飞行器迎风前缘逆喷与发汗防热机理及复杂流动算法研究
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摘要
逆喷发汗迎风前缘结构为高超声速远程飞行器提供了一种结构简单、可靠性高、成本低而效率高的主动热防护方法。这种前缘结构由层板组合成型，融发汗冷却功能和空气针减阻降热功能为一体。在正常飞行时应用层板冷却结构周向发汗满足高超声速远程飞行器的热防护要求；同时在顶部由层板结构形成的微型空气针产生逆向喷流，实现在高马赫数飞行时获得减少阻力降低热流强度的效果。该热防护方法的应用能够实现高超声速飞行器头锥长时间工作和可重复使用的目的。
本文围绕逆喷发汗前缘的热防护机理，主要针对两大防热功能展开研究，即逆向喷流热防护功能与层板发汗热防护功能，内容涉及外部流场与固壁热参数分布的研究与分析，并对现有计算方法进行改进与优化。
逆向喷流热防护功能方面，为了更详尽细致地了解逆向喷流热防护方法的工作机理，本文进行了数值方法研究，采用有限体积法，结合AUSMPW格式、MUSCL方法和LU-SGS方法编制了计算程序，通过与实验及文献算例对比验证了程序可靠性。应用模拟程序，数值模拟了超声速逆向喷流复杂流场，缜密对比分析了数值模拟结果。将总压比率和流量相结合，提出了新的参数表征逆向喷流的强度，探讨分析了新参数对流场特征，阻力系数及相对总传热量的影响效果，给出了该参数与流场特征，阻力系数及相对总传热量等参数的定性函数关系。准确获得再入飞行器绕流流场特性及物面传热量是获得逆喷发汗前缘热环境特性的重要基础，为分析、研究其整体结构热特性提供有效的数据支持。
在带逆向喷流的复杂超声速流场中，同时存在着高速流动区与低速流动区。由于低速流动区域存在低速效应，导致数值模拟计算收敛速度变慢，数值误差变大。引入预处理方法，可实现加速收敛与正确求解的目的。数值求解时，根据通量分裂格式与系统特征值的关系推导出应用预处理方法的系统下对应的通量分裂格式。为了将通量矢量分裂格式应用于预处理方法中，本文基于Roe格式提出了一种新的通量矢量分裂格式，并成功应用于预处理方法中。通过具体算例成功实现了低马赫数条件下数值计算的收敛加速与正确求解，验证了预处理方法的有效性与正确性。
层板发汗热防护功能方面，构建了层板发汗鼻锥前缘的物理外形；再针对外部热源的计算，本文采用了外部流场对壁面气动加热的工程计算方法及CFD方法，针对内部冷却机制建立了发汗鼻锥冷却槽道内冷却液分布模型，获得了槽道内冷却液的流动换热参数；在此基础上应用有限体积法计算鼻锥整体的热状态，获得了鼻锥壁面的温度分布以及各槽道冷却液温度分布。为分析鼻锥整体结构参数与冷却效果提供了分析途径与计算方法。
利用温度场计算方法计算鼻锥整体的热状态，获得了鼻锥壁面的温度分布，分析研究了层板发汗冷却对鼻锥的冷却效果，成功将最高壁面温度控制在材料的耐热温度以内，确保鼻锥在严重的气动加热环境下仍能保持在允许的温度范围内持续工作。并在此基础上讨论了层板冷却通道结构参数对冷却效果的影响，并分析了其影响效果的发生机理。最后将层板发汗与逆向喷流的热防护效果相结合，综合分析了整体热防护效果，验证了逆喷发汗前缘结构的有效性。
最后对逆喷发汗前缘结构的外流场进行了实验研究。首先对超声速静风洞实验系统的各个组成部分进行了介绍，然后由所研究的问题提出实验模型，并根据逆向喷流鼻锥绕流的流动特性设计了实验方案，最后将实验结果与数值计算进行了对比分析。通过高分辨率NPLS流场观测技术，能够清晰地观察到逆向喷流流场的复杂结构，包括钝体前端的弓形激波以及在喷流层的回流再附点附近形成再压缩激波。将实验结果与数值计算进行了对比分析，结果显示两者相当吻合，再次验证了数值计算的正确性和可靠性。
The opposing-jet and inspiration leading edge offers long-range hypersonic vehiclean active thermal protection method with simple structure, high reliability, low cost andhigh efficiency. The leading edge is built with plantlets, and has both transpirationcooling and opposing jet for thermal protection. When it works, the transpirationcooling structure will cool down the long-range hypersonic vehicle, and the opposing jetwill also reduce the drag while reducing the heat flux. At the same time, it can make thehypersonic vehicle work for a long time and reusable.
