航空发动机气冷涡轮叶片的气热耦合数值的研究.pdf

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1、国内图书分类号：TK471学校代码：10213国际图书分类号：621.438密级：公开工学博士学位论文航空发动机气冷涡轮叶片的气热耦合数值模拟研究博 士 研 究 生：董平导 师：冯国泰 教授申请学位：工学博士学科：动力机械及工程所 在 单 位：能源科学与工程学院答 辩 日 期：2009 年 5 月授予学位单位：哈尔滨工业大学Classified Index:TK471U.D.C.:621.438Dissertation for the Doctoral Degree in EngineeringRESEARCH ON CONJUGATE HEATTRANSFER SIMULATION OF A

2、EROTURBINE ENGINE AIR-COOLED VANECandidate：Dong PingSupervisor：Prof.Feng GuotaiAcademic Degree Applied for：Doctor of EngineeringSpeciality：Power Machine and EngineeringAffiliation：Energy Science and EngineeringDate of Defence：May,2009Degree-Conferring-Institution：Harbin Institute of Technology摘 要-I-

3、摘 要摘 要现代航空燃气涡轮发动机为了获得更高的推重比和热效率，不断提高涡轮入口温度，目前涡轮进口温度已经远远超过叶片材料的熔点温度，必须采用复杂的冷却技术来保持涡轮叶片的正常工作，准确预测涡轮叶片的温度场是提高冷却效率、延长叶片工作寿命的关键问题，随着的计算流体力学的不断发展，气热耦合数值模拟技术已经成在航空发动机工程设计的重要工具，本文的主要工作就是对航空发动机气冷涡轮叶片进行气热耦合数值模拟，并对进一步提高涡轮叶片气热耦合数值模拟的准确性和可靠性进行深入分析。本文首先研究了如何准确预测边界层转捩流动的问题，航空发动机中普遍存在转捩流动，边界层转捩前后的流动状态、壁面切应力和换热系数等均不

4、相同，准确预测边界层转捩流动的起始位置对于涡轮叶片的气动和传热设计都具有非常重要的意义，能否准确模拟边界层转捩流动的关键在于如何选择合适的物理模型，本文选择常见的二方程湍流模型和转捩模型对平板边界层转捩实验进行数值模拟，通过详细分析各模型的计算特点和对转捩流动的识别能力，发现常见的高雷诺数 K-湍流模型和低雷诺数 K-湍流模型都是基于全湍流假设，不能准确模拟边界层转捩流动，目前情况下准确模拟边界层转捩流动必须选用合适的转捩模型，本文选用 g-Req转捩模型对平板边界层转捩流动进行数值模拟获得了比较令人满意的结果。为了研究涡轮典型工作环境下边界层转捩的特点，本文还研究了压力梯度和温度梯度对转捩流

5、动的影响，发现顺压梯度有助于稳定边界层流动，延缓转捩的发生，而逆压力梯度容易促使边界层流动失稳，促进转捩提前发生；垂直于壁面方向较大的温度梯度所引起的密度分层抑制了湍流的脉动运动，延缓边界层转捩的发生。航空发动机涡轮叶片的气动和传热过程非常复杂而且相互影响，本文通过对带有内部径向对流冷却的 MarkII 和 C3X 叶片的传热实验进行了气热耦合数值模拟，评估了不同湍流模型和转捩模型的气热耦合计算精度，发现边界层的流动状态对涡轮叶片的传热过程有很大的影响，当涡轮叶片存在转捩流动的时候必须选用合适的物理模型才能提高气热耦合数值计算的准确性和可靠性，本文选用 g-Req转捩模型比较准确的模拟了涡轮叶

6、片表面的传热过程。上述关于提高涡轮叶片气动和传热计算精度的研究对进一步模拟航空发动机涡轮叶片气冷结构的传热过程做了数值计算上的准备。哈尔滨工业大学工学博士学位论文-II-叶片外表面冷却射流与高温主流燃气之间的相互掺混是一个复杂三维气动与传热过程，本文通过对带有气膜冷却 C3X 叶片传热实验进行气热耦合数值模拟，发现气膜冷却射流在叶片表面的流动和传热过程和主流边界层的流动状态也密切相关，当主流边界层为层流状态时，气膜冷却射流能保持自身的独特流动状态，冷却射流覆盖区域的冷却效果较好，但当主流边界层转捩为湍流以后，由于冷却射流与主流边界层之间掺混流动增强促使气膜冷却射流与主流边界层迅速融合，冷却射流

