【作者】 熊宴斌；

【导师】 姜培学；

【作者基本信息】 清华大学 ， 动力工程及工程热物理， 2013， 博士

【摘要】 随着航空航天飞行器技术的发展，液体火箭及吸气式发动机燃烧室壁面以及高超声速飞行器外壳的工作热环境越来越恶劣，发汗冷却技术被认为是最有可能解决未来航空航天飞行器中关键部位热防护问题的技术方案之一。研究真实条件下发汗冷却的流动换热规律，对下一代液体火箭发动机以及高超声速飞行器的发展具有重要意义。已有的发汗冷却研究多采用低温低速主流条件，不能直接反映真实工况下的发汗冷却规律。本文以真实发汗冷却应用环境为背景，搭建了Ma为3的超声速主流风洞实验台，采用纹影仪观测主流激波结构，红外热像仪测量发汗壁面温度信息，研究分析了颗粒直径为90μm的青铜、不锈钢粉末烧结多孔平板，烧结金属丝网多孔平板以及不锈钢烧结粉末多孔曲面结构的发汗冷却规律，同时通过数值模拟，分析了发汗出流和主流相互作用以及激波结构对发汗冷却影响的具体规律。研究结果表明，发汗冷却二次流注入能够有效减小壁面速度梯度，降低壁面温度；高固体骨架热导率会有助于热量向多孔冷端传递，提高冷却效率；冷却效率基本不随主流总温的变化而变化。针对超燃冲压发动机燃烧室支板热防护的具体问题，提出、设计并加工了不同结构的烧结金属粉末多孔支板结构，采用发汗冷却方式对其进行热防护。在高温高速风洞中的搭载实验结果表明此种结构能够有效对支板进行保护，并且通过对支板底部表面测量的温度分布可以看出，此种发汗冷却支板热防护方案存在很大的优化设计空间。最后利用解析解及数值计算相结合的方法，对发汗冷却多孔介质区域采用局部非热平衡模型进行了计算和分析，研究了冷端边界条件设置以及计算模型的简化和选择对多孔壁面温度的影响。研究结果表明：多孔区冷端边界处应考虑冲击换热的影响，否则在小雷诺数流动时会造成计算温度低于实际值；在求解发汗冷却解析解过程中所进行的常物性、忽略流体热扩散项假设会使得计算的压力、温度信息与实际情况有所偏差，尤其不能忽略热弥散效应对多孔介质内换热的影响；温度跳跃效应会使微多孔内流固相温度升高，速度滑移效应会使多孔壁面进出口边界压差减小。更多还原

【Abstract】 Thermal protection for aerospace vehicles has become a great challenge due to thehigh thermal loads experienced in liquid fuel rockets and scramjet chambers.Transpiration cooling is one of the most promising cooling techniques for protecting thesurface from ablation in the high temperature supersonic mainstream. Studies of thetranspiration cooling mechanism with real working conditions are very important for thedevelopment of liquid fuel rockets and super/hypersonic vehicles.Transpiration cooling studies using low velocities mainstream and temperaturescan not directly model the transpiration cooling characteristics for real workingconditions. This study used a supersonic wind tunnel with Mach number3toexperimentally investigate supersonic transpiration cooling. The wall temperatures weremeasured by an infrared camera and a schlieren system to observe the shock wavestructures. The flow and heat transfer mechanisms of sintered porous structures, wovenwire mesh structures and curved sintered porous structures were investigatedexperimentally and numerically. The results show that the velocity gradients near thesurface are significantly reduced by the coolant transpiration, with bronze porous walltemperatures lower than stainless steel wall temperature because of the higher thermalconductivity and that the cooling effectiveness did not change much with themainstream total temperature.A transpiration cooling scheme was developed for struts using a sintered metalporous medium provide thermal protection for struts in scramjet chambers.Transpiration cooling experiments using methane as the coolant with supersonic, hightemperature main flow demonstrated that this strut cooling scheme provides effectivethermal protection. The temperature distribution measured by thermocouples welded tothe bottom of the strut indicated that the system parameters can be optimized tosignificantly improve the cooling.In addition, analytical and numerical methods were used to calculate thetranspiration cooling temperature distribution in the porous media using the LocalThermal Non-Equilibrium model. The analytical model predicted the temperaturedistribution in the porous wall with different coolant entrance boundary conditions. Numerical simulations were used to evaluate the temperature changes caused by thesimplifications in the analytical solutions. Transpiration cooling using a microporousmedia was also numerically studied. The analytical solutions showed that theconvection heat transfer at the entrance should be considered with low Reynoldsnumber flow; otherwise, the solution will underpredict the temperatures. Comparisonsof the numerical results with the analytical solutions indicated that the material thermalproperties and the thermal diffusion term significantly affect the pressure andtemperature distributions in the porous matrix. The transpiration cooling simulations inthe microporous media showed that the wall temperature jump increased thetemperatures in the microporous wall while the velocity slip effect reduced the pressuredrops between the inlet and outlet surfaces.更多还原