In the drive to increase cycle efficiency, gas turbine designers have increased turbine inlet temperatures well beyond the metallurgical limits of engine components. In order to prevent failure and meet life requirements, turbine components must be cooled well below these hot gas temperatures. Film cooling is a widely employed cooling technique whereby air is extracted from the compressor and ejected through holes on the surfaces of hot gas path components. The cool air forms a protective film around the surface of the part. Accurate numerical prediction of film cooling performance is extremely difficult so experiments are required to validate designs and CFD tools. In this study, a first stage turbine vane with five rows of showerhead cooling was instrumented with platinum thin-film gauges to experimentally characterize film cooling performance. The vane was tested in a transonic vane cascade in Virginia Tech's heated, blow-down wind tunnel. Two freestream exit Mach numbers of 0.76 and 1.0—corresponding to exit Reynolds numbers based on vane chord of 1.1x106 and 1.5x106, respectively—were tested at an inlet freestream turbulence intensity of two percent and an integral length scale normalized by vane pitch of 0.05. The showerhead cooling scheme was tested at blowing ratios of 0 (no cooling), 1.5, and 2.0 and a density ratio of 1.35. Midspan Nusselt number and film cooling effectiveness distributions over the surface of the vane are presented. Film cooling was found to augment heat transfer and reduce adiabatic wall temperature downstream of injection. In general, an increase in blowing ratio was shown to increase augmentation and film cooling effectiveness. Increasing Reynolds number was shown to increase heat transfer and reduce effectiveness. Finally, comparing low turbulence measurements (Tu = 2%) to measurements performed at high freestream turbulence (Tu = 16%) by Nasir et al. [13] showed that large-scale high freestream turbulence can reduce heat transfer coefficient downstream of injection. / Master of Science

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.

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.

In modern, low emission, gas turbine combustion systems the amount of air available for cooling of the flame tube liner is limited. This has led to the development of more complex cooling systems such as cooling tiles i.e. a double skin system, as opposed to the use of more conventional cooling slots i.e. a single skin system. An isothennal experimental facility has been constructed which can incorporate 10 times full size single and double skin (cooling tile) test specimens. The specimens can be tested with or without effusion cooling and measurements have been made to characterise the flow through each cooling system along with the velocity field and cooling effectiveness distributions that subsequently develop along the length of each test section. The velocity field of the coolant film has been defined using pneumatic probes, hot-wire anemometry and PIV instrumentation, whilst gas tracing technique is used to indicate (i) the adiabatic film cooling effectiveness and (ii) mixing of the coolant film with the mainstream flow. Tests have been undertaken both with a datum low turbulence mainstream flow passing over the test section, along with various configurations in which large magnitudes and scales of turbulence were present in the mainstream flow. These high turbulence test cases simulate some of the flow conditions found within a gas turbine combustor. Results are presented relating to a variety of operating conditions for both types of cooling system. The nominal operating condition for the double skin system was at a coolant to mainstream blowing ratio of approximately 1.0. At this condition, mixing of the mainstream and coolant film was relatively small with low mainstream turbulence. However, at high mainstream turbulence levels there was rapid penetration of the mainstream flow into the coolant film. This break up of the coolant film leads to a significant reduction in the cooling effectiveness. In addition to the time-averaged characteristics, the time dependent behaviour of the .:coolantfilm was. also investigated. In particular, unsteadiness associated with large scale structures in the mainstream flow was observed within the coolant film and adjacent to the tile surface. Relative to a double skin system the single skin geometry requires a higher coolant flow rate that, along with other geometrical changes, results in typically higher coolant to mainstream velocity ratios. At low mainstream turbulence levels this difference in velocity between the coolant and mainstream promotes the generation of turbulence and mixing between the streams so leading to some reduction in cooling effectiveness. However, this higher momentum coolant fluid is more resistant to high mainstream turbulence levels and scales so that the coolant film break up is not as significant under these conditions as that observed for the double skin system. For all the configurations tested the use of effusion cooling helped restore the coolant film along the rear of the test section. For the same total coolant flow, the minimum value of cooling effectiveness observed along the test section was increased relative to the no effusion case. In addition the effectiveness of the effusion patch depends on the amount of coolant injected and the axial location of the patch. The overall experimental data suggested the importance of the initial cooling film conditions together with better understanding of the possible mechanisms that results in the rapid cooling film break-up, such as high turbulence mainstream flow and scales, and this will lead to a more effective cooling system design. This experimental data is also thought to be ideal for the validation of numerical predictions.