7、对叶片表面的冷却效果大大降低；当气膜冷却射流的流动状态保持较好时，气膜冷却射流的动量相对当地的主流边界层要大得多，气膜冷却射流在叶片表面形成若干平行的“气流柱”，主流边界层流体只能从相邻“气流柱”之间压缩流过，并且气膜射流的对转涡对不断裹挟周围的高温流体到边界层内，使这些区域的换热增强；另外叶片表面多排气膜冷却相互叠加作用也很强烈，前缘气膜射流的较大径向流动分速诱使下游气膜冷却射流也明显向叶顶偏斜，下游冷却射流形成密集连续的气膜将前缘冷却射流和主流边界层抬离壁面，在下游较远处才向壁面发生再附。实验研究不能完全模拟航空发动机的实际工作条件，利用数值模拟的优势深入研究涡轮叶片的实际工作状态具有很重

8、要的工程意义，本文对航空发动机实际工作条件下涡轮叶片前缘复合冷却结构的传热过程进行了气热耦合数值模拟，复合冷却结构包括气膜冷却、对流冷却、冲击冷却、热障涂层和高温镍基合金，涡轮叶片前缘复合冷却结构的隔热降温效果比较明显，叶片的实际工作温度被有效降低到比较安全的范围，冲击冷却配合气膜冷却使用能有效降低叶片前缘的热负荷，在叶片表面添加热障涂层可以在相同的工作条件下大大提高涡轮叶片的抗氧化和抗热腐蚀的能力，并适当降低金属叶片的工作温度和温度梯度。本文最后研究了燃烧室出口温度不均匀分布对涡轮叶片前缘传热的影响，由于燃烧室内掺混不充分以及壁面冷却形成的涡轮进口温度沿径向不均匀分布是航空发动机运行过程中常

9、见的现象，这种温度的径向不均匀分布使涡轮叶片表面的热负荷变得不均匀，同时使叶片前缘和下游气膜冷却在局部区域的冷却效果降低，在冷却设计中应该考虑进口温度径向分布不均匀对涡轮叶片传热的影响；“热斑”是由涡轮进口温度不均匀性的一种畸变状态，热斑对涡轮叶片表面形成非常集中、强烈的热冲击，叶片局部形成很大的温度梯度，极大影响了涡轮叶片的工作寿命，热斑与涡轮叶片的相对位置变化存在“时钟效应”，当热斑正对叶片前缘时，对叶片的热冲击最大；热障涂层能在一定程度上减小热斑对涡轮叶片的热冲击。关键词：涡轮；数值模拟；气热耦合；转捩；复合冷却；温度不均匀分布Abstract-III-AbstractIn order

10、to improve aero turbine engine thermal efficiency and thrust-to-weightratio,combustor exit temperature have clearly exceeded the permissible materialtemperature of turbine bladings,and complex cooling techniques are commonlyused to maintain turbine bladings working in safe conditions.Precise heat tr

11、ansferanalysis of turbine bladings is essential to increase cooling configuration efficiencyand extend bladings operating life.With the incessant maturity of numericalsimulation technology,conjugate heat transfer methodology has gradually became aprevailing tools in the desigh process of aero turbin

12、e engine.Base this point,themain task of this dissertation is to study the heat transfer of turbine bladings byconjugate heat transfer methodology,and explore how to increase the accuracy and reliabilityofconjugate heat transfer calculation.First of all,the method to precisely simulate the flow of b

13、oundary layertransition was investigated.Boundary layer transition is the common flowphenomenon in aero turbine engine.Before and after transition flow occurs,theboundary velocity profile,wall shear stress and heat transfer coefficient are totallydissimilar in boundary layer.Precise numerical simula

14、tion of boundary layertransition is essential to the aerodynamics and heat transfer design of turbinebladings.Different turbulence models and turbulence models were employed in thenumerical simulation of flat plate boundary layer transition experiments.Bycomparing calculated results with different t

