A Project on Optimizing Cooling Passages in Turbine Blades

The gas turbine in its most common from is a rotary heat engine operating by means of series of processes consisting of air taken from the atmosphere increase of gas temperature by constant pressure combustion of the fuel the whole process being continuous. The turbine inlet temperature in modern gas turbines is far above the permissible metal temperature. Turbines need to run at very high temperatures to reduce fuel burn, but they require internal cooling to maintain structural integrity and meet service-life requirements. A blade can be defined as the medium of transfer of energy from the gases to the turbine rotor. To increase the efficiency of gas turbine inlet Temperature should be high as possible. Turbine blades have internal passages that provide cooling during operation in a high temperature engine. The design of the cooling passages is critical to achieve near uniform temperature of the blade during operation. The temperature of the blade is dependent on the thermal properties of the blade material as well as the fluid dynamics of the air circulating in the cooling passages. Computational optimization methods have successfully been applied to design lighter and more efficient structures for many aerospace structures. An extension of these techniques is now applied to guiding the thermal design of a turbine blade by designing the optimal cooling passage layout. Optimization methods will be applied to determine the optimum pattern of the cooling passages and then to optimize the size of the individual cooling passages. The goal is to produce a more thermally efficient turbine blade design that will produce blades with longer lives and better performance. In this project the model blade of the gas turbine is created in PTC Creo Parametric 3.0. The cooling passages were modeled into the blade shape and the entire model was meshed in Hyper Mesh 2019.