Influence of film hole arrangement on cooling and aerodynamic performance of blade tip with squealer structure

Due to the high thermal load of the gas turbine, effective cooling of the blade is important, in which film cooling is a common used method. However, there is no universal consensus on the cooling hole arrangement, especially on the blade tip. In this paper, three typical cooling hole arrangements on blade tip with squealer structure have been investigated, namely Near Pressure Side (NPS), Camber Line (CL), and Near Suction Side (NSS). Numerical and experimental methods are applied to obtain the influence of cooling hole arrangement on cooling and aerodynamic performance at blade tip region. The distribution of film cooling effectiveness on blade tip surface and total pressure loss of blade cascade are analyzed. Results show that, with high blowing ratio (M), the coolant decreases the tip leakage flow with CL and NSS. For cooling effect, the CL and NPS arrangements result in better film cooling effectiveness than NSS arrangement. A topology model describing the cooling jet migration in the tip region is proposed, which shows that coolant coverage area increases where the cooling hole is arranged at the pressure side cavity vortex (PSCV) separation line. The proposed topology model provides direct guide on the optimized arrangement of cooling holes. Results show that much improved cooling effectiveness can be obtained, which is 68.6% higher than that of the CL arrangement. Albeit very simple, the proposed model can be used as a guideline for film hole arrangement on blade tip.