迎风凹腔与逆向喷流组合强化防热结构复杂流场和传热特性研究
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摘要
热防护系统是高速飞行器设计与制造的关键技术之一，它直接关系到飞行器的飞行安全。由于高速飞行器飞行环境的恶劣和复杂，使得飞行器对热防护系统的要求很高，热防护系统设计一直是一项极富挑战性的前沿课题。
本文针对高速飞行器的热防护系统设计展开研究。建立了超声速、高超声速飞行器头部复杂流场的数值计算模型，通过流场显影实验结合公开文献中的实验结果对数值方法进行了验证。在对独立的逆向喷流防热方法以及迎风凹腔防热方法开展数值研究的基础上，提出了将两者结合应用，构建迎风凹腔与逆向喷流组合热防护系统的防热设计方案，数值计算证明了该设计方案的有效性。
数值研究了逆向喷流热防护系统对飞行器鼻锥流场、气动受力与气动热性能的影响。讨论了逆向喷流热防护系统的攻角特性，得到了该热防护系统在相应条件下的使用攻角范围。当飞行攻角为10°，逆向喷流鼻锥迎风面上的最大热流值已经接近没有热防护系统的鼻锥最大热流。可以认为在本文研究的飞行条件下，当飞行攻角为10°，逆向喷流热防护系统不再适用。
对独立迎风凹腔热防护系统对飞行器鼻锥流场、气动受力与气动加热性能的影响进行了数值模拟。得到了带迎风凹腔鼻锥的全壁面热流分布并讨论了不同的迎风凹腔几何参数选择对鼻锥流场、气动受力以及结构冷却效果的影响。迎风凹腔结构能够有效的对高超声速飞行器的头部进行冷却，尤其是驻点区域附近，冷却效果很好。凹腔越深，飞行器鼻锥外壁面的气动加热热流值越低。鼻锥表面的最大热流并不在尖锐唇缘的顶点，而是位于凹腔的侧壁面接近唇缘的位置，凹腔的深度变化对最大热流的出现位置影响很小。外壁面的最大热流也位于壁面唇缘顶点之后。虽然整个鼻锥气动加热的极大值出现在凹腔侧壁面上，但是凹腔侧壁其他位置的热流普遍较低。除了很浅的凹腔（凹腔的长度与直径之比小于等于0.5），凹腔底面的热流值都非常小，基本可以忽略。
提出并探讨了迎风凹腔与逆向喷流组合防热结构的设计方案及其冷却效果。对该组合防热结构的流场、气动力以及防热效能进行了数值分析。分别对组合热防护系统中，逆向喷流总压、逆喷马赫数以及迎风凹腔的深度与直径的选择对组合热防护系统鼻锥的影响进行了讨论。得到了不同逆喷工况、不同凹腔几何参数条件下，组合热防护系统的防热效能。迎风凹腔与逆向喷流组合热防护系统能够有效地对高速飞行器头部进行冷却，大大削弱鼻锥物面的气动加热。回流区域在鼻锥冷却过程中扮演着关键的角色。在相同的逆向喷流总压条件下，逆向喷流的速度越高，逆喷流量越大，鼻锥外壁面的冷却效果越好，采用组合防热结构鼻锥的阻力越小。在相同的逆向喷流马赫数条件下，飞行器鼻锥的气动阻力随逆喷总压提高单调下降，而外壁面热流的变化趋势较为复杂。在鼻锥流动条件（包含自由来流与逆向喷流）不变的情况下，组合热防护结构凹腔直径不变时，凹腔越深，鼻锥外壁面热流越小，组合结构的冷却效果越好；凹腔深度的变化对鼻锥气动阻力的影响不大。当凹腔长度不变，对于鼻锥外壁面热流，存在一个最优的凹腔直径选择，在本文的研究中，当凹腔直径取6mm时有最小的鼻锥外表面热流分布。
最后，使用国防科技大学基于纳米粒子的平面激光流动成像系统（Nanoparticle-based Planar Laser Scattering，NPLS）对单纯迎风凹腔、逆向喷流以及两者的组合结构三种热防护结构进行了流场成像观测。得到了清晰的采用三种不同热防护方案鼻锥的流场图像，包括使用单纯逆向喷流和组合结构时，由自由来流与逆向流动作用形成的鼻锥前端的脱体激波以及再压缩波等复杂波系结构；使用单纯迎风凹腔时，鼻锥前端弓形脱体激波的往复震荡。进一步的，将实验结果与数值计算的结果相对比，两者非常吻合，再次验证了数值计算的正确性。
Thermal protection system (TPS) is a key technique of the high speed vehicles’design and manufacture, it is connected with the safety of the aerocraft directly. Thedesign of TPS is always a challenging work for the high requirement from the aerocraftby its inclement and complex aviation environment.
