导向器叶片冲击冷却结构设计及传热特性研究
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摘要
提高涡轮入口温度对改善航空发动机推重比和热效率有重要意义，但也给发动机的正常运行带来一系列严重问题。目前入口燃气温度的提高速度远远高于叶片材料耐高温性能的发展速度，在继续发展耐高温材料的同时，必须采取有效的冷却技术和热防护措施来保证涡轮叶片在高温环境下安全可靠的工作。准确预测涡轮叶片的温度场是延长叶片工作寿命、提高冷却效率的关键问题，随着计算流体力学的不断发展，气热耦合方法已经成为航空发动机冷却结构设计的重要工具。本文借助气热耦合方法，重点针对小尺寸高压涡轮导向器叶片进行冲击冷却结构设计，完成气动和传热的三维耦合分析，为实际涡轮导向器叶片冷却结构设计提供参考。
叶轮机械中边界层转捩流动普遍存在，其直接影响叶片壁面附近传热过程，本文首先研究了如何准确预测边界层转捩流动问题，选择Transition k-kl-转捩模型对平板边界层有无压力梯度的跨越转捩实验进行数值计算，并对边界层转捩前后流动状态以及影响转捩流动的边界条件进行研究，计算结果表明：Transition k-kl-转捩模型对有无压力梯度的平板跨越转捩表现良好，压力梯度对边界层转捩流动有较大影响，顺压梯度有助于稳定边界层流动，延缓转捩的发生，而逆压梯度则容易使边界层流动失稳，促使转捩提前发生。
本文进一步对NASA-MarkII叶片传热实验进行气热耦合数值计算，研究了各湍流模型的计算特点以及它们对转捩流动的识别能力，分析了叶片表面边界层流动状态对涡轮叶片传热的影响，发现Transition k-kl-转捩模型相比其它全湍流模型，能够更准确预测附面层内的分离转捩状况。由于转捩前期采用层流动能来描述扰动的发展，避免了使用含来流湍流度的经验公式所带来的计算层流误差，引入“分裂机制”来描述层流与湍流脉动间的相互作用，并且在旁路转捩和自然转捩源项模块中加入了Tollmien-Schlichting波的影响，对强激波后的温度计算相比常用的间歇因子转捩模型与实验值更吻合。最后采用变湍流普朗特数方法对Transition k-kl-转捩模型的热涡流扩散系数项进行修正，保证湍流普朗特数从边界层到主流区光滑过渡，使得对流换热系数计算值与实验值更加接近。
选用上述转捩模型进行数值计算，得出平板单孔冲击冷却结果，并与三维单孔冲击冷却分析解以及实验关联式进行对比，同时对模型网格无关性进行计算，验证数值结果的准确性，得到单孔冲击冷却的相应规律，对于理解冲击流动的流场和冲击冷却的机理有很大帮助。再利用已确定的数值计算方法，分析阵列射流不同的几何参数和物理参数对冲击冷却传热和流动的影响，与实验关联式进行对比，得到平板阵列射流的传热规律，为接下来导向器叶片冷却结构设计提供依据。
实验研究不能完全仿照航空发动机的实际工作条件，利用数值计算的优势研究涡轮叶片的实际工作状态具有非常重要的工程意义，接下来本文对导向器叶片进行冲击冷却结构设计，首先计算得出未加冷却时导向器叶片的温度场分布，发现叶片最高温度存在于前缘外表面中径处，再根据具体温度分布进行导向器叶片气冷结构设计，完成气动和传热的三维耦合分析，实现在保持冷却气体流量不变的条件下提高冷却效果，降低材料的性能要求。当导流片孔数和孔径满足关系式(4n-1)D=叶高（mm）时，在导向器叶片叶高方向上能满足冷却要求。以结构简单为原则，选择导流片每列冲击孔轴向距离为X/d=5~7时，能够满足中弦区冷却要求。前缘区域要想满足要求，所设计的冷却结构前缘至少须布置两列冲击孔。冲击间距较小时，冲击间距变化对叶片的内、外表面平均温度影响不大，但随冲击间距的减小，内、外表面最高温度会降低，而最低温度则会升高。单纯从冷却效果看，采用小间距为宜。与开孔排气结构相比，尾缘劈缝排气结构可以有效改善尾缘的冷却效果，且减小了总压损失。
本文最后应用热弹耦合方法对温度和冲击载荷作用的导向器叶片进行了应力分析，探讨了网格数量对计算结果的影响。结果发现：总变形量最大值出现在结构强度较低的叶片中弦位置，等效应力最大值出现在叶根和叶尖位置，分析认为是由于轮毂、轮缘给定固定支撑的约束条件，叶体由于热应力产生的形变，均会传递到两个端壁面上产生压缩效应。
It is important to improve aero turbine engine thermal efficiency and thrust-to-weightratio by increasing turbine inlet temperature, which also brings a series of serious problemsfor the normal operation of the engine. Now the inlet temperature increases much faster thanthe development of the temperature resistance of the blade material. The efficient coolingtechnology and the thermal protective measures must be taken into consideration to ensure theturbine blades working safely and reliably at a high temperature, in addition to continuing toimprove high temperature materials resistance. The precise heat transfer analysis of turbineblades is essential to improve the cooling efficiency and extend the operating life of the blades.With the incessant maturity of computational fluid dynamics, the conjugate heat transfermethodology has gradually become a prevailing tools in the design process of aero turbineengine. With the help of numerical simulation of conjugate heat transfer methodology, theimpingement cooling structure which is focused on the small size high pressure turbine guideblade, has been designed, and then the three-dimensional aerodynamics analysis and the heattransfer coupled analysis of this structure have been completed, which can provide referencefor the cooling structure design of the actual engine guide blades.
