Computations and experiments were run to study adiabatic effectiveness and thermal field contours for a simulated turbine blade leading edge. The RKE and SST k-[omega] turbulence models were used for the computational simulations. Predictions of RKE model for laterally averaged adiabatic effectiveness matched the experimental values. The computational simulations showed different flowfield for the coolant exiting the stagnation line row of holes. Both the experiments and SST k-[omega] simulations predicted coolant separation at the stagnation plane. Also, the downstream spreading of the coolant exiting the stagnation row of exit holes was better predicted by the SST k-[omega] model. At the stagnation plane, experimental thermal field measurements showed greater diffusion of the coolant into the mainstream than predicted by both turbulence models. Reasons for increased diffusion were examined. Thermal field comparison downstream of the offstagnation row of exit holes showed that the computational simulations and the experiments had the same general shape for the offstagnation coolant jet. But the computational simulations predicted greater diffusion of coolant in the direction normal to the surface than seen in the experiments. / text

Gas turbines are widely used for power generation, producing megawatts of usable energy, but consume fossil fuels in order to do so. With gas prices on the rise, all eyes have turned to operating cost and fuel efficiency. To increase efficiency, manufactures raise the temperature of the gas that is combusted. This temperature is high above the melting point of the turbine components. In order for the gas turbine to work under these conditions, its parts must be protected. This study focuses on two aspects of cooling for turbine components. Over the last decades, researchers have investigated many aspects of film cooling, The present study investigates the impact of the stagnation region created by a downstream airfoil on endwall film cooling effectiveness with and without the presence of wake. Experimental measurements are presented for a single row of cylindrical holes inclined at 35° with hole length to diameter ratio, LID= 7.5, pitch to diameter ratio, Pl/D = 3 with a constant density ratio of 1.26, and with nitrogen as the coolant. Twelve different configurations were studied. The airfoil was positioned at X/D equal to 6.35, 12.7, and 25.4. A wake plate was added upstream of the film holes at -12.7 and -50.8 X/D. The effect of stagnation and wake was combined by placing both the airfoil and the wake plate in the test section, combining all positions of each. Baseline cases for the cooling holes alone, and the cooling holes with the airfoil and wake individually were compared to the combined effects. The experimental data shows that as the airfoil stagnation region inhibits film cooling close to the airfoil, and strong wake decreases film effectiveness. With both stagnation region and wake combined, an overall decrease in film cooling performance is observed. Higher blowing ratio increase lateral spreading of the jet promoting jet to jet interaction and mainstream interaction enhancing mixing. The presence of wake promotes jet mixing with the mainstream resulting in lower film cooling effectiveness. High performance turbine airfoils are typically cooled with a combination of internal cooling channels and impingement/film cooling. In such applications, the jets impinge against a target surface, and then exit along the channel formed by the jet plate, target plate, and side walls. Local convection coefficients are the result of both the jet impact, as well as the channel flow produced from the exiting jets. Numerous studies have explored the effects of jet array and channel configurations on both target and jet plate heat transfer coefficients. However, little work has been done in examining effects of height variation and heating on all channel walls, in which both target wall and side wall data is taken, as was neglected by previous literature. This study examines the local and averaged effects of channel height on heat transfer coefficients for target and side walls. High resolution local heat transfer coefficient distributions were measured using temperature sensitive paint and recorded via a scientific grade CCD camera. Streamwise pressure distributions for both the target and side walls was recorded and used to explain heat transfer trends. Results are presented for average jet based Reynolds numbers 17K to 45K. All experiments were carried out on a large scale single row, 15 hole impingement channel, with X/D of 5, YID of 4, and Z/D of 1, 3 and 5. Providing high quality results will aid in the validation of predictive tools and development of physics-based models.

Turbine blade cooling techniques have been around for many years, and the addition of coolant into the turbine blade section will remain to be a viable cooling option for many years to come. Film cooling, which will be the main subject of this research, is a form of convection cooling where holes are placed through the surface of the metal components. With the addition of this film coolant into the main flow, an increased amount of total pressure loss will be found downstream. This is caused by the difference of flow momentum of the coolant and main flow when the two fluids are mixed. The test rig used for the upcoming research will be the NASA-designed E3 rig. E3, standing for Energy Efficient Engine, was established to develop technology for improving the energy efficiency of future commercial transport aircraft engines. These engines were designed to provide real-world, actual test configurations in order to produce more efficient turbine engines, mainly to be used for propulsion. Tests were not focused just on heat transfer as its use will be, but with all aspects of the engine's components, especially aero. The annular cascade with 3-dimensional blade profiles as well as high Reynolds numbers make this setup an accurate test bed in which actual turbine conditions can be compared to. The focus of this research is on the increased amount of total pressure loss seen downstream in an annular cascade with 3D blade profiles with the addition of inner endwall film cooling in a high-speed setting. Also, the rig setup of closed vs. open-loop and its effect on the inlet conditions as well as total pressure loss will be investigated.

