환형 섹터 터빈 노즐 캐스케이드에서 팬 형상 막냉각 홀의 복합분사각도 적용에 따른 막냉각 효율 특성

Abstract

The present study investigated the effect of compound angle injection of fan-shaped film cooling holes on film cooling performance on the nozzle vane surface in an annular sector cascade. Two different configurations were examined to investigate effect of compound angle injection. The baseline has streamwise injection holes on the pressure and suction side surfaces and normal injection holes in leading edge region. The other configuration has compound angle injection holes on the leading edge and the pressure side surface. An annular sector turbine cascade test facility was used and the measurements using an infrared thermography method were conducted at the exit Reynolds number of 2.2x10<SUP>6</SUP> and exit Mach number of 0.8. Total coolant mass flow rate ranges between 5 and 10% of mainstream. The results showed that the baseline configuration shows skewed trajectory of coolant downstream of holes on the pressure side surface, which is not observed in ‘flat-plate’ study and strongly related to mainstream flow direction near the nozzle surface. On the other hands, the coolant follows the mainstream direction and spread more uniformly on the pressure side surface with compound angle injection configuration. However as coolant mass flow rate increases, compound angle injection induces stronger interaction of mainstream and coolant and film cooling effectiveness drops quickly in the downstream region. Also, film cooling effectiveness in the suction side surface shows lower values for compound angle injection configuration since compound angle injection on the showerhead region causes non-uniform distributions of film cooling effectiveness due to stronger secondary flow.