The work in this thesis is about the thermal protection mechanism of theopposing-jet and inspiration leading edge for hypersonic vehicle, which includes twoparts: one is opposing jet thermal protection function and another is transpirationcooling thermal protection function. The complex hypersonic flowfield and walltemperature are simulated and analyzed. And the simulation method is improved withpreconditioning method.
The opposing jet thermal protection function is one of the important sides. In orderto know how it works, numerical simulation of hypersonic flow field with opposing jetis performed. The3D computational code to solve Navier-Stoke function has beenestablished by using AUSMPW scheme, MUSCL method, LU-SGS method and finitevolume method. The computational results are consistent with those of the referencesand the code has been validated using experimental data. By the use of the code, thecomplex hypersonic flow field with opposing jet is obtained and analyzed. To study theeffect of the intensity of opposing jet more reasonably, a new parameter has beendefined by combining the flux and the total pressure ratio. The study shows that thesame shock wave position, drag coefficient and total heat load can be obtained with thesame new parameter with different fluxes and the total pressures, and the new parameterhas qualitative relationship with the flowfield coefficients. As a base to calculate thewall temperature of the opposing-jet and inspiration leading edge, the flowfieldcharacteristic and heat load should be obtained correctly, which will support the wholethermal analysis.
In the hypersonic flow field with opposing jet, there are both high speed area andlow speed area. In the low speed area, there is low speed effect, which results inconvergence deterioration and incorrect solution. However, we can introducepreconditioning method in order to accelerate the convergence of the steady solutionand obtain the numerical solution correctly. When the low speed problem is solvednumerically, the flux splitting schemes with preconditioning is deduced based on therelationship between the flux splitting schemes and system eigenvalues. In order to useflux vector splitting scheme in the preconditioning method, a new scheme is proposed based on the Roe scheme and can be used in the preconditioning method. The numericaltests of a low speed cases succeed with convergence acceleration and correct solution,which validates the preconditioning method.
The plantlet transpiration cooling thermal protection function is another importantside. The physical shape of the plantlet transpiration cooling nose is built. In order toanalysis the thermal characteristic of the whole nose, both heating and cooling wayshould be considered. The methods of calculating the aerodynamic heating, such asengineering method and CFD method, are introduced. For cooling way, a distributionmodel of coolant is proposed for the later temperature calculation. With the heating andcooling data, the whole nose thermal state can be obtained with wall temperaturedistribution and coolant temperature distribution in the grooves by use of Finite VolumeElement.
With the method of obtaining the thermal state of the whole nose, the transpirationcooling effect is studied. It is showed that the transpiration cooling can keep the highestwall temperature of the nose within the working temperature range of the material,which makes sure the nose keep working within the allowed temperature with seriousaerodynamic heating. The cooling effect with the plantlet cooling groove structure isdiscussed and the physical mechanism is analyzed. Combined with opposing jet thermalprotection method, the plantlet nose will work better. The synthetical thermal analysisof the opposing jet thermal protection function and the transpiration cooling thermalprotection function testifies the validity of the opposing-jet and inspiration leading edge.
