FILM COOLING OF TURBINE BLADE SURFACE WITH EXTENDED EXIT HOLES

Gas turbines are extensively used for air craft propulsion, land based power generation and industrial applications. The current rotor inlet temperature in advanced gas turbine is for above the melting point of blade material. A sophisticated cooling scheme must be developed for continuous safe operation of gas turbines with high performance. The techniques that involve to cool the blades and vanes by using cooling methods is to have radial holes to pass high velocity cooling air along the blade span. The turbine blade is designed with no holes, 5 holes, 9 holes, 13 holes. CFD analysis is done to determine the pressure distribution, velocity, temperature distribution and heat transfer rate by applying the inlet velocities. Thermal analysis is done to determine the heat transfer rates of the blade. The present used material for blade is chromium steel. In this thesis, it is replaced with Inconel 718. The better material for blade is analyzed.