15、urbulence models to the measured data,it is clear that calculation with g-Req transition model can better simulate the flowin boundary layers because dual-equation turbulence models regard the wholeboundary layer field as a full turbulence flow.Pressure gradient and temperaturegradient are common ph

16、enomena in aero turbine engine and influence the onsetlocation of boundary layer transition.Favor pressure gradient stabilize boundarylayer flow,and postpone the onset location of transition.Adverse pressure gradientinduce boundary layer separation and pre-act transition occurs.Higher temperaturegra

17、dient normal to plate induces higher density gradient which could reduceturbulence characteristics in boundary layer,and delay transition process.哈尔滨工业大学工学博士学位论文-IV-Secondly,conjugate heat transfer methodology was employed to MarkII andC3X turbine guide vane which has internal radial convective cool

18、ing channels inthis dissertation.By comparing conjugate heat transfer simulation with differentturbulence models,the calculation with g-Req transition model showed goodagreement with measured data.Boundary layer transition had significant influenceon heat transfer of turbine bladings and application

19、 of transition models was crucialto the accuracy and reliability of conjugate heat transfer methodology at the presentday.The study mentions above which focused on improving the calculated accuracyof aerodynamics and heat transfer in turbine bladings laid the foundation for furtherinvestigation to c

20、ooling techniques of aero turbine engine.Conjugate heat transfer methodology was employed to C3X turbine vane withfilm cooling in this dissertation.The results indicated that the flow and heat transferof film cooling ejections were influenced by flow of mainstream boundary layer onvane surface.The u

21、nique 3-D turbulent characteristics of film cooling ejects was,but quickly disappeared owing to its intensified mixing with hot gas aftermainstream boundary layer transition,and cooling performance to vane surface wassimultaneously weakened.When the flow characteristics of film cooling ejects wascle

22、arly maintained before transition,film cooling ejects separate mainstreamboundary layer into parallel street on vane surface because the momentum of filmcooling ejects was greater than local mainstream boundary layer,the heat transferwas reinforce in those streets because of compressed hot gas flow

23、in those streetsand the entrainment phenomenon of film cooling ejection.The interaction ofoverlap between leading edge film cooling and downstream film was so strong thathad greater influenced on heat transfer process on vane surface.The radial velocityof leading edge film cooling ejection induced d

24、ownstream film cooling ejections toflow in same direction,Down stream film cooling ejections raised upper streamcooling ejections away from vane surface,and reattachment occurred at a distance.Finally,with aid of numerical simulation technology,detail investigation ofengine realistic operating condi

25、tions had great significance in aero turbine enginedesign process.Lab environment could not totally simulate aero turbine enginerealistic operating conditions.Conjugate heat transfer methodology was employedto study the heat transfer of compound cooling structure employed in the leadingedge of C3X v

26、ane under engine-realistic operating conditions in this dissertation.Abstract-V-The compound cooling techniques include film cooling,convection cooling,impingent cooling,thermal barrier coating and nickel based heat resistingsuperalloy.Compound cooling techniques had excellent capability of coolingt

27、urbine bladings to a permissible temperature under harsh working conditions.Thecombined usage of internal impingent cooling and outer film cooling couldeffectively reduce the heat load of turbine bladings.Thermal barrier coating couldnot only effectively increase the capability of oxidation resistan

28、t and corrosionresistant,but also moderately lower the working temperature of turbine bladings.The non-uniform turbine inlet temperature distribution in radial direction caused byinsufficient mixing and wall film cooling in combustor could usually induceinhomogeneous heat load of vane surface,and we

29、aken the performance ofdownstream film cooling,so researchers should pay more attention to the influenceof non-uniform temperature in cooling technique design process.Hot streakphenomenon was a turbine inlet temperature distortion which had both radial andcircumferential severe temperature gradients

30、.The thermal shock of hot streakusually caused great temperature gradient in turbine bladings,and was very harmfulto operating life.The location of hot streak has a clocking effect on the surfacetemperature distribution of turbine bladingsTurbine vane received the biggestthermal shock when the core

31、of hot streak was positively aligned with turbine vaneleading edge.Thermal barrier coating could little reduce the thermal shock of hotstreak.Keywords:turbine,numerical simulation,conjugate heat transfer methodology,transition,compound cooling techniques,non-uniform inlet temperature哈尔滨工业大学工学博士学位论文-