The experiment on flowfield around the opposing-jet and inspiration leading edgehas been conducted in the supersonic wind tunnel. The high-definitionNanoparticle-based Planar Laser Scattering (NPLS) is used to observe the flowfield.The experiment is designed based on the characteristic of the supersonic flow withopposing jet. By NPLS, the complex structure of the flowfield with opposing jet can beobserved in detail, including the bowl shock wave in front of the nose and therecompression shock wave near the reattachment of the jet layer. It is showed that theexperiment results are consistent with the calculation results, which validates thecalculation method again.
本文围绕逆喷发汗前缘的热防护机理，主要针对两大防热功能展开研究，即逆向喷流热防护功能与层板发汗热防护功能，内容涉及外部流场与固壁热参数分布的研究与分析，并对现有计算方法进行改进与优化。
逆向喷流热防护功能方面，为了更详尽细致地了解逆向喷流热防护方法的工作机理，本文进行了数值方法研究，采用有限体积法，结合AUSMPW格式、MUSCL方法和LU-SGS方法编制了计算程序，通过与实验及文献算例对比验证了程序可靠性。应用模拟程序，数值模拟了超声速逆向喷流复杂流场，缜密对比分析了数值模拟结果。将总压比率和流量相结合，提出了新的参数表征逆向喷流的强度，探讨分析了新参数对流场特征，阻力系数及相对总传热量的影响效果，给出了该参数与流场特征，阻力系数及相对总传热量等参数的定性函数关系。准确获得再入飞行器绕流流场特性及物面传热量是获得逆喷发汗前缘热环境特性的重要基础，为分析、研究其整体结构热特性提供有效的数据支持。
在带逆向喷流的复杂超声速流场中，同时存在着高速流动区与低速流动区。由于低速流动区域存在低速效应，导致数值模拟计算收敛速度变慢，数值误差变大。引入预处理方法，可实现加速收敛与正确求解的目的。数值求解时，根据通量分裂格式与系统特征值的关系推导出应用预处理方法的系统下对应的通量分裂格式。为了将通量矢量分裂格式应用于预处理方法中，本文基于Roe格式提出了一种新的通量矢量分裂格式，并成功应用于预处理方法中。通过具体算例成功实现了低马赫数条件下数值计算的收敛加速与正确求解，验证了预处理方法的有效性与正确性。
层板发汗热防护功能方面，构建了层板发汗鼻锥前缘的物理外形；再针对外部热源的计算，本文采用了外部流场对壁面气动加热的工程计算方法及CFD方法，针对内部冷却机制建立了发汗鼻锥冷却槽道内冷却液分布模型，获得了槽道内冷却液的流动换热参数；在此基础上应用有限体积法计算鼻锥整体的热状态，获得了鼻锥壁面的温度分布以及各槽道冷却液温度分布。为分析鼻锥整体结构参数与冷却效果提供了分析途径与计算方法。
利用温度场计算方法计算鼻锥整体的热状态，获得了鼻锥壁面的温度分布，分析研究了层板发汗冷却对鼻锥的冷却效果，成功将最高壁面温度控制在材料的耐热温度以内，确保鼻锥在严重的气动加热环境下仍能保持在允许的温度范围内持续工作。并在此基础上讨论了层板冷却通道结构参数对冷却效果的影响，并分析了其影响效果的发生机理。最后将层板发汗与逆向喷流的热防护效果相结合，综合分析了整体热防护效果，验证了逆喷发汗前缘结构的有效性。
最后对逆喷发汗前缘结构的外流场进行了实验研究。首先对超声速静风洞实验系统的各个组成部分进行了介绍，然后由所研究的问题提出实验模型，并根据逆向喷流鼻锥绕流的流动特性设计了实验方案，最后将实验结果与数值计算进行了对比分析。通过高分辨率NPLS流场观测技术，能够清晰地观察到逆向喷流流场的复杂结构，包括钝体前端的弓形激波以及在喷流层的回流再附点附近形成再压缩激波。将实验结果与数值计算进行了对比分析，结果显示两者相当吻合，再次验证了数值计算的正确性和可靠性。
The opposing-jet and inspiration leading edge offers long-range hypersonic vehiclean active thermal protection method with simple structure, high reliability, low cost andhigh efficiency. The leading edge is built with plantlets, and has both transpirationcooling and opposing jet for thermal protection. When it works, the transpirationcooling structure will cool down the long-range hypersonic vehicle, and the opposing jetwill also reduce the drag while reducing the heat flux. At the same time, it can make thehypersonic vehicle work for a long time and reusable.