The turbine blade platform can be protected from hot mainstream gases by injecting cooler air through the gap between stator and rotor. The effectiveness of this film cooling method depends on the geometry of the slot, the quantity of injected air, and the secondary flows near the platform. The purpose of this study was to measure the effect of the upstream vane or stator on this type of platform cooling, as well as the effect of changes in the width of the gap. Film cooling effectiveness distributions were obtained on a turbine blade platform within a linear cascade with upstream slot injection. The width of the slot was varied as well as the mass flow rate of the injected coolant. Obstacles were placed upstream to model the effect of the upstream vane. The coolant was injected through an advanced labyrinth seal to simulate purge flow through a stator-rotor seal. The width of the opening of this seal was varied to simulate the effect of misalignment. Stationary rods were placed upstream of the cascade in four phase locations to model the unsteady wake formed at the trailing edge of the upstream vane. Delta wings were also placed in four positions to create a vortex similar to the passage vortex at the exit of the vane. The film cooling effectiveness distributions were measured using pressure-sensitive paint (PSP). Reducing the width of the slot was found to decrease the area of coolant coverage, although the film cooling effectiveness close to the slot was slightly increased. The unsteady wake was found to have a trivial effect on platform cooling, while the passage vortex from the upstream vane may significantly reduce the film cooling effectiveness.

En vue de leur refroidissement, les parois de chambres de combustion aéronautiques sont perforées de trous à travers lesquels de l’air plus froid est injecté. La paroi est ainsi refroidie par convection et un film isolant est créé en surface chaude (film cooling). Cette thèse a pour objectif d’utiliser les possibilités de la fabrication additive pour proposer de nouvelles solutions architecturées qui permettraient d’augmenter les échanges de chaleur internes et d’obtenir ainsi de meilleures efficacités de refroidissement.La première approche consiste à élaborer de nouveaux designs de plaques multiperforées par Electron Beam Melting (EBM) et Selective Laser Melting (SLM) aux limites de résolution des procédés. Les architectures sont caractérisées en microscopie, en tomographie X et en perméabilité. Des simulations aérothermiques permettent de mettre en évidence l’effet de ces nouveaux designs sur l’écoulement et les échanges de chaleur, et de proposer des voies d’amélioration de la géométrie.La deuxième approche consiste à élaborer de façon simultanée une pièce architecturée par EBM, avec des zones denses et poreuses. A partir d’analyse d’images associée à une cartographie EBSD grand champ, il est possible de remonter aux mécanismes de formation du matériau poreux et de relier la perméabilité et la porosité aux paramètres procédé. Afin de favoriser le film cooling, il pourrait être avantageux que les zones microporeuses soient orientées dans le sens de l’écoulement. Pour ce faire, un nouveau procédé dénommé Magnetic Freezing, où des poudres métalliques forment une structure orientée par un champ magnétique, est mis au point.Les diverses solutions développées durant cette thèse sont testées sur un banc aérothermique. Les essais montrent qu’elles offrent un refroidissement plus efficace et plus homogène que la référence industrielle. Enfin, de premiers tests en combustion sur l’une des structures retenues, plus légère et plus perméable que la référence, montrent qu’il s’agit d’une solution aussi efficace à un débit traversant donné, et donc a priori plus efficace à une surpression donnée. / Combustion chamber walls are perforated with holes so that a cooling air flow can be injected through them. The wall is cooled by convection and an insulating film is created on the hot surface (film cooling). This PhD thesis aims to use the possibilities of additive manufacturing to provide new architectured solutions that could enhance the internal heat exchanges, and lead to a higher cooling effectiveness.The first approach is to develop new designs of multiperforated walls by Electron Beam Melting (EBM) and Selective Laser Melting (SLM) used at the resolution limits of the processes. They are characterized by microscopy, X-ray tomography and permeability tests. Some aerothermal simulations help understanding the effects of these new designs on the flow and on heat exchanges. These results lead to a geometry adaptation.The second approach is to simultaneously manufacture an architectured part with dense and porous zones by EBM. Thanks to image analysis combined with large field EBSD, it is possible to investigate the mechanisms leading to the porous zones and to link them to permeability and porosity. The film cooling effect could be favoured by the orientation of pores towards the cooling flow. Therefore, a new powder-based manufacturing process named Magnetic Freezing, where metallic powders organize into an oriented structure thanks to a magnetic field, is developed.The various solutions studied during this thesis are tested on an aerothermal bench. They all show a more efficient and homogeneous cooling than the industrial reference. Some first tests on one of the selected solutions are performed on a combustion bench. This lighter and more permeable structure proves to be a solution as efficient as the industrial reference at a given flow rate. It should therefore be a more efficient solution for a given overpressure.

