A Review of Secondary Combustion on Turbine Blade Cooling

Abstract The gas turbine engine is the powerhouse of most large modern military and civil aircraft. These engines operate at temperatures above the melting point of the materials that the combustor and turbine components are made from. Film cooling is used extensively to cool the hot surfaces and extend the life of the gas turbine’s hot end components. In some modern and future engines, the average turbine inlet temperature is increased to about 2,400K and the length of the combustor is reduced. The turbine inlet temperature is increased to improve the thermal efficiency while the combustor is shortened to increase the thrust-to-weight ratio. Both developments are meant to reduce the amount of fuel burnt and the operational cost of the power plant. Increasing the turbine inlet temperature to above 1,850 K CO2 dissociation starts to compete with CO oxidation. Reducing the combustor length reduces the residence time of fuel and increases the likelihood of unburnt hydrocarbons entering the turbine. When carbon monoxide and/or unburnt hydrocarbons enter the turbine, they could react with oxygen in the cooling air and potentially increase the blade metal temperature. An increase of about 30 K can reduce the blade life by half: secondary combustion of reactive species entering the turbine section could therefore lead to serious durability concerns. In a review of the literature, it was found that an estimated 10% of fuel energy is available for combustion in the turbine section and a maximum heat flux augmentation of 18% due to secondary combustion occurs. Secondary combustion in the turbine components is reviewed through a discussion of the analysis of reactive film cooling, developments driving the need to develop an in-depth understanding of reactive film cooling, scaling of reaction kinetics and heat release potential, performance of cooling hole geometries and configurations and mitigation strategies.