航空航天技术近几十年来发展迅速，对飞行器的大推力、高速度、可重复使用等方面提出了新的要求。发汗冷却由于其理论极限冷却能力高、可降低冷却流体需求的技术优点，是满足上述飞行器发展需求的有效方法，开展高热流密度和超声速条件下的发汗冷却基础问题研究具有重要的意义。目前发汗冷却技术在多孔基底材料、实验研究、理论模型和数值模拟方法方面已有大量的研究，但仍存在一些不足：如超声速条件下的相变发汗冷却基础研究较为缺乏，相变发汗冷却流量控制难；自抽吸发汗冷却系统极限冷却能力较低，易断流等。同时高速飞行器中实际环境因素对发汗冷却的影响研究也较为缺乏。因此，本文针对上述问题开展了较为系统的研究。为了揭示超声速条件下的平板相变发汗冷却规律，在超声速中温主流和高焓超声速主流风洞中研究了注入率、主流温度、平均粒径、平板攻角等参数对平板相变发汗冷却的影响规律。实验中发现单一冷却流体注入压力可能对应多个冷却流体流量，平均粒径对相变发汗冷却效率的分布有显著影响。在总温3017 K、马赫数3.4、攻角8°的条件下，仅需注入率F =0.28%的冷却水即可实现多孔平板表面高达97%的均匀的发汗冷却效率。为了提升自抽吸发汗冷却系统的最大冷却能力和稳定性，发展了微纳结构改性增强自抽吸发汗冷却方法，采用热化学生长法和热化学腐蚀法在微米青铜颗粒烧结多孔材料上加工了纳米表面结构，显著提升了系统的最大冷却能力以及在热冲击和热流过载时的稳定性，研究了自抽吸发汗冷却系统的冷却流量对环境热流密度的自适应规律并提出了准则关联式。为了探索自抽吸发汗冷却在高超飞行器上的应用潜力，设计了飞行器外壳的自抽吸发汗冷却蒙皮模块和高温壁面自抽吸发汗冷却模块，研究了临近空间的低压工作环境对自抽吸发汗冷却的影响机理，提出了考虑压力影响的自抽吸发汗冷却理论分析方法，在热考核实验中获得了良好的热防护效果。为了对大推力火箭喷注面板进行有效的热防护，研究了不同参数的发汗/气膜复合冷却对喷注面板的保护效果，分析了局部热平衡模型在燃烧室高温燃气的极高热流环境中的偏差，提出了喷注面板的局部非热平衡计算方法。

With the rapid development of aerospace technology, the requirements for large thrust rockets, high speed vechicles, and repeatability of aircraft are getting higher and higher. Transpiration cooling technology is an effective method to meet the requirements due to its advantages of large theoretical cooling capacity and low coolant demand. It is of great significance to carry out research on the transpiration cooling under the conditions of high heat flux and supersonic flow. Transpiration cooling has been studied by a lot of researchors in porous materials, experimental research, theoretical models, and numerical simulation methods. There are still some shortcomings: such as the lack of transpiration cooling with phase-change under supersonic conditions with a problem of controlling the cooling flow rate; the cooling capacity of self-pumgping transpiration cooling system is still limited, and the coolant flow could break down easily. There is alos a lack of research on the influence of actual environmental factors on transpiration cooling under high-speed flight conditions. Therefore, this dissertation systematically studies the above shortcomings and difficulties:In order to explore the law of flat under supersonic conditions, the effects of injection rate, mainstream parameters, particle diameter and attack angle were studied in the supersonic medium temperature uniform mainstream wind tunnel and the high enthalpy supersonic mainstream wind tunnel. The experiment found that there is a one-to-many mapping relationship between the injection pressure and mass flow of the coolant, and the particle diameter has a significant influence on the distribution of phase-change transpiration cooling efficiency. Under the conditions of a total temperature of 3017 K, a Mach number of 3.4, and an angle of attack of 8°, transpiration cooling with on a cooling water injection rate of F = 0.28% can achieve a uniform cooling efficiency of up to 97%.In order to improve the cooling capacity and stability of the self-pumping transpiration cooling system, a micro-nano structure enhanced self-pumping transpiration cooling method has been developed. The thermochemical growth method and thermochemical corrosion method are used to fabricate nano-scale surface structures on bronze particles sintered porous material. The surface structure effectively improves the wettability of the porous material and significantly improves the maximum cooling capacity and the stability of self-pumping transpiration cooling during thermal shock and heat flow overload. The empirical correlation between the mass flux of the coolant flow and the heat flux is proposed based on the experimental data of self-pumping transpiration cooling.In order to explore the potential applications of the self-pumping transpiration cooling method in supersonic aircraft, the self-pumping transpiration cooling shell module of the aircraft and the self-pumping transpiration cooling module on the high temperature wall were designed. The effects of actual low-pressure working environment on self-pumping transpiration cooling were studied, and a theoretical analysis method of self-pumping transpiration cooling considering the influence of pressure is proposed. The self-pumgping transpiration modules were protected well during the experiments.In order to effectively protect the injection panel in the thrust chamber of large thrust rocket, the combined transpiration/film cooling method is proposed. The effect of different parameters on the cooling perfomence of the combined cooling method was studied. The deviation of the local thermal equilibrium model under the condition of extremely high heat flux in the combustion chamber is analyzed, and the numerical method based on local non-thermal equilibrium model was presented for transpiration cooling on the injection panel.