Numerical Study On Cooling Performance Of Air-cooled Turbine Blade With Thermal Barrier Coating

The design of turbine airfoil is extremely important for the increasing of aeroengine thermal efficiency and thrust ratio.The developed cooling technology and thermal protection technologies are used to reduce the blade temperature.In the work,the cooling performance coated vane with film cooling are studied with conjugate heat transfer method(CHT).The results show that:(1)The geometric model of NASA C3X vane with thermal barrier coating(TBC)was established,three mesh models are used to assure grid independence.Compared with different turbulence models;transition k-kl-ω turbulence model is selected for simulation.The effect of cooling flow rate on the cooling performance of a coated turbine vane with internal cooling is studied with the conjugate heat transfer method.The analysis shows that,cooling efficiency increases with the increase of cooling flow.However,the efficiency increment declines with the increasing of the cooling flow.Thermal barrier coating on the suction side is more effective than that of the pressure side.In the base case,the decrease amplitude in temperature is about 72.6K which leads to an increase of cooling efficiency about 6.5%.In the trailing edge,thermal barrier coating prevents the heat transfer from the metal surface to the mainstream on the regions with insufficient internal cooling which increase the metal surface temperature,and thus decline the overall cooling efficiency those regions vane.Thus,sufficient internal cooling is the promise of improving the overall cooling performance for a coated turbine vane.(2)The effect of film cooling on the cooling performance of a coated turbine vane with internal cooling is studied by using the conjugate heat transfer method.The analysis shows that,film cooling on the pressure side is more effective than suction side;For uncoated vane with leading edge(LE)film,temperature decline is about15.6K at x/L =-0.36 in comparison with the no film case.On the suction side,the value is decrease to about 6.0K near the same radial channel.The interconnected influence between various coolant ejections leads to part of coolant ejections lifting off the surface of vane and thus offset the cooling performances on those regions;Thermal barrier coating show positive roles on the cooling performance increment in the area near the cooling film hole,but in the area far away from the cooling film hole and the area covered by the cooling film has negative effect;Thus,improving the convective heat transfer of coolant with the wall of coolant plenums and film cooling holes is the guarantee of improving the cooling performance of the coated vane with film and internal cooling.(3)The effect of inlet hot streak on the cooling performance of thermal barrier coating gas-cooled turbine blades is studied with the conjugate heat transfer method.The analysis shows that,thermal barrier coating can improve the cooling performance of vane surface,especially on the suction surface.The inlet hot streak will migrate on the suction surface,which makes the cooling performance decrease on the pressure surface near the trailing edge and on the suction surface near the leading edge,and the maximum temperature increases about 69 K.The migration of inlet hot streak to the top of vane is affected by the direction of internal cooling flow,which cannot be changed by thermal barrier coating.In the case of hot streak condition with thermal barrier coating added,the cooling performance of vane surface is improved more than that under uniform temperature.Thermal barrier coatings can reduce the effect of hot streak on cooling performance,even in areas with insufficient internal cooling,and the maximum growth rate of the overall cooling effectiveness can reach 50%.