Investigation On Film Cooling Characteristics For Backward Injections With Swirling Coolant

Film cooling has been successfully used in cooling gas turbine components that are exposed to very high temperature environments as an advanced technology.Several experts research that the performance of turbine end-wall film cooling is mostly affected by the secondary flow in the cascade channel,and there is a cooling dead zone in the pressure corner region.In this paper,backward injections with swirling coolant at the exit of the cooling hole are proposed to improve the film cooling performance of the end-wall pressure side corner region.In this paper,the cooling characteristics of backward injections are numerically studied compared to the forward injection on a flat plate,and the flow field characteristics and the vortex structure are analyzed.The cooling performance and aerodynamic losses of different blow ratios for backward injections with different compound angle are discussed in detail.Compared with the forward injection,the results show that the film cooling performance of backward injections is improved,but the aerodynamic loss of backward injections is larger.In terms with the disadvantage of aerodynamic losses for the forward injection,A swirling coolant at the exit of the cooling hole are leaded for the backward injections.In this paper,adiabatic cooling performance and aerodynamic losses of film cooling for forward and backward injections and backward injections with swirling coolant at the exit of the cooling hole is explored for compound angle cylindrical hole.The results show that,compared with the backward injections,the film cooling performance of the backward injections with swirling coolant at the exit of the cooling hole is significantly improved,and the aerodynamic loss is significantly declined.The average cooling effectiveness of the case of backward injections with a swirling flow raises up approximately 17% in all cooling areas,and the total pressure loss coefficient declines about 30% at a blowing ratio M=1.0.Compared with the forward injections,the film cooling effectiveness raises up obviously.With the same total pressure loss coefficient(? = 2% and ? = 3%),the film cooling effectiveness was increased by 20% and 40% respectively.In order to improve the film cooling performance of the pressure side corner region of the turbine end wall,the end-wall film cooling characteristics of the backward injections with swirling coolant at the exit of the cooling hole are further studied for 40° compound angle cylindrical hole.The results show that the secondary flow in the cascade channel are suppressed.the degree of migration of the jet cold air to the suction side decreases,the cooling gas moves to the pressure side,and the cooling performance of the film at the root of the pressure side is obviously improved.In the upstream cooling zone,the film cooling efficiency at the root of the pressure side blade is increased by about 100%;in the downstream cooling zone,the film cooling efficiency at the root of the pressure side blade is increased by about 35%.In terms with aerodynamic losses,backward injections with a swirling flow are improved obviously.Compared with the forward injection,the total pressure loss coefficient is declined by up to about 6.2% at a blowing ratio M=0.8.