Modernization Of Cooling System Of The High-pressure Turbine Blade For Bypass Turbofan Engine

In a gas turbine engine temperature of hot gas before inlet of turbine is one of the determining parameters.In modern turbine research community,cooling system design has been raised to the key issues with the rise of the inlet temperature.The purpose of this study is to conduct a numerical study of the film cooling of the initial model of the High-Pressure Turbine rotor blade and the model with optimized geometry of the exhaust cooling channels.Modern technical solutions for reducing the temperature of the cooling film on the turbine blade surface were considered.Since it is of great practical interest to search for new technical solutions that will provide the required cooling efficiency in combination with an acceptable coolant flow rate and a simpler manufacturing technique.Ansys CFX software package was used for numerical simulation.Optimization in the original model was subjected to a double row of cylindrical holes placed in high pressure side of the blade.Cylindrical cooling channels have been replaced by fan-shaped cooling channels,and their number has been reduced in order to preserve the mass-flow rate characteristics of the coolant.The result of this work is to obtain reduced characteristics for the temperature of the film on the blade profile directly behind optimized row of holes compared with the initial model.From what the conclusion about reasonability of application of fan-shaped channels in local places of overheating where they show the best indicators than cylindrical channels were made.Fan-shaped holes made the main contribution,compared to conventional cylindrical ones,is to prevent or disrupt renal vortices tearing the cooling air flow away from the blade surface.