Advanced cooling techniques are essential for further improvement in the efficiency and the power output of gas turbines. Turbine inlet temperatures of 1900 K are typical of current gas turbines, and there is an interest in increasing the temperatures for the next generation of gas turbine engines. Over the past decades, significant effort has been devoted to increase the turbine efficiency and to develop effective cooling strategies to maintain the blade temperature below the melting point of the alloys used to construct the airfoils. As a result, various cooling strategies have been developed such as film, impingement, and muti-pass cooling for the blades, and evaporative cooling for the inlet air. In this work, a state-of-the-art thermal turbomachinery test rig was designed and constructed to investigate the film-cooling performance of advanced film cooling schemes over a flat plate. Designing and constructing mechanical parts, as well developing software codes (Labview and image processing) for transient film cooling measurement was the foremost part of the current experimental work. The thermochromic liquid crystal (TLC) technique was used to measure wall surface temperature. A circular film hole was used to validate the current experimental technique and methodology. The validation results showed that the current experimental technique and methodology were deemed reliable. Subsequently, the film cooling performance of the louver and new hybrid schemes were investigated, experimentally. The louver scheme was proposed by Pratt and Whitney Canada (PWC) to allow the cooling flow to pass through a bend and to encroach an airfoil material (impingement effect), then exit to the outer surface of the airfoil through a designed film hole. Immarigeon and Hassan (2006) then Zhang and Hassan (2006) numerically investigated the film cooling effectiveness performance of the louver scheme. The hybrid scheme was proposed in the current study, which includes two consecutive film hole configurations with interior bending. The cooling performances for the two advanced schemes have been analyzed experimentally over a flat plate across blowing ratios of 0.5, 1.0 and 1.5 at a density ratio of 0.94. The results showed that the louver and the hybrid schemes enhanced the local and the average film cooling performance in terms of film cooling effectiveness, and the net heat flux reductions are better than other published film hole configurations. In addition, both schemes provided an extensively wide spray of 'secondary flow over the outer surface, and thus enhanced the lateral film cooling performance over the downstream surface area. Moreover, the two schemes produced an average heat transfer coefficient ratio near unity at low and high blowing ratios. As a result, the louver and the hybrid schemes are expected to reduce the temperature of the outer surface of the gas turbine airfoil and to provide superior cooling performance, which increases airfoil lifetime. In addition, the adiabatic film cooling performance and flow characteristics for the hybrid scheme were investigated numerically. The numerical investigation was analyzed across blowing ratio, of 0.5, 1, and 2. The flow structures of the hybrid scheme are presented at different blowing ratios to provide a better physical understanding. The results showed that the hybrid scheme directed the secondary flow in the horizontal direction and reduced the jet liftoff at different blowing ratios. Finally, conjugate heat transfer (CHT) and film-cooling analyses were performed to investigate the hybrid scheme performance with different flow configurations. Different geometries of parallel flow and jet impingement with different gap heights as well as the adiabatic case study were investigated at blowing ratios of 0.5 and 1.0. The results showed that the adiabatic case provided downstream centerline superlative cooling performance near the hybrid film hole exit compared to other conjugate geometries studied. At the downstream location, the impingement configuration with a large gap height provided the highest downstream performance at blowing ratio of 0.5 and 1.0 with respect to other cases studied. Moreover, the downstream film cooling performance was enhanced far along the spanwise direction for the CHT cases studied and it has the highest value near the scheme exit for parallel configuration. In addition, the impingement configuration enhanced the upper stream cooling performance compared to parallel flow and it was further enhanced for large gap heights. Keywords: film cooling effectiveness, heat transfer coefficient ratio, louver, hybrid, TLC, NHFR, CHT.