The present dissertation is about research of thermal protection system design forhigh speed vehicles. The numerical simulation mode of the flow field of surpersonic andhypersonic vehicle nose-tip is established, and this numerical method is validated byflow imaging experiment and experimental date in the open literature. Based on thenumerical simulation of the single opposing jet and single forward-facing cavity coolingmethod, a kind of the forward-facing cavity and opposing jet combined TPS is designedand investigated numerically.
The influence of the opposing jet TPS on the flow field, aerodynamic force, andsurface aerodynamic heating of the vehicle nose-tip was investigated numerically. Theproperties of attack angle for opposing jet TPS was discussed, the useful attach anglerange of opposing jet TPS was obtained. When the attack angle reaches10, the heatflux on the windward generatrix is close to the maximal heat flux on the wall surface ofthe nose-tip without thermal protection system, thus the thermal protection is failure.
A numerical study on the effect of forward-facing cavity upon aerodynamicheating on the hypersonic vehicle nose was conducted. The flow field parameters, heatflux distribution along the outer body surface and the cavity wall were obtained, thecooling effect of the forward-facing cavity with different dimension was analyzed. Theresults show that the forward-facing cavity configuration dose well in cooling the noseof hypersonic vehicles especially at the stagnation point area. The deeper the cavity is,the smaller the heat flux is. The maximal heat flux dose not locates at the peak of thesharp lip, the maximal heat flux of the outer body surface and upper wall of the cavityare all behind the lip peak. The cavity length has little effect on the location of themaximal heat flux. In sum, there is a low heat flux on the upper wall, but the maximalaerodynamic heating point is located on it. There is a very low surface heat flux alongthe base wall of the cavity except the cavity is shallow enough.
The forward-facing cavity and opposing jet combined TPS was investigatedprimarily. The flow field parameters, aerodynamic force and surface heat fluxdistribution of this combined TPS nose-tip were obtained by numerical method. Theinfluence of the opposing jet stagnation pressure, Mach number and cavity length,diameter on the thermal protection efficiency of the TPS was discussed respectively.The performance parameter of the combined TPS with different opposing jet workcondition and cavity geometry parameter were obtained. This kind of combined TPS has an excellent impact on cooling the nose-tip of high speed vehicles. The recirculationregion plays an pivotal role for the reduction of heat flux. Under the same opposing jetstagnation pressure condition, with the opposing jet speed increasing, the coolingefficiency of the combined TPS is improved and the aerodynamic resistance isdecreasing. Under the same jet Mach number condition, with the stagnation pressure ofopposing jet decreases, the aerodynamic resistance is increasing but the heat flux alongouter body surface of the nose-tip does not increase monotonically. Under the sameflow condition (include the free stream and opposing jet flow), with a constant cavitydiameter, the combined TPS with larger length cavity has the higher heat flux reduction.The cavity length has little influence on the aerodynamic force. Both of theaerodynamic heating and the force do not variate monotonically with the cavitydiameter. There is an optimal choice of the cavity diameter for the aerodynamic heatingreducing of the nose-tip, in this researching, the aerodynamic heating has the minimumdistribution when the cavity diameter is6mm.
The high-definition Nanoparticle-based Planar Laser Scattering (NPLS) is used toobserve the flowfield of nose-tip with the three different thermal protectionconfiguration, forward-facing cavity, opposing jet and their combined configuration.The complex structure of the flowfield with different thermal protection method can beobserved in detail, including the bowl shock wave in front of the nose and therecompression shock wave when the thermal protection method contains the opposingjet, and the bowl shock wave oscillation when the thermal protection method containsforward-facing cavity only. At last, these experiment images are compared with thecalculation results, there is a good agreement, which validates the calculation methodagain.
