Numerical Analysis of Cooling Performance of 1st Stage Gas Turbine With Damaged Thermal Barrier Coatings

The compound gas burnt in the combustor passes into the turbine, its temperature is highly correlated with the turbine efficiency. The modern turbine components are being operated at excessively high turbine inlet temperatures. Therefore, adequate cooling techniques should be exploited to protect the components and can be properly operated within the allowed range against hot gases. In gas turbine cooling technology, thermal barrier coating (TBC) is designed with the material having a low heat transfer coefficient. Film cooling allows the cooling air to apply internal channel cooling and is sprayed through the holes to reduce the temperature rise from the hot gas on the surface. In this study, the cooling performance of the gas turbine 1^st vanes and blades is analyzed according to the thickness of the TBC and film cooling. The TBC on the outer surface of the 1^st gas turbine is applied to the NiCrAlY of bond coating and the YSZ of top coating. Numerical simulation results show that the surface temperature of the vanes and blades increases as the thickness of the TBC decreases with the same film cooling condition. The thickness of the thermal barrier coating also affects the flow characteristics around the turbine components. We suggest that our numerical simulation of cooling performance according to TBC thickness a basis for selecting coating thickness for gas turbines with increased turbine inlet temperature.