가스터빈 동익에서 분사비에 따른 막 냉각 효과에 관한 수치해석

Abstract

Recently desired for more efficient, turbine inlet total temperature has been increased, but the higher temperature gas causes the thermal stress seriously. The film cooling method is one of the ways to cool the turbine blade with hole of coolant at blade surface. The blowing ratio is the critical element to affect the blade temperature. This paper presents a numerical analysis to investigate the film cooling effectiveness for the variation of coolant blowing ratio in 1.5 stage axial turbine. The rotor blade leading edge of the rotor blade is film cooled with three rows of film cooling holes. Each hole has same area for ejection of coolant. In these processes, Navier-Stokes solver(CFX-solver) was used to calculate the flow inside the axial turbine and SST turbulent model. Consequently the increased blowing ratio has been resulted to raise the film cooling effectiveness at rotor blade pressure side, but the suction side shows opposite results. The holes at each side have no influence to the other side, but the holes at the leading edge have extensive effect on the both side of the rotor.