Impingement is the most commonly used method of cooling in the hot stages of gas turbines. This is often combined with film cooling to further increase the cooling performance. The mainstream flow where in the coolant films discharge often has large stream wise pressure variations. All existing studies on coupled film and impingement cooling concentrated on the effect of the film depletion on the impingement heat transfer. This study investigates the impact of impingement on film cooling, where the jets impinging on a flat plate are depleted through arrays of film cooling holes in the presence of pressure gradient in the main gas path. The main characteristic of the test setup is that there is an impingement wall on the backside of the film effusion wall. The fluid used for both impingement flow and main flow is air. The impingement flow is heated as opposed to the usual practice of heating mainflow, and the array of film holes are configured under the impingement jet hole arrays such that there is no direct impingement on the film holes. The static pressure variations and Mach number (0.01 to 0.3) in the mainstream underneath the flat plate are controlled by inserts with varying flow area. The detailed temperature distribution on the film-covered surface is measured using the Temperature Sensitive Paint (TSP) technique, and film cooling effectiveness is calculated from the measurements. Results are presented for averaged impingement jet Reynolds numbers of 5000 and 8000. The effect of impingement on film effectiveness is studied by comparing the results from the two cases: one where film flow is directly supplied from a plenum and the other where the post- impingement flow is depleted through film effusion holes. The results are presented for cylindrical film cooling holes which are inclined at angles of 20 deg and 30 deg with respect to the target plate surface. The variation of the effectiveness of the film hole arrays along the mainstream are studied in detail. It is observed that the impingement through jet effects the pressure distribution on the target plate with film holes, which in turn affects the blowing rates of each row. The change in the blowing ratios because of a different pressure distribution on the impingement side of the target plate causes the effectiveness to change. From the results it is observed that the farther rows of impingement are affected by the pressure distribution underneath the film holes and have more flow through the film cooling rows, this increases the inlet flow of the films which increase the blowing ratios and in turn decreases the effectiveness of the film cooling holes. The pressure distribution and the change of effectiveness are studied in detail.

This thesis is an experimental and numerical full-coverage film cooling study. The objective of this work is the quantification of local heat transfer augmentation and adiabatic film cooling effectiveness for two full-coverage film cooling geometries. Experimental data was acquired with a scientific grade CCD camera, where images are taken over the heat transfer surface, which is painted with a temperature sensitive paint. The CFD component of this study served to evaluate how well the v2-f turbulence model predicted film cooling effectiveness throughout the array, as compared with experimental data. The two staggered arrays tested are different from one another through a compound angle shift after 12 rows of holes. The compound angle shifts from ?=-45° to ?=+45° at row 13. Each geometry had 22 rows of cylindrical film cooling holes with identical axial and lateral spacing (X/D=P/D=23). Levels of laterally averaged film cooling effectiveness for the superior geometry approach 0.20, where the compound angle shift causes a decrease in film cooling effectiveness. Levels of heat transfer augmentation maintain values of nearly h/h0=1.2. There is no effect of compound angle shift on heat transfer augmentation observed. The CFD results are used to investigate the detrimental effect of the compound angle shift, while the SST k-? turbulence model shows to provide the best agreement with experimental results.

