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Aero-thermal Investigation of a Film-cooled Blade in a High-pressure Turbine

Aero-thermal Investigation of a Film-cooled Blade in a High-pressure Turbine
Jinuk Kim
Issue Date
A conjugate heat transfer (CHT) analysis was performed to improve the aerodynamic and cooling performance of the high-pressure turbine of a turbofan engine for small and medium aircrafts. For accurate computational analysis of a high-pressure cooled turbine, a computational fluid dynamics (CFD) method was established by using a NASA C3X vane. Reliable CFD results for high-pressure turbine (HPT) cooled nozzles can be obtained by using a k-ω based shear stress transport turbulence model, viscous work term, and γ-transition model. The established CFD method was used to perform CHT analyses on the HPT of a turbofan engine designed by the Korea Aerospace Research Institute. The aerodynamic and heat transfer characteristics of the vanes and blades including those of the internal cooling components and external film-cooling holes, were verified by CHT analysis. In addition, it was confirmed that the blade material (CMSX-4) does not cause thermal deformation when the thermal barrier coating is applied to vane and blade surfaces. The optimal blade tip cavity shapes that can improve the aerodynamic and heat transfer performance were derived as well. For the optimization procedure, an experimental design using central composite design and latin hypercube design, an approximate model using kriging regression, and an optimization method using multi-objective generic analysis were employed. In the airfoil type cavity, the cavity depth, leading edge radius, and trailing edge radius were selected as design parameters. It was confirmed that the larger the cavity depth, the larger the front and rear radius of the cavity and the better the heat transfer characteristics
however, this would also imply a larger aerodynamic loss. To improve the aerodynamic performance of the HPT, the shape of the endwall (nozzle hub, nozzle shroud, and rotor hub) was optimized. The optimization procedure used in this study is the same as the blade tip cavity shape optimization procedure. The nozzle hub, nozzle shroud, and rotor hub endwall were independently optimized, and when the three optimized endwall shapes were combined, the turbine efficiency could be increased by 0.7%p compared to the basic shape. It is expected that the enhancement of the aero-thermal characteristics of a high-pressure cooled turbine will improve the problems of the existing design model, and can be applied to the design of gas turbine engines in the future.
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