Browsing by Subject "Turbine cooling"
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Item Experimental investigation of film cooling and thermal barrier coatings on a gas turbine vane with conjugate heat transfer effects(2013-05) Kistenmacher, David Alan; Bogard, David G.In the United States, natural gas turbine generators account for approximately 7% of the total primary energy consumed. A one percent increase in gas turbine efficiency could result in savings of approximately 30 million dollars for operators and, subsequently, electricity end-users. The efficiency of a gas turbine engine is tied directly to the temperature at which the products of combustion enter the first stage, high-pressure turbine. The maximum operating temperature of the turbine components’ materials is the major limiting factor in increasing the turbine inlet temperature. In fact, current turbine inlet temperatures regularly exceed the melting temperature of the turbine vanes through advanced vane cooling techniques. These cooling techniques include vane surface film cooling, internal vane cooling, and the addition of a thermal barrier coating (TBC) to the exterior of the turbine vane. Typically, the performance of vane cooling techniques is evaluated using the adiabatic film effectiveness. However, the adiabatic film effectiveness, by definition, does not consider conjugate heat transfer effects. In order to evaluate the performance of internal vane cooling and a TBC it is necessary to consider conjugate heat transfer effects. The goal of this study was to provide insight into the conjugate heat transfer behavior of actual turbine vanes and various vane cooling techniques through experimental and analytical modeling in the pursuit of higher turbine inlet temperatures resulting in higher overall turbine efficiencies. The primary focus of this study was to experimentally characterize the combined effects of a TBC and film cooling. Vane model experiments were performed using a 10x scaled first stage inlet guide vane model that was designed using the Matched Biot Method to properly scale both the geometrical and thermal properties of an actual turbine vane. Two different TBC thicknesses were evaluated in this study. Along with the TBCs, six different film cooling configurations were evaluated which included pressure side round holes with a showerhead, round holes only, craters, a novel trench design called the modified trench, an ideal trench, and a realistic trench that takes manufacturing abilities into account. These film cooling geometries were created within the TBC layer. Each of the vane configurations was evaluated by monitoring a variety of temperatures, including the temperature of the exterior vane wall and the exterior surface of the TBC. This study found that the presence of a TBC decreased the sensitivity of the thermal barrier coating and vane wall interface temperature to changes in film coolant flow rates and changes in film cooling geometry. Therefore, research into improved film cooling geometries may not be valuable when a TBC is incorporated. This study also developed an analytical model which was used to predict the performance of the TBCs as a design tool. The analytical prediction model provided reasonable agreement with experimental data when using baseline data from an experiment with another TBC. However, the analytical prediction model performed poorly when predicting a TBC’s performance using baseline data collected from an experiment without a TBC.Item Experimental simulation and mitigation of contaminant deposition on film cooled gas turbine airfoils(2011-05) Albert, Jason Edward; Bogard, David G.; da Silva, Alexandre K.; Ezekoye, Ofodike A.; Webber, Michael E.; Wenglarz, Richard A.Deposition of contaminant particles on gas turbine surfaces reduces the aerodynamic and cooling efficiency of the turbine and degrades its materials. Gas turbine designers seek a better understanding of this complicated phenomenon and how to mitigate its effects on engine efficiency and durability. The present study developed an experimental method in wind tunnel facilities to simulate the important physical aspects of the interaction between deposition and turbine cooling, particularly film cooling. This technique consisted of spraying molten wax droplets into the mainstream flow that would deposit and solidify on large scale, cooled, turbine airfoil models in a manner consistent with inertial deposition on turbine surfaces. The wax particles were sized to properly simulate the travel of particles in the flow path, and their adhesion to the surface was modeled by ensuring they remained at least partially molten upon impact. Initial development of this wax spray technique was performed with a turbine blade leading edge model with three rows of showerhead film cooling. It was then applied to turbine vane models with showerhead holes and row on pressure side consisting of either standard cylindrical holes or similar holes situated in a spanwise, recessed trench. Vane models were either approximately adiabatic or had a thermal conductivity selected to simulate the conjugate heat transfer of turbine airfoils at engine conditions. These models were also used to measure the adiabatic film effectiveness and overall cooling effectiveness in order to better assess how the cooling design interacted with deposition. Deposit growth was found to be sensitive to the mainstream air and the model surface temperatures and the solidification temperature of the wax. Deposits typically grew to an equilibrium thickness caused by a balance between erosion and adhesion. The existence of film cooling substantially redistributed deposit growth, but changes in blowing ratio had a minor effect. A hypothesis was proposed and substantiated for the physical mechanisms governing wax deposit growth, and its applicability to engine situations was discussed.Item Using contoured endwalls to achieve proper scaling for a gas turbine vane model using a low speed testing facility(2015-05) Vaclavik, Adam William; Bogard, David G.; Ezekoye, OfodikeThe testing of gas turbine vane and blade models is often performed in low speed, large scale infinite cascade facilities to allow for more precise machining of parts and more accurately measured data. However, flow in engine scale turbines reaches well into the compressible gas range while low speed facilities run in the incompressible fluid range, and engines have three dimensional flow effects due to having contoured endwall while traditional cascade testing has not accounted for three dimensional effects. This means that matching pressure distributions cannot be achieved between engine scale and experimental scale through simple geometric scaling of the model. In the past, these differences in pressure distributions were often overcome by changing the geometry of the test model. An alternate to this is to use contoured enwalls inside the test facility to allow the decreased area to correct for the differences in pressure distributions. In this work, the concept of using contoured endwalls in the test facility to achieve a matching pressure distribution on a vane was tested. Three dimensional computation fluid dynamics (CFD) simulations were used to find the correct geometry for the contoured endwalls. The proposed endwalls and vanes were then built and tested in a low speed simulated infinite cascade testing facility. The pressure distribution was measured at low turbulence levels and Re = 1.1×10⁶. It was shown that the pressure distribution in the test model with contoured endwalls did match within uncertainty the pressure distribution predicted for the engine scale using CFD. Thus, contoured endwalls can be said to be a viable option to force the matching of pressure distribution of a model test vane to that of engine conditions. Additionally, a vane model with a constant heat flux surface was tested at the same conditions, and the heat transfer coefficient distribution for the vane was determined. It was shown that the endwalls had minimal effects on the spanwise uniformity of the heat transfer coefficient distribution.