The hot gas temperature in gas turbine engines is far above the permissible metal temperatures. Advanced cooling technologies must be applied to cool the blades, so they can withstand the extreme conditions. Film cooling is widely used in modern high temperature and high pressure blades as an active cooling scheme. 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 four parts: 1) effect of upstream wake on blade surface film cooling, 2) effect of upstream vortex on platform purge flow cooling, 3) influence of hole shape and angle on leading edge film cooling and 4) slot film cooling on trailing edge. Pressure sensitive paint (PSP) technique was used to get the conduction-free film cooling effectiveness distribution. For the blade surface film cooling, the effectiveness from axial shaped holes and compound angle shaped holes were examined. Results showed that the compound angle shaped holes offer better film effectiveness than the axial shaped holes. The upstream stationary wakes have detrimental effect on film effectiveness in certain wake rod phase positions. For platform purge flow cooling, the stator-rotor gap was simulated by a typical labyrinth-like seal. Delta wings were used to generate vortex and modeled the passage vortex generated by the upstream vanes. Results showed that the upstream vortex reduces the film cooling effectiveness on the platform. For the leading edge film cooling, two film cooling designs, each with four film cooling hole configurations, were investigated. Results showed that the shaped holes provide higher film cooling effectiveness than the cylindrical holes at higher average blowing ratios. In the same range of average blowing ratio, the radial angle holes produce better effectiveness than the compound angle holes. The seven-row design results in much higher effectiveness than the three-row design. For the trailing edge slot cooling, the effect of slot lip thickness on film effectiveness under the two mainstream conditions was investigated. Results showed thinner lips offer higher effectiveness. The film effectiveness on the slots reduces when the incoming mainstream boundary layer thickness decreases.

High turbine inlet temperature becomes necessary for increasing thermal efficiency of modern gas turbines. To prevent failure of turbine components, advance cooling technologies have been applied to different portions of turbine blades. The detailed film cooling effectiveness distributions along a rotor blade has been studied under combined effects of upstream trailing edge unsteady wake with coolant ejection by the pressure sensitive paint (PSP). The experiment is conducted in a low speed wind tunnel with a five blade linear cascade and exit Reynolds number is 370,000. The density ratios for both blade and trailing edge coolant ejection range from 1.5 to 2.0. Blade blowing ratios are 0.5 and 1.0 on suction surface and 1.0 and 2.0 on pressure surface. Trailing edge jet blowing ratio and Strouhal number are 1.0 and 0.12, respectively. Results show the unsteady wake reduces overall effectiveness. However, the unsteady wake with trailing edge coolant ejection enhances overall effectiveness. Results also show that the overall effectiveness increases by using heavier coolant for ejection and blade film cooling. Leading edge film cooling has been investigated using PSP. There are two test models: seven and three-row of film holes for simulating vane and blade, respectively. Four film holes' configurations are used for both models: radial angle cylindrical holes, compound angle cylindrical holes, radial angle shaped holes, and compound angle shaped holes. Density ratios are 1.0 to 2.0 while blowing ratios are 0.5 to 1.5. Experiments were conducted in a low speed wind tunnel with Reynolds number 100,900. The turbulence intensity near test model is about 7%. The results show the shaped holes have overall higher effectiveness than cylindrical holes for both designs. As increasing density ratio, density effect on shaped holes becomes evident. Radial angle holes perform better than compound angle holes as increasing blowing and density ratios. Increasing density ratio generally increases overall effectiveness for all configurations and blowing ratios. One exception occurs for compound angle and radial angle shaped hole of three-row design at lower blowing ratio. Effectiveness along stagnation row reduces as increasing density ratio due to coolant jet with insufficient momentum caused by heavier density coolant, shaped hole, and stagnation row. The electronic version of this dissertation is accessible from http://hdl.handle.net/1969.1/148288