In this study, the film-cooling effectiveness in different regions of gas turbine blades was investigated with various film hole/slot configurations and mainstream flow conditions. The study consisted of three parts: 1) turbine blade span film-cooling, 2) turbine platform film-cooling, and 3) blade tip film-cooling. Pressure sensitive paint (PSP) technique was used to get the conduction-free film-cooling effectiveness distribution. Film-cooling effectiveness is assessed in terms of cooling hole geometry, blowing ratio, freestream turbulence, and coolant-to-mainstream density ratio. Blade span film-cooling test shows that the compound angle shaped holes offer better film effectiveness than the axial shaped holes. Greater coolant-to-mainstream density ratio prevents coolant to lift-off. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of suction side. Results are also correlated with momentum flux, compound shaped hole has the greatest optimum momentum flux ratio, and then followed by axial shaped hole, compound cylindrical hole, and axial cylindrical hole. For platform purge flow cooling, the stator-rotor gap was simulated by a typical labyrinth-like seal. Two different film-cooling hole geometries, three blowing ratios and density ratios, and two freestream turbulence are examined. Results showed that the shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. Moreover, the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity. The blade tip study was performed in a blow-down flow loop. Results show that a blowing ratio of 2.0 is found to give best results on the tip floor. Lift-off of the coolant jet can be observed for the holes closer to the leading edge as blowing ratio increases from 1.5 to 2.0. A stator vane suction side heat transfer study was conducted in a partial annular cascade. The heat transfer coefficients were measured by using the transient liquid crystal technique. At X/L=0.15, a low heat transfer region where transition occurs. The heat transfer coefficients increase toward the trailing edge as flow accelerates; a spanwise variation can be found at neat tip and hub portions due to passage and horseshoe vortices.

Energy is one of the most important engineering challenges of this time. Gas turbine engines a,re responsible for nearly twenty-percent of all electricity produced in the United States today. A small increase in the operating efficiency of these engines could lead to massive reduction in the emission of greenhouse gases into the atmosphere as well as the financial burden on the average homeowner paying the monthly energy bill. In order to improve the efficiency of the engine, the Turbine Inlet Temperature of the hot gas coming from the combustor is continually increased. This requires increasingly advanced active cooling methods to maintain component life in the hot stages of the turbo machine. In this study, a complete experiment is developed for accurate testing of the complex heat transfer and aerodynamic characteristics present in the active cooling design applied to the transition duct of a land-based gas turbine. The transition duct is the component that channels the hot gases from the combustor to the first stage of the turbihe. It is in contact with the hottest mainstream gas flow in the entire machine. The unique cooling design applied to this component is a combination the three main cooling methods. It is characterized by an impinging jet inlet, which splits into two identical channels flowing in exactly opposite directions. The flow travels through these channels, cooling the hot surfaces of the duct through which they are formed. At the flow exit, it is expelled into the hot gas stream flowing from the can-annular combustor to the turbine stage. The channel exit provides a thin film of cool air coverage that protects the metal surface from the harsh temperatures of the hot gas.