Numerical analysis of film cooling effectiveness using compound cooling holes at the end of gas turbine engine combustor

Abstract

By using the well-known Bryton cycle, great turbine industries strive to extend the turbine inlet temperature and increase engine performance. However the turbine inlet temperature increment creates harsh environment for the downstream components of the combustor and so it is needed to design a cooling technique. The blowing ratio increase, caused to cooling effectiveness enhancement, however, for the traditional cooling system, the coolant not attached better on the surface at higher blowing ratios and therefore, it is required to restructure the cooling holes. Compound cooling holes is the useful way to this achievement. But, Most of the previous studies paid attention on the using compound cooling holes on the turbine blades and there is a lack of research on the application of these holes at the end wall of combustor. This study was accomplished in order to investigate the effects of cylindrical and row compound cooling holes with alignment angle of 30 degree, 60 degree and 90degree. The combustor used in the study is Pratt and Whitney gas turbine engine. This model was simulated and analyzed with a commercial finite volume package ANSYS FLUENT 14.0 to gain fundamental data.The entire findings of the study showed that with using the row compound cooling holes near the end wall surface, film cooling effectiveness is doubled compared to the cooling performance of baseline case. To conclude, as different arrangements of cooling holes affect the film cooling performance, it is strongly recommended to use a combination of compound cooling holes with different alignment angles for cooling panels.