本文针对高速飞行器的热防护系统设计展开研究。建立了超声速、高超声速飞行器头部复杂流场的数值计算模型，通过流场显影实验结合公开文献中的实验结果对数值方法进行了验证。在对独立的逆向喷流防热方法以及迎风凹腔防热方法开展数值研究的基础上，提出了将两者结合应用，构建迎风凹腔与逆向喷流组合热防护系统的防热设计方案，数值计算证明了该设计方案的有效性。
数值研究了逆向喷流热防护系统对飞行器鼻锥流场、气动受力与气动热性能的影响。讨论了逆向喷流热防护系统的攻角特性，得到了该热防护系统在相应条件下的使用攻角范围。当飞行攻角为10°，逆向喷流鼻锥迎风面上的最大热流值已经接近没有热防护系统的鼻锥最大热流。可以认为在本文研究的飞行条件下，当飞行攻角为10°，逆向喷流热防护系统不再适用。
对独立迎风凹腔热防护系统对飞行器鼻锥流场、气动受力与气动加热性能的影响进行了数值模拟。得到了带迎风凹腔鼻锥的全壁面热流分布并讨论了不同的迎风凹腔几何参数选择对鼻锥流场、气动受力以及结构冷却效果的影响。迎风凹腔结构能够有效的对高超声速飞行器的头部进行冷却，尤其是驻点区域附近，冷却效果很好。凹腔越深，飞行器鼻锥外壁面的气动加热热流值越低。鼻锥表面的最大热流并不在尖锐唇缘的顶点，而是位于凹腔的侧壁面接近唇缘的位置，凹腔的深度变化对最大热流的出现位置影响很小。外壁面的最大热流也位于壁面唇缘顶点之后。虽然整个鼻锥气动加热的极大值出现在凹腔侧壁面上，但是凹腔侧壁其他位置的热流普遍较低。除了很浅的凹腔（凹腔的长度与直径之比小于等于0.5），凹腔底面的热流值都非常小，基本可以忽略。
提出并探讨了迎风凹腔与逆向喷流组合防热结构的设计方案及其冷却效果。对该组合防热结构的流场、气动力以及防热效能进行了数值分析。分别对组合热防护系统中，逆向喷流总压、逆喷马赫数以及迎风凹腔的深度与直径的选择对组合热防护系统鼻锥的影响进行了讨论。得到了不同逆喷工况、不同凹腔几何参数条件下，组合热防护系统的防热效能。迎风凹腔与逆向喷流组合热防护系统能够有效地对高速飞行器头部进行冷却，大大削弱鼻锥物面的气动加热。回流区域在鼻锥冷却过程中扮演着关键的角色。在相同的逆向喷流总压条件下，逆向喷流的速度越高，逆喷流量越大，鼻锥外壁面的冷却效果越好，采用组合防热结构鼻锥的阻力越小。在相同的逆向喷流马赫数条件下，飞行器鼻锥的气动阻力随逆喷总压提高单调下降，而外壁面热流的变化趋势较为复杂。在鼻锥流动条件（包含自由来流与逆向喷流）不变的情况下，组合热防护结构凹腔直径不变时，凹腔越深，鼻锥外壁面热流越小，组合结构的冷却效果越好；凹腔深度的变化对鼻锥气动阻力的影响不大。当凹腔长度不变，对于鼻锥外壁面热流，存在一个最优的凹腔直径选择，在本文的研究中，当凹腔直径取6mm时有最小的鼻锥外表面热流分布。
最后，使用国防科技大学基于纳米粒子的平面激光流动成像系统（Nanoparticle-based Planar Laser Scattering，NPLS）对单纯迎风凹腔、逆向喷流以及两者的组合结构三种热防护结构进行了流场成像观测。得到了清晰的采用三种不同热防护方案鼻锥的流场图像，包括使用单纯逆向喷流和组合结构时，由自由来流与逆向流动作用形成的鼻锥前端的脱体激波以及再压缩波等复杂波系结构；使用单纯迎风凹腔时，鼻锥前端弓形脱体激波的往复震荡。进一步的，将实验结果与数值计算的结果相对比，两者非常吻合，再次验证了数值计算的正确性。
Thermal protection system (TPS) is a key technique of the high speed vehicles’design and manufacture, it is connected with the safety of the aerocraft directly. Thedesign of TPS is always a challenging work for the high requirement from the aerocraftby its inclement and complex aviation environment.