The boundary layer transition flow is widespread in turbo machinery, and it directlyaffects the heat transfer process of the turbine blade near the wall. The problem of how toaccurately predict the boundary layer transition flow was firstly studied, then the Transitionk-kl-model was selected to simulate boundary layer transition on the flat plate, and then theflow state before and after the boundary layer transition was in-depth studied. The resultsshow that: Transition k-kl-model has a good performance on the plate bypass transitionwith or without pressure gradient, and the pressure gradient has influence to the boundarylayer transition flow. The favor pressure gradient can stabilize boundary layer flow, postponethe onset location of transition. Otherwise the adverse pressure gradient can induce theboundary layer separation, and pre-act transition occurs.
The conjugate heat transfer methodology has been employed to NASA-MarkII turbineguide vane in this dissertation. By discussing the calculation characteristics of the turbulencemodel and the ability to identifying transitional flows, the influence of the blade surfaceboundary layer flow on turbine blade heat transfer was analyzed and corrected. Compared to all the other turbulence models, the transition model can predict the separation transition stateof the boundary layer more accurately. As the transition model uses laminar kinetic energy todescribe the development of the disturbance, so this model can avoid the errors introduced bythe empirical formula, which is related with the inflow turbulence intensity. This modelintroduces the "split system" to describe the interaction between the laminar and the turbulentfluctuation, besides Tollmien-Schlichting wave is added into the bypass transition and thenatural transition source term. The result of temperature calculation behind the strong shock ismore consistent with the experimental value, compared to that of the common models withthe intermittent transition factor. Finally, the mutative turbulent Prandtl number method isadded to correct the thermal eddy diffusion coefficient term of the transition model, which canguarantee the smooth transition of the turbulent Prandtl number from the boundary layer intothe mainstream area, and furthermore, the heat transfer coefficient is closer to the actualexperiment value.
The above-mentioned transition model was selected to carry out the numericalsimulation of the single-hole cooling. By comparing the analytical solution ofthree-dimensional impingement cooling with the theoretical and experimental solutions, thegrid independence was calculated to verify the accuracy of the numerical results. Thecorresponding law of the single-hole impingement cooling has a great help in understandingthe impact of the flow structure and the cooling mechanism. The identified numericalmethods were used to analyze the influence of the different geometrical and physicalparameters on the array jet impingement cooling heat transfer and flow. Compared with theexperimental correlations, the laws of the plat array jet heat transfer was determined toprovide the basis for the design of the next deflector impingement cooling holes.
With aid of the numerical simulation technology, the detail investigation of the enginerealistic operating conditions has a great significance in aero turbine engine design process.Lab environment can not totally realize the realistic operating conditions of the aero turbineengine. Then the guide blade impingement cooling structure was designed. Firstly the leaftemperature distribution with no-cooling is calculated, and it is found that the maximumtemperature of the leaves exists at the central surface of the leading edge. The air-cooledstructure was designed to complete the analysis of the three-dimensional aerodynamic andheat transfer coupled, to achieve a goal of enhancing the cooling effect and reducing the performance requirements of the material under a condition of the fixed cooling gas flow rate.When the number of holes and diameter along blade height satisfy the relationship（4n-1）D=h,the temperature can meet the cooling requirements. When each column of the jet hole X/d=5~7in the axial direction to the principle of simple structure, it can meet the coolingrequirements of the leaves chord region. In order to meet the cooling requirements of theleading edge, there shall be arranged in two cooling holes at least. In the small impact spacing,the impact spacing has little effect on the variation of the average temperature of the inner andouter surfaces, the maximum temperature will lower and the minimum temperature will riserwith smaller spacing. Just from the view of the cooling effect, the use of small spacing isappropriate. Compared with the cutout exhaust structure, the split seam structure of thetrailing edge can improve the cooling effect of the trailing edge and reduce the total pressureloss effectively.
