Research On Conjugate Heat Transfer Simulation Of Aero Turbine Engine Air-Cooled Vane

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.