Analysis of Compound Angled Gas Turbine Blade Leading Edge Model on Adiabatic Film Cooling Effectiveness By Numerical Investigation using CFD

This study aims at investigating the film cooling effectiveness using numerical analysis for the scaled up gas turbine blade leading edge model. A typical gas turbine blade leading edge model is selected for the study with five rows of holes, one at stagnation line, two rows of holes at 30° on either side of stagnation line and two rows of holes at 60° on either side of stagnation line respectively and each row consisting of five holes. This scaled up geometry has the film cooling hole diameter of 5.6mm, considered based on the literature survey and calculations. The film cooling hole rows are arranged in staggered manner to cover the more flow area on the blade surface. Each row has the five holes at a pitch of 22.4 mm with the varied hole angles of 25°, 35°, 40°, 50°, and 60° oriented with the stream line direction. Film cooling effectiveness analysis are found using CFD (Fluent) simulation by varying the blowing ratios (B.R) in the range of 1.0, 1.50, 2.0 and 2.50 at the density ratio (D.R) of 1.30 with nominal mainstream flow Reynolds number of 1,00,000 based on the leading edge diameter. The generated CFD results are compared with the previously published results in journal for the validation. The CFD results indicate similar trends of the cooling effectiveness results as that of experimental results. The CFD results has shown the increase in cooling effectiveness with the increase in blowing ratio up to 2.0 and found there is no increase in effectiveness above the blowing ratio 2.0 for this model. Hence, the optimized blowing ratio for this LECA model can be considered as 2.0 with the higher cooling effectiveness.