Finally, the thermal stress was analyzed with temperature load and impact forceconditions on the blade by applying the thermoplastic coupling, and the influence of gridnumber was discussed. The results show that: the total maximum deformation exists in theblade central location with lower structural strength. The maximum thermal stress exists inthe blade root and the tip position, and it is due to the end wall which is given the constraintsof a fixed support. The deformation will be delivered to the two end walls producingcompression effect due to thermal stress generated in the guide blade.
叶轮机械中边界层转捩流动普遍存在，其直接影响叶片壁面附近传热过程，本文首先研究了如何准确预测边界层转捩流动问题，选择Transition k-kl-转捩模型对平板边界层有无压力梯度的跨越转捩实验进行数值计算，并对边界层转捩前后流动状态以及影响转捩流动的边界条件进行研究，计算结果表明：Transition k-kl-转捩模型对有无压力梯度的平板跨越转捩表现良好，压力梯度对边界层转捩流动有较大影响，顺压梯度有助于稳定边界层流动，延缓转捩的发生，而逆压梯度则容易使边界层流动失稳，促使转捩提前发生。
本文进一步对NASA-MarkII叶片传热实验进行气热耦合数值计算，研究了各湍流模型的计算特点以及它们对转捩流动的识别能力，分析了叶片表面边界层流动状态对涡轮叶片传热的影响，发现Transition k-kl-转捩模型相比其它全湍流模型，能够更准确预测附面层内的分离转捩状况。由于转捩前期采用层流动能来描述扰动的发展，避免了使用含来流湍流度的经验公式所带来的计算层流误差，引入“分裂机制”来描述层流与湍流脉动间的相互作用，并且在旁路转捩和自然转捩源项模块中加入了Tollmien-Schlichting波的影响，对强激波后的温度计算相比常用的间歇因子转捩模型与实验值更吻合。最后采用变湍流普朗特数方法对Transition k-kl-转捩模型的热涡流扩散系数项进行修正，保证湍流普朗特数从边界层到主流区光滑过渡，使得对流换热系数计算值与实验值更加接近。
选用上述转捩模型进行数值计算，得出平板单孔冲击冷却结果，并与三维单孔冲击冷却分析解以及实验关联式进行对比，同时对模型网格无关性进行计算，验证数值结果的准确性，得到单孔冲击冷却的相应规律，对于理解冲击流动的流场和冲击冷却的机理有很大帮助。再利用已确定的数值计算方法，分析阵列射流不同的几何参数和物理参数对冲击冷却传热和流动的影响，与实验关联式进行对比，得到平板阵列射流的传热规律，为接下来导向器叶片冷却结构设计提供依据。
实验研究不能完全仿照航空发动机的实际工作条件，利用数值计算的优势研究涡轮叶片的实际工作状态具有非常重要的工程意义，接下来本文对导向器叶片进行冲击冷却结构设计，首先计算得出未加冷却时导向器叶片的温度场分布，发现叶片最高温度存在于前缘外表面中径处，再根据具体温度分布进行导向器叶片气冷结构设计，完成气动和传热的三维耦合分析，实现在保持冷却气体流量不变的条件下提高冷却效果，降低材料的性能要求。当导流片孔数和孔径满足关系式(4n-1)D=叶高（mm）时，在导向器叶片叶高方向上能满足冷却要求。以结构简单为原则，选择导流片每列冲击孔轴向距离为X/d=5~7时，能够满足中弦区冷却要求。前缘区域要想满足要求，所设计的冷却结构前缘至少须布置两列冲击孔。冲击间距较小时，冲击间距变化对叶片的内、外表面平均温度影响不大，但随冲击间距的减小，内、外表面最高温度会降低，而最低温度则会升高。单纯从冷却效果看，采用小间距为宜。与开孔排气结构相比，尾缘劈缝排气结构可以有效改善尾缘的冷却效果，且减小了总压损失。
本文最后应用热弹耦合方法对温度和冲击载荷作用的导向器叶片进行了应力分析，探讨了网格数量对计算结果的影响。结果发现：总变形量最大值出现在结构强度较低的叶片中弦位置，等效应力最大值出现在叶根和叶尖位置，分析认为是由于轮毂、轮缘给定固定支撑的约束条件，叶体由于热应力产生的形变，均会传递到两个端壁面上产生压缩效应。
It is important to improve aero turbine engine thermal efficiency and thrust-to-weightratio by increasing turbine inlet temperature, which also brings a series of serious problemsfor the normal operation of the engine. Now the inlet temperature increases much faster thanthe development of the temperature resistance of the blade material. The efficient coolingtechnology and the thermal protective measures must be taken into consideration to ensure theturbine blades working safely and reliably at a high temperature, in addition to continuing toimprove high temperature materials resistance. The precise heat transfer analysis of turbineblades is essential to improve the cooling efficiency and extend the operating life of the blades.With the incessant maturity of computational fluid dynamics, the conjugate heat transfermethodology has gradually become a prevailing tools in the design process of aero turbineengine. With the help of numerical simulation of conjugate heat transfer methodology, theimpingement cooling structure which is focused on the small size high pressure turbine guideblade, has been designed, and then the three-dimensional aerodynamics analysis and the heattransfer coupled analysis of this structure have been completed, which can provide referencefor the cooling structure design of the actual engine guide blades.