32、VI目录摘要.IAbstract.III第 1 章 绪论.11.1 课题来源.11.2 课题研究的背景及意义.11.3 航空燃气涡轮冷却技术概述.31.3.1 涡轮冷却技术发展历程简介.31.3.2 涡轮叶片的典型冷却方式.31.3.3 航空发动机涡轮叶片冷却术的设计要求和主要问题.71.4 航空发动机涡轮叶片的换热研究综述.81.4.1 航空发动机燃烧室出口条件对涡轮叶片传热的影响.81.4.2 涡轮叶片的传热过程的实验研究.111.5 航空发动机涡轮叶片的气热耦合数值模拟.201.6 转捩模型研究发展.251.6.1 转捩的基本概念.251.6.2 叶轮机械内的转捩流动.261.6.3 转

33、捩流动的常见数值模拟方法.281.6.4 基于局部变量的转捩模型的发展.311.7 本文的主要研究内容.331.8 本章小结.35第 2 章 数值模拟方法.362.1 引言.362.2 数值模拟程序.362.2.1 Fluent.362.2.2 CFX.372.3 数值计算方法.382.3.1 求解器.382.3.2 湍流模型.43目录-VII-2.3.3 转捩模型.452.3.4 分块结构化网格.472.4 数值模拟程序气热耦合计算精度校核.482.5 本章小结.49第 3 章 平板转捩的数值模拟.503.1 引言.503.2 T3 系列平板转捩实验的数值模拟.513.2.1 T3 系列平板

34、转捩实验工况和计算条件.513.2.2 零压力梯度平板转捩的数值模拟.523.2.3 带有压力梯度平板转捩的数值模拟.573.3 温度梯度对转捩的影响.633.3.1 实验工况和计算条件.633.3.2 计算结果与讨论.633.4 本章小结.66第 4 章 无气膜冷却涡轮叶片的气热耦合数值模拟.694.1 引言.694.2 数值方法和模拟工况.694.3 MarkII 叶型计算结果讨论.724.3.1 计算结果与实验值的对比.724.3.2 使用不同湍流模型计算的叶片表面边界层流动状态分析.764.4 C3X 叶型计算结果讨论.794.5 本章小结.82第 5 章 带有气膜冷却涡轮叶片的气热耦

35、合数值模拟.845.1 引言.845.2 带有前缘气膜冷却的涡轮叶片的气热耦合数值模拟.845.2.1 计算模型和边界条件.845.2.2 计算结果和实验结果的对比.875.2.3 前缘气膜冷却射流对叶片表面的换热影响.885.2.4 前缘气膜冷却射流在叶片表面的流动状态分析.945.3 带有多排气膜冷却的涡轮叶片的气热耦合数值模拟.995.3.1 计算模型和边界条件.995.3.2 计算结果和实验结果的对比.1015.3.3 多排气膜冷却的对叶片表面传热的影响.102哈尔滨工业大学工学博士学位论文-VIII5.3.4 多排气膜冷却的叠加作用.1035.3.5 多排气膜冷却在叶片表面的流动状态

36、分析.1065.4 本章小结.110第 6 章 涡轮叶片前缘复合冷却的气热耦合数值模拟.1126.1 引言.1126.2 计算模型和实验工况.1126.3 涡轮叶片内部冲击冷却的数值模拟.1156.4 涡轮进口温度的不均匀分布对叶片前缘传热的影响.1206.4.1 涡轮进口温度的径向正弦分布对叶片前缘传热的影响.1206.4.2 涡轮进口温度热斑对叶片前缘传热的影响.1236.5 叶片表面热障涂层的气热耦合数值模拟.1306.6 实际工作条件下的涡轮叶片前缘的气热耦合数值研究.1346.7 本章小结.138结论.141参考文献.145攻读学位期间发表的学术论文.157哈尔滨工业大学博士学位论文

37、原创性声明.159哈尔滨工业大学博士学位论文使用授权书.159致谢.160个人简历.161Contents-IX-ContentsAbstract(in Chinese).IAbstract(in English).IIIChapter 1 Intruduction.11.1 Programe backound.11.2 Motivation and purport of the dissertation.11.3 Research status of aero turbine engine cooling techniques.31.3.1 Development of turbine bl