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

The present work has been an experimental investigation to evaluate the applicability of gas turbine cooling technology. With the temperature of the mainstream gas entering the turbine elevated above the melting temperature of the metal components, these components must be cooled, so they can withstand prolonged exposure to the mainstream gas. Both external and internal cooling techniques have been studied as a means to increase the life of turbine components. Detailed film cooling effectiveness distributions have been obtained on the turbine blade platform with a variety of cooling configurations. Because the newly developed pressure sensitive paint (PSP) technique has proven to be the most suitable technique for measuring the film effectiveness, it was applied to a variety of platform seal configurations and discrete film flows. From the measurements it was shown advanced seals provide more uniform protection through the passage with less potential for ingestion of the hot mainstream gases into the engine cavity. In addition to protecting the outer surface of the turbine components, via film cooling, heat can also be removed from the components internally. Because the turbine blades are rotating within the engine, it is important to consider the effect of rotation on the heat transfer enhancement within the airfoil cooling channels. Through this experimental investigation, the heat transfer enhancement has been measured in narrow, rectangular channels with various turbulators. The present experimental investigation has shown the turbulators, coupled with the rotation induced Coriolis and buoyancy forces, result in non-uniform levels of heat transfer enhancement in the cooling channels. Advanced turbulator configurations can be used to provide increased heat transfer enhancement. Although these designs result in increased frictional losses, the benefit of the heat transfer enhancement outweighs the frictional losses.

Summary: Transpiration and film cooling promise to be effective methods of cooling gas-turbine blades; consequently, analytical and experimental investigations are being conducted to obtain a better understanding of these processes. This report serves as an introduction to these cooling methods, explains the physical processes, and surveys the information available for predicting blade temperatures and heat-transfer rates. In addition, the difficulties encountered in obtaining a uniform blade temperature are discussed, and the possibilities of correcting these difficulties are indicated. Air is the only coolant considered in the application of these cooling methods.

The focus of this program was to experimentally investigate advanced gas turbine cooling schemes and the effects of and factors that contribute to surface deposition from particulate matter found in coal syngas exhaust flows on turbine airfoil heat transfer and film cooling, as well as to characterize surface roughness and determine the effects of surface deposition on turbine components. The program was a comprehensive, multi-disciplinary collaborative effort between aero-thermal and materials faculty researchers and the Department of Energy, National Energy Technology Laboratory (NETL). The primary technical objectives of the program were to evaluate the effects of combustion of syngas fuels on heat transfer to turbine vanes and blades in land-based power generation gas turbine engines. The primary questions to be answered by this investigation were; What are the factors that contribute to particulate deposition on film cooled gas turbine components? An experimental program was performed in a high-temperature and pressure combustion rig at the DOE NETL; What is the effect of coal syngas combustion and surface deposition on turbine airfoil film cooling? Deposition of particulate matter from the combustion gases can block film cooling holes, decreasing the flow of the film coolant and the film cooling effectiveness; How does surface deposition from coal syngas combustion affect turbine surface roughness? Increased surface roughness can increase aerodynamic losses and result in decreased turbine hot section efficiency, increasing engine fuel consumption to maintain desired power output. Convective heat transfer is also greatly affected by the surface roughness of the airfoil surface; Is there any significant effect of surface deposition or erosion on integrity of turbine airfoil thermal barrier coatings (TBC) and do surface deposits react with the TBC in any way to decrease its thermal insulating capability? Spallation and erosion of TBC is a persistent problem in modern turbine engines; and What advancements in film cooling hole geometry and design can increase effectiveness of film cooling in turbines burning high-hydrogen coal syngas due to the higher heat loads and mass flow rates of the core flow? Experimental and numerical investigations of advanced cooling geometries that can improve resistance to surface deposition were performed. The answers to these questions were investigated through experimental measurements of turbine blade surface temperature and coolant coverage (via infrared camera images and thermocouples) and time-varying surface roughness in the NETL high-pressure combustion rig with accelerated, simulated surface deposition and advanced cooling hole concepts, coupled with detailed materials analysis and characterization using conventional methods of Scanning Electron Microscopy (SEM), Transmission Electron Microscopy (TEM), X-Ray Diffraction (XRD), 3-D Surface Topography (using a 3-D stylus profilometer). Detailed surface temperatures and cooling effectiveness could not be measured due to issues with the NETL infrared camera system. In collaboration with faculty startup funding from the principal investigator, experimental and numerical investigations were performed of an advanced film cooling hole geometry, the anti-vortex hole (AVH), focusing on improving cooling effectiveness and decreasing the counter-rotating vortex of conventional cooling holes which can entrain mainstream particulate matter to the surface. The potential benefit of this program is in gaining a fundamental understanding of how the use of alternative fuels will effect the operation of modern gas turbine engines, providing valuable data for more effective cooling designs for future turbine systems utilizing alternative fuels.