The boundary layer transition flow is widespread in turbo machinery, and it directlyaffects the heat transfer process of the turbine blade near the wall. The problem of how toaccurately predict the boundary layer transition flow was firstly studied, then the Transitionk-kl-model was selected to simulate boundary layer transition on the flat plate, and then theflow state before and after the boundary layer transition was in-depth studied. The resultsshow that: Transition k-kl-model has a good performance on the plate bypass transitionwith or without pressure gradient, and the pressure gradient has influence to the boundarylayer transition flow. The favor pressure gradient can stabilize boundary layer flow, postponethe onset location of transition. Otherwise the adverse pressure gradient can induce theboundary layer separation, and pre-act transition occurs.
The conjugate heat transfer methodology has been employed to NASA-MarkII turbineguide vane in this dissertation. By discussing the calculation characteristics of the turbulencemodel and the ability to identifying transitional flows, the influence of the blade surfaceboundary layer flow on turbine blade heat transfer was analyzed and corrected. Compared to all the other turbulence models, the transition model can predict the separation transition stateof the boundary layer more accurately. As the transition model uses laminar kinetic energy todescribe the development of the disturbance, so this model can avoid the errors introduced bythe empirical formula, which is related with the inflow turbulence intensity. This modelintroduces the "split system" to describe the interaction between the laminar and the turbulentfluctuation, besides Tollmien-Schlichting wave is added into the bypass transition and thenatural transition source term. The result of temperature calculation behind the strong shock ismore consistent with the experimental value, compared to that of the common models withthe intermittent transition factor. Finally, the mutative turbulent Prandtl number method isadded to correct the thermal eddy diffusion coefficient term of the transition model, which canguarantee the smooth transition of the turbulent Prandtl number from the boundary layer intothe mainstream area, and furthermore, the heat transfer coefficient is closer to the actualexperiment value.
The above-mentioned transition model was selected to carry out the numericalsimulation of the single-hole cooling. By comparing the analytical solution ofthree-dimensional impingement cooling with the theoretical and experimental solutions, thegrid independence was calculated to verify the accuracy of the numerical results. Thecorresponding law of the single-hole impingement cooling has a great help in understandingthe impact of the flow structure and the cooling mechanism. The identified numericalmethods were used to analyze the influence of the different geometrical and physicalparameters on the array jet impingement cooling heat transfer and flow. Compared with theexperimental correlations, the laws of the plat array jet heat transfer was determined toprovide the basis for the design of the next deflector impingement cooling holes.
With aid of the numerical simulation technology, the detail investigation of the enginerealistic operating conditions has a great significance in aero turbine engine design process.Lab environment can not totally realize the realistic operating conditions of the aero turbineengine. Then the guide blade impingement cooling structure was designed. Firstly the leaftemperature distribution with no-cooling is calculated, and it is found that the maximumtemperature of the leaves exists at the central surface of the leading edge. The air-cooledstructure was designed to complete the analysis of the three-dimensional aerodynamic andheat transfer coupled, to achieve a goal of enhancing the cooling effect and reducing the performance requirements of the material under a condition of the fixed cooling gas flow rate.When the number of holes and diameter along blade height satisfy the relationship（4n-1）D=h,the temperature can meet the cooling requirements. When each column of the jet hole X/d=5~7in the axial direction to the principle of simple structure, it can meet the coolingrequirements of the leaves chord region. In order to meet the cooling requirements of theleading edge, there shall be arranged in two cooling holes at least. In the small impact spacing,the impact spacing has little effect on the variation of the average temperature of the inner andouter surfaces, the maximum temperature will lower and the minimum temperature will riserwith smaller spacing. Just from the view of the cooling effect, the use of small spacing isappropriate. Compared with the cutout exhaust structure, the split seam structure of thetrailing edge can improve the cooling effect of the trailing edge and reduce the total pressureloss effectively.
Finally, the thermal stress was analyzed with temperature load and impact forceconditions on the blade by applying the thermoplastic coupling, and the influence of gridnumber was discussed. The results show that: the total maximum deformation exists in theblade central location with lower structural strength. The maximum thermal stress exists inthe blade root and the tip position, and it is due to the end wall which is given the constraintsof a fixed support. The deformation will be delivered to the two end walls producingcompression effect due to thermal stress generated in the guide blade.
引文
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