Electricity has become so ingrained in everyday life that the current generation has no knowledge of life without it. The majority of power generation in the United States is the result of turbines of some form. With such widespread utilization of these complex rotating machines, any increase in efficiency translates into improvements in the current cost of energy. These improvements manifest themselves as reductions in greenhouse emissions or possible savings to the consumer. The most important temperature regarding turbine performance is the temperature of the hot gas entering the turbine, denoted turbine inlet temperature. Increasing the turbine inlet temperature allows for increases in power production as well as increases in efficiency. The challenge with increasing this temperature, currently the hottest temperature seen by the turbine, is that it currently already exceeds the melting point of the metals that the turbine is manufactured from. Active cooling of stationary and rotating components in the turbine is required. Cooling flows are taken from bleed flows from various stages of the compressor as well as flow from the combustor shell. This cooling flow is considered wasted air as far as performance is concerned and can account for as much as 20% of the mass flow in the hot gas path. Lowering the amount of air used for cooling allows for more to be used for performance gain. Various technologies exist to allow for greater turbine inlet temperatures such as various internal channel features inside of turbine blades, film holes on the surface to cool the outside of the airfoil as well as thermal barrier coatings that insulate the airfoils from the hot mainstream iv flow. The current work is a study of the potential performance impact of coupling two effusion technologies, transpiration and discrete hole film cooling. Film cooling and transpiring flows are individually validated against literature before the two technologies are coupled. The coupled geometries feature 13 film holes of 7.5mm diameter and a transpiring strip 5mm long in the streamwise direction. The first coupled geometry features the porous section upstream of the film holes and the second features it downstream. Both geometries use the same crushed aluminum porous insert of nominal porosity of 50%. Temperature sensitive paint along with an ‘adiabatic’ Rohacell surface (thermal conductivity of 0.029W/m-K) are used to measure adiabatic film cooling effectiveness using a scientific grade high resolution CCD camera. The result is local effectiveness data up to 50 film hole diameters downstream of injection location. Data is laterally averaged and compared with the baseline cases. Local effectiveness contours are used to draw conclusions regarding the interactions between transpiration and discrete hole film cooling. It is found that a linear superposition method is only valid far downstream from the injection location. Both coupled geometries perform better than transpiration or the discrete holes far downstream of the injection location. The coupled geometry featuring the transpiring section downstream of the film holes matches the transpiration effectiveness just downstream of injection and surpasses both transpiration and film cooling further downstream.

Full-coverage film cooling is investigated both experimentally and numerically. First, surface measurements local of adiabatic film cooling effectiveness and heat transfer augmentation for four different arrays are described. Reported next is a comparison between two very common turbulence models, Realizable k-ε and SST k-ω, and their ability to predict local film cooling effectiveness throughout a full-coverage array. The objective of the experimental study is the quantification of local heat transfer augmentation and adiabatic film cooling effectiveness for four surfaces cooled by large, both in hole count and in non-dimensional spacing, arrays of film cooling holes. The four arrays are of two different hole-to-hole spacings (P/D = X/D = 14.5, 19.8) and two different hole inclination angles (α = 30◦ , 45◦ ), with cylindrical holes compounded relative to the flow (β = 45◦ ) and arranged in a staggered configuration. Arrays of up to 30 rows are tested so that the superposition effect of the coolant film can be studied. In addition, shortened arrays of up to 20 rows of coolant holes are also tested so that the decay of the coolant film following injection can be studied. Levels of laterally averaged effectiveness reach values as high as ¯η = 0.5, and are not yet at the asymptotic limit even after 20 − 30 rows of injection for all cases studied. Levels of heat transfer augmentation asymptotically approach values of h/h0 ≈ 1.35 rather quickly, iii only after 10 rows. It is conjectured that the heat transfer augmentation levels off very quickly due to the boundary layer reaching an equilibrium in which the perturbation from additional film rows has reached a balance with the damping effect resulting from viscosity. The levels of laterally averaged adiabatic film cooling effectiveness far exceeding ¯η = 0.5 are much higher than expected. The heat transfer augmentation levels off quickly as opposed to the film effectiveness which continues to rise (although asymptotically) at large row numbers. This ensures that an increased row count represents coolant well spent. The numerical predictions are carried out in order to test the ability of the two most common turbulence models to properly predict full coverage film cooling. The two models chosen, Realizable k − ε (RKE) and Shear Stress T ransport k − ω (SSTKW), are both two-equation models coupled with Reynolds Averaged governing equations which make several gross physical assumptions and require several empirical values. Hence, the models are not expected to provide perfect results. However, very good average values are seen to be obtained through these simple models. Using RKE in order to model full-coverage film cooling will yield results with 30% less error than selecting SSTKW.

