항공기용 고압터빈의 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 of k-ω 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. The flow characteristics were analyzed by comparing with uncooled nozzle vane. The Mach number around the vane increased due to the increase of coolant mass flow flowed in the main flow passage. The maximum cooling effectiveness (91%) at the vane surface is located in the middle of pressure side which is effected by the film cooling and the rib turbulrator. The region of the minimum cooling effectiveness (44.8%) was positioned at the leading edge. And the results show that the TBC layer increases the average cooling effectiveness up to 18%.