38、adings cooling techniques.31.3.2 Typical cooling techiques of turbine bladings.31.3.3 Problem and objectives in design of turbine bladings cooling techniques.71.4 Research status of turbine bladeings heat transfer.81.4.1 Influence of combustor outlet conditions to turbine bladings heat transfer 81.4

39、.2 Experimental Research of turbine bladings heat transfer.111.5 Conjugate heat transfer simulation of turbine bladings.201.6 Research status of transition models.251.6.1 Principles of transition.251.6.2 Typical transition in turbine machinery.261.6.3 Numerical simulation methodology of transition.2

40、81.6.4 Transition models based on local variavle.311.7 Main content of this dissertation.331.8 Summery of the chapter.35Chapter 2 Numerical simulation methodology.362.1 Introduction.362.2 Numerical soler.362.2.1 Fluent.362.2.2 CFX.372.3 Numerical method.382.3.1 Solver.382.3.2 Turbulence models.43哈尔滨

41、工业大学工学博士学位论文-X2.3.3 Transition models.452.3.4 Multiple block grids.472.4 Validation of the solver conjugate heat transfer calculation precision.482.5 Summery of the chapter.49Chapter 3 Numerical simulation of flat plate boundary layer transition.503.1 Introduction.503.2 Numerical simulation of T3 se

42、ries flat plate boundary layer transitionexperiments.513.2.1 Geometric parameters and boundary conditions of T3 series experiments513.2.2 Numerical simulation of flat plate boundary layer transition at zeropressure gradient.523.2.3 Numerical simulation of flat plate boundary layer transition underpr

43、essure gradient.573.3 Effects of temperature gradient to flat plate boundary layer transition.633.3.1 Numerical model and boundary condtitions.633.3.2 Analysis of numerical results.633.4 Summery of the chapter.66Chapter 4 Conjugate heat transfer simulation of turbine vane without filmcooling.694.1 I

44、ntroduction.694.2 Numerical model and boundary conditions.694.3 Conjugate heat transfer simulation of MarkII vane.724.3.1 Contrast between numerical results and measured data.724.3.2 Analysis of boundary layer flow calculated with different turbulencemodels.764.4 Conjugate heat transfer simulation o

45、f C3X vane.794.5 Summery of the chapter.82Chapter 5 Conjugate heat transfer simulation of turbine vane with filmcooling.845.1 Introduction.845.2 Conjugate heat transfer simulation of turbine vane with leading edge filmcooling.845.2.1 Numerical models and boundary conditions.84Contents-XI-5.2.2 Contr

46、ast between calculated result and measured data.875.2.3 Influence of leading edge film cooling ejects to heat transfer on vanesurface.885.2.4 Flow characteristic of leading edge film cooling ejects.945.3 Conjugate heat transfer simulation of turbine vane with multiple film cooling.995.3.1 Numerical

47、model and boundary conditions.995.3.2 Contrast between calculated results and measured data.1015.3.3 Influence of multiple film cooling to heat transfer on vane surface.1025.3.4 Interaction of multiple rows film cooling ejects.1035.3.5 Flow characteristics of multiple film cooling ejects.1065.4 Summ

48、ery of the chapter.110Chapter 6 Conjugate heat transfer simulation of leading edge compoundcooling techniques.1126.1 Introduction.1126.2 Numerical models and boundary conditions.1126.3 Conjugate heat transfer simulation of internal impingent cooling.1156.4 Effects of non-uniform turbine inlet temper

49、ature to heat transfer on vanesurface.1206.4.1 Effects of non-uniform turbine inlet temperature in raidal direction toheat transfer on vane leading edge.1206.4.2 Effects of hot streak to heat transfer on vane leading edge.1236.5 Conjugate heat transfer simulation of turbine vane with thermal barrier

50、coating.1306.6 Conjugate heat transfer simulation of turbine vane under engine realisticworking conditions.1346.6 Summery of the chapter.138Conclusion.141Reference.145Papers published in the period of Ph.D.education.157Statement of Copyright.159Letter of Authorization.159Acknowledgement.160Resume.16