The modern gas turbine engine requires innovative cooling techniques to protect its internal components from the harsh operating environment typically seen downstream of the combustor. Much research has been performed on the design of these cooling techniques thus allowing for combustion temperatures higher than the melting point of the parts within the turbine. As turbine inlet temperatures and efficiencies continue to increase, it becomes vitally important to correctly and realistically model all of the turbine's external cooling features so as to provide the most accurate representation of the associated heat transfer to the metal surfaces. This study examines the effect of several realistic endwall features for a turbine vane endwall. The first study addresses the effects of a mid-passage gap, endwall misalignment, and roughness on endwall film-cooling. The second study focuses on the effect of varying the combustor-to-turbine gap width. Both studies were performed in a large-scale low speed wind tunnel with the same vane geometry. Geometric and flow parameters were varied and the variation in endwall cooling effectiveness was evaluated. Results from these studies show that realistic features, such as surface roughness, can reduce the effectiveness of endwall cooling designs while other realistic features, such as varying the combustor-to-turbine gap width, can significantly improve endwall cooling effectiveness. It was found that, for a given coolant mass flowrate, a narrow combustor-turbine gap width greatly increased the coverage area of the leaked coolant, even increasing adiabatic effectiveness upstream of the vane stagnation point. The turbine designer can also more efficiently utilize leaked coolant from the combustor-to-turbine gap by controlling endwall misalignment, thereby reducing the overall amount of film-cooling needed for the first stage. / Master of Science

One of the main goals of recent turbine film cooling research has been to improve the overall efficiency of the turbine by slightly increasing film cooling efficiency. This has a twofold effect. Firstly by increasing the effectiveness of the cooling being done. it is possible to increase the inlet temperature of the combusted air coming into the turbine which in tum increases turbine performance. Secondly by increasing the cooling efficiency less air is required. for cooling. This means that less air will be redirected from the compressor for cooling purposes, allowing more air to reach the combustor to be burned and used for power or thrust generation. Even though much bulk flow pulsation research has been conducted in the past, little research has studied the effect of film coolant pulsation on cooling effectiveness. Previous studies that have been conducted on the effect of film pulsation have provided conflicting results, therefore more research is required. This project provides experimental data and analysis which study, and show the effects that low frequency pulsations (5.55 and 11.11 Hz) at two different blowing rations (.5 and .75) have on film cooling effectiveness. In addition a Kulite dynamic pressure probe was placed at the entrance to the coolant holes in order to provide the actual blowing ratio felt by the holes. The study concluded that film pulsation increases film cooling effectiveness mainly through. a reduction in the amount of coolant gas needed to provide adequate film cooling. In addition to providing some initial data, the study also lays the groundwork for additional research to delve further into film pulsation and answer unanswered questions, which will be conducted at a later time.

A segment of transpiring wall is installed near a row of unshaped film holes. The effects on the aerodynamic performance and cooling downstream of the row of cylindrical holes in the presence of transpiration is studied numerically. The changes in behavior of the film due to relative positioning of the injection sources and blowing ratios are predicted to understand the sensitivity of cooling and aerodynamic losses on the relative positioning of the two sources and each blowing ratio. The results indicate that a coupling of the two sources allows a more efficient use of coolant by generating a more uniform initial film resulting in improved component durability through reduction of hot- streaks. With careful optimization the discrete holes can be placed farther apart laterally operating at a lower blowing ratio with a transpiration segment making the large deficits in cooling effectiveness mid-pitch less severe, overall minimizing coolant usage. Addition of transpiration increases the aerodynamic losses associated with injection. This effect can be arguably small compared to corresponding thermal benefits seen by coupling the two. Comparisons of linear superposition predictions of the two independent sources with the corresponding coupled scenario indicate the two films positively influence one another and outperform predictions. The interaction between the two films is dependent upon the relative placement of the transpiration; all relative placements have an overall beneficial effect on the cooling seen by the protected wall. An increase in area-averaged film cooling effectiveness of 300% is seen along with only a 50% increase in loss coefficient by injecting an additional 10% coolant. In this study the downstream placement of transpiration is found to perform best of the three geometries tested while considering cooling, aerodynamic losses, local uniformity and manufacturing feasibility. With further study and optimization this technique can potentially provide more effective thermal protection at a lower cost of aerodynamic losses and spent coolant. A method of measuring the local temperature of a porous wall is also discussed. Measurements are taken with temperature sensitive paint applied in thin coats to the wall. This technique was validated on a 40PPI, 7% relative density aluminum porous coupon. Measurements of discharge coefficients as well as downstream effectiveness data are included to verify the flow through the porous wall was unaltered by applying the paint. A maximum deviation in film-cooling effectiveness of 9% between the two cases with the majority of data falling within 4% was found, very similar to the experimental uncertainty of the rig. This excellent agreement between the repeated tests showed that by applying thermal paint to a wall of such porosity does not significantly affect the flow exiting the wall and hence the measurement technique can readily be applied to transpiration cooling studies at this scale. Methods of filtering the temperature sensitive paint on the porous wall are presented.