to an increase trend in a certain distance downstream of the film hole. The study of supersonic inlet Mach number shows that the film cooling efficiency decreases along the flow direction with the increase of the main flow Mach number, and the uneven distribution along the extension direction is more. The larger the Mach number is, the stronger the shock wave is, and the greater the adverse impact on the film cooling efficiency is. At a certain distance downstream of the film hole, the wall friction decreases with the increase of incoming Mach number. At the same time, the shock wave intensity also has an impact on the wall friction reduction effect, the greater the shock wave intensity, the greater the adverse impact on the wall friction reduction effect.

In this paper, the basic variation law of film cooling efficiency in real supersonic combustor is obtained through experimental study. Different configurations of cooling test pieces were used in the experiment to study the effect of film flow rate on its cooling efficiency, including film cooling under combustion condition and pure pneumatic heating condition. A high temperature heat flux sensor based on the principle of Gardon heat flux meter was used to directly measure the change of heat flux on the wall of combustion chamber with and without air film, so as to evaluate the change rule of film cooling performance and film cooling efficiency. The experimental results show that the film cooling has high cooling performance under supersonic condition. When the film flow rate accounts for 9.7% of the main flow rate, the heat flux of the measuring point downstream of the film hole on the combustion chamber wall decreases by 53%. And the larger the film flow rate, the higher the cooling efficiency. At the same time, the cooling efficiency of 90mm downstream of the film hole is higher than that of 20mm downstream of the film hole. Combined with numerical simulation, it can be seen that a certain flow distance is required to form a stable film and obtain a higher cooling efficiency after coolant injection.

Theory research, based on the two-dimensional Goldstein film cooling efficiency theory model, and considering the three-dimensional effect of discrete holes and hole superposition effect, etc., for the first time established the supersonic combustion chamber row more uniform configuration of discrete holes film cooling efficiency theory prediction model, the predicted results are in good agreement with the experimental data, the biggest difference is less than 5%.