The work in this thesis is about the thermal protection mechanism of theopposing-jet and inspiration leading edge for hypersonic vehicle, which includes twoparts: one is opposing jet thermal protection function and another is transpirationcooling thermal protection function. The complex hypersonic flowfield and walltemperature are simulated and analyzed. And the simulation method is improved withpreconditioning method.
The opposing jet thermal protection function is one of the important sides. In orderto know how it works, numerical simulation of hypersonic flow field with opposing jetis performed. The3D computational code to solve Navier-Stoke function has beenestablished by using AUSMPW scheme, MUSCL method, LU-SGS method and finitevolume method. The computational results are consistent with those of the referencesand the code has been validated using experimental data. By the use of the code, thecomplex hypersonic flow field with opposing jet is obtained and analyzed. To study theeffect of the intensity of opposing jet more reasonably, a new parameter has beendefined by combining the flux and the total pressure ratio. The study shows that thesame shock wave position, drag coefficient and total heat load can be obtained with thesame new parameter with different fluxes and the total pressures, and the new parameterhas qualitative relationship with the flowfield coefficients. As a base to calculate thewall temperature of the opposing-jet and inspiration leading edge, the flowfieldcharacteristic and heat load should be obtained correctly, which will support the wholethermal analysis.
In the hypersonic flow field with opposing jet, there are both high speed area andlow speed area. In the low speed area, there is low speed effect, which results inconvergence deterioration and incorrect solution. However, we can introducepreconditioning method in order to accelerate the convergence of the steady solutionand obtain the numerical solution correctly. When the low speed problem is solvednumerically, the flux splitting schemes with preconditioning is deduced based on therelationship between the flux splitting schemes and system eigenvalues. In order to useflux vector splitting scheme in the preconditioning method, a new scheme is proposed based on the Roe scheme and can be used in the preconditioning method. The numericaltests of a low speed cases succeed with convergence acceleration and correct solution,which validates the preconditioning method.
The plantlet transpiration cooling thermal protection function is another importantside. The physical shape of the plantlet transpiration cooling nose is built. In order toanalysis the thermal characteristic of the whole nose, both heating and cooling wayshould be considered. The methods of calculating the aerodynamic heating, such asengineering method and CFD method, are introduced. For cooling way, a distributionmodel of coolant is proposed for the later temperature calculation. With the heating andcooling data, the whole nose thermal state can be obtained with wall temperaturedistribution and coolant temperature distribution in the grooves by use of Finite VolumeElement.
With the method of obtaining the thermal state of the whole nose, the transpirationcooling effect is studied. It is showed that the transpiration cooling can keep the highestwall temperature of the nose within the working temperature range of the material,which makes sure the nose keep working within the allowed temperature with seriousaerodynamic heating. The cooling effect with the plantlet cooling groove structure isdiscussed and the physical mechanism is analyzed. Combined with opposing jet thermalprotection method, the plantlet nose will work better. The synthetical thermal analysisof the opposing jet thermal protection function and the transpiration cooling thermalprotection function testifies the validity of the opposing-jet and inspiration leading edge.
The experiment on flowfield around the opposing-jet and inspiration leading edgehas been conducted in the supersonic wind tunnel. The high-definitionNanoparticle-based Planar Laser Scattering (NPLS) is used to observe the flowfield.The experiment is designed based on the characteristic of the supersonic flow withopposing jet. By NPLS, the complex structure of the flowfield with opposing jet can beobserved in detail, including the bowl shock wave in front of the nose and therecompression shock wave near the reattachment of the jet layer. It is showed that theexperiment results are consistent with the calculation results, which validates thecalculation method again.
引文
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