The present dissertation is about research of thermal protection system design forhigh speed vehicles. The numerical simulation mode of the flow field of surpersonic andhypersonic vehicle nose-tip is established, and this numerical method is validated byflow imaging experiment and experimental date in the open literature. Based on thenumerical simulation of the single opposing jet and single forward-facing cavity coolingmethod, a kind of the forward-facing cavity and opposing jet combined TPS is designedand investigated numerically.
The influence of the opposing jet TPS on the flow field, aerodynamic force, andsurface aerodynamic heating of the vehicle nose-tip was investigated numerically. Theproperties of attack angle for opposing jet TPS was discussed, the useful attach anglerange of opposing jet TPS was obtained. When the attack angle reaches10, the heatflux on the windward generatrix is close to the maximal heat flux on the wall surface ofthe nose-tip without thermal protection system, thus the thermal protection is failure.
A numerical study on the effect of forward-facing cavity upon aerodynamicheating on the hypersonic vehicle nose was conducted. The flow field parameters, heatflux distribution along the outer body surface and the cavity wall were obtained, thecooling effect of the forward-facing cavity with different dimension was analyzed. Theresults show that the forward-facing cavity configuration dose well in cooling the noseof hypersonic vehicles especially at the stagnation point area. The deeper the cavity is,the smaller the heat flux is. The maximal heat flux dose not locates at the peak of thesharp lip, the maximal heat flux of the outer body surface and upper wall of the cavityare all behind the lip peak. The cavity length has little effect on the location of themaximal heat flux. In sum, there is a low heat flux on the upper wall, but the maximalaerodynamic heating point is located on it. There is a very low surface heat flux alongthe base wall of the cavity except the cavity is shallow enough.
The forward-facing cavity and opposing jet combined TPS was investigatedprimarily. The flow field parameters, aerodynamic force and surface heat fluxdistribution of this combined TPS nose-tip were obtained by numerical method. Theinfluence of the opposing jet stagnation pressure, Mach number and cavity length,diameter on the thermal protection efficiency of the TPS was discussed respectively.The performance parameter of the combined TPS with different opposing jet workcondition and cavity geometry parameter were obtained. This kind of combined TPS has an excellent impact on cooling the nose-tip of high speed vehicles. The recirculationregion plays an pivotal role for the reduction of heat flux. Under the same opposing jetstagnation pressure condition, with the opposing jet speed increasing, the coolingefficiency of the combined TPS is improved and the aerodynamic resistance isdecreasing. Under the same jet Mach number condition, with the stagnation pressure ofopposing jet decreases, the aerodynamic resistance is increasing but the heat flux alongouter body surface of the nose-tip does not increase monotonically. Under the sameflow condition (include the free stream and opposing jet flow), with a constant cavitydiameter, the combined TPS with larger length cavity has the higher heat flux reduction.The cavity length has little influence on the aerodynamic force. Both of theaerodynamic heating and the force do not variate monotonically with the cavitydiameter. There is an optimal choice of the cavity diameter for the aerodynamic heatingreducing of the nose-tip, in this researching, the aerodynamic heating has the minimumdistribution when the cavity diameter is6mm.
The high-definition Nanoparticle-based Planar Laser Scattering (NPLS) is used toobserve the flowfield of nose-tip with the three different thermal protectionconfiguration, forward-facing cavity, opposing jet and their combined configuration.The complex structure of the flowfield with different thermal protection method can beobserved in detail, including the bowl shock wave in front of the nose and therecompression shock wave when the thermal protection method contains the opposingjet, and the bowl shock wave oscillation when the thermal protection method containsforward-facing cavity only. At last, these experiment images are compared with thecalculation results, there is a good agreement, which validates the calculation methodagain.
引文
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