항공기용 가스터빈의 고압 냉각터빈 노즐에 대한 복합열전달 해석

Abstract

Conjugate heat transfer analysis was performed to investigate the flow and cooling performance of the high pressure turbine nozzle of gas turbine engine. The CHT code was verified by comparison between CFD results and experimental results of C3X vane. The combination with k-w based SST turbulence model and transition model was used to solve the flow and thermal field of the fluid zone and the material property of CMSX-4 was applied to the solid zone. The turbine nozzle has two internal cooling channels and each channel has a complex cooling configurations such as the film cooling, jet impingement, pedestal and rib turbulator. The parabolic temperature profile was given to the inlet condition of the nozzle to simulate the combustor exit condition. High pressure rise at trailing edge of cooled vane was occurred. This phenomenon was not happened in case of an uncooled vane. The overall cooling effectiveness were evaluated for the nozzle. The overall cooling effectiveness of vane surface was satisfied the design goal which is the creep temperature, 1,223K. However, over-cooled region was founded at pressure side. The effect of thermal barrier coating on the overall cooling effectiveness was examined. The TBC layer of 300 ㎛ increase 18% of average cooling effectiveness.