基于冲击冷却的翼前缘主动冷却结构构建方法及其换热特性研究

The wing leading edge of hypersonic vehicle works in a complex aerothermal heating environment. Its temperature is extremely high and the distribution of which is uneven. The temperature of stagnation region is ultimate high which forms hot spot. Accordingly thermal stress is generated from temperature gradient and accompanies with aerodynamics force, which appeals demand for mechanical ability as well as thermal endurance ability of the leading edge structure. Active cooling is an effective way to cool local high temperature region and promote temperature uniformity of the whole structure, which can satisfy the need of thermal protection system of wing leading edge. The present study focuses on the active cooling structure of leading edge, based on impingement cooling which has outstanding cooling effect and impressive thermal protection potential, with taking thermal load, pneumatic force, construction thermal stress into consideration, aiming at design a effective internal cooling structure to fulfill the need of hypersonic environment. Method of theoretics analysis, numerical simulation and experimental validation are adopted to design proper active cooling configuration to meet the demand of heat and pneumatic load for the leading edge. Numerical model is built up to investigate internal flow, heat transfer characteristics and stress distribution of the active cooling structure with progressive coupling method. Optimized precept of the given structure can be brought forward on the base of calculation results. Experiments are conducted to conclude flow and heat transfer characteristics of the optimized active cooling structure, as well as to detect the mechanism of enhancing heat transfer. Optimised combination of low temperature and high uniformity is present to improve the design method of active cooling structure.

高超声速翼前缘处于复杂的气动热环境中，整体温度高且温度分布不均匀，在驻点区有局部热斑存在，由此产生的附加热应力与气动载荷叠加，从热承载和机械承载两方面对翼前缘结构提出了严格要求。主动冷却方法能有效降低局部高温，提高结构温度均匀性，符合翼前缘热防护系统需求。本项目以翼前缘为研究对象，采用理论分析、数值模拟和试验验证方法，充分考虑热载荷、气动力和结构热应力等多项影响因素，基于具有工程实践经验和深厚挖掘潜力的冲击冷却方法，设计满足翼前缘热载荷和机械承载要求的主动冷却结构。建立翼前缘计算模型，发展耦合条件下的计算方法，研究翼前缘温度分布规律，以及主动冷却结构的流动、换热特性和应力分布规律，提出结构优化方案。针对优化结构开展试验验证研究，归纳该结构的流阻特性和换热特性规律，提炼实验经验关联式，揭示强化换热机理，提出降低局部温度与提高温度均匀性的优化组合，完善翼前缘主动冷却结构的构建方法。

高超声速翼前缘处于复杂的气动热环境中，整体温度高且温度分布不均匀，在驻点区有局部热斑存在，由此产生的附加热应力与气动载荷叠加，从热承载和机械承载两方面对翼前缘结构提出了严格要求。主动冷却方法能有效降低局部高温，提高结构温度均匀性，符合翼前缘热防护系统需求。本项目以翼前缘为研究对象，采用理论分析、数值模拟和试验验证方法，基于具有工程实践经验和深厚挖掘潜力的冲击冷却方法，设计满足翼前缘需求的主动冷却结构。本研究设计了六边形翼型，采用工程方法估算了高超声速条件下的翼面热环境，利用结构化网格建立翼型计算模型，在对湍流模型进行高超声速条件下的可压缩修正后开展数值计算；翼型内部主动冷却结构以前缘梯形冲击腔为基础，辅以内部蜿蜒通道，经优化后有效提高了换热性能，降低了压力损失和温度不均匀度，初步实现了设计目的；在常温条件下对翼前缘冲击腔内换热特性开展了实验研究，获得了基于射流雷诺数和横流强度的冲击腔内的换热准则方程，并验证了前缘冲击冷却结构对换热的提升能力为光滑管的2-3倍；在热冲击平台上开展温度-时间响应特性实验，获得的实验结果对后续数值计算中热边界条件的设定具有很强的指导性。