【作者】 董平；

【导师】 冯国泰；

【作者基本信息】 哈尔滨工业大学 ， 动力机械及工程， 2009， 博士

【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-ǜ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。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。更多还原

【Abstract】 In order 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 transferanalysis 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 turbine 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 reliabilityof conjugate heat transfer calculation.First of all, the method to precisely simulate the flow of boundary 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 simulation 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 turbulence 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 phenomena 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 temperaturegradient normal to plate induces higher density gradient which could reduceturbulence characteristics in boundary layer, and delay transition process. Secondly, conjugate heat transfer methodology was employed to MarkII andC3X turbine guide vane which has internal radial convective cooling 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 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 cooling 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 unique 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 wasclearly 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 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 downstream 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 conditions 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 vane under engine-realistic operating conditions in this dissertation. 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 coolingturbine 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 resistant 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 weaken 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. 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 bladings.Turbine vane received the biggestthermal shock when the core of hot streak was positively aligned with turbine vaneleading edge. Thermal barrier coating could little reduce the thermal shock of hotstreak.更多还原