This research work investigates the thermal performance of a film-cooled gas turbine endwall under two different mainstream flow conditions. In the first part of the research investigation, the effect of unsteady passing wakes on a film-cooled pitchwise-curved surface (representing an endwall without airfoils) was experimentally studied for heat transfer characteristics on a time-averaged basis. The temperature sensitive paint technique was used to obtain the local temperatures on the test surface. The required heat flux input was provided using foil heaters. Discrete film injection was implemented on the test surface using cylindrical holes with a streamwise inclination angle of 35? and no compound angle relative to the mean approach velocity vector. The passing wakes increased the heat transfer coefficients at both the wake passing frequencies that were experimented. Due to the increasing film cooling jet turbulence and strong jet-mainstream interaction at higher blowing ratios, the heat transfer coefficients were amplified. A combination of film injection and unsteady passing wakes resulted in a maximum pitch-averaged and centerline heat transfer augmentation of ? 28% and 31.7% relative to the no wake and no film injection case. The second part of the research study involves an experimental and numerical analysis of secondary flow and coolant film interaction in a high subsonic annular cascade with a maximum isentropic throat Mach number of ? 0.68. Endwall (platform) thermal protection is provided using discrete cylindrical holes with a streamwise inclination angle of 30? and no compound angle relative to the mean approach velocity vector. The surface flow visualization on the inner endwall provided the location of the saddle point and the three-dimensional separation lines. Computational predictions showed that the leading-edge horseshoe vortex was confined to approximately 1.5% of the airfoil span for the no film injection case and intensified with low momentum film injection. At the highest blowing ratio, the film cooling jet weakened the horseshoe vortex at the leading-edge plane. The passage vortex was intensified with coolant injection at all blowing ratios. It was seen that increasing average blowing ratio improved the film effectiveness on the endwall. The discharge coefficients calculated for each film cooling hole indicated significant non-uniformity in the coolant discharge at lower blowing ratios and the strong dependence of discharge coefficients on the mainstream static pressure and the location of three-dimensional separation lines. Near the airfoil suction side, a region of coalesced film cooling jets providing close to uniform film coverage was observed, indicative of the mainstream acceleration and the influence of three-dimensional separation lines.

The study presented herein describes and analyzes a detailed experimental investigation of the effects of density ratio on endwall thermal performance at varying blowing rates for a typical nozzle guide vane platform purge jet cooling scheme. An axisymmetric converging endwall with an upstream doublet staggered cylindrical hole purge jet cooling scheme was employed. Nominal exit flow conditions were engine representative and as follows: {rm Ma}_{Exit} = 0.85, {rm Re}_{Exit,C_{ax}} = 1.5 times {10}^6, and large-scale freestream Tu = 16%. Two blowing ratios were investigated corresponding to the upper and lower engine extrema. Each blowing ratio was investigated amid two density ratios; one representing typical experimental neglect of density ratio, at DR = 1.2, and another engine representative density ratio achieved by mixing foreign gases, DR = 1.95. All tests were conducted on a linear cascade in the Virginia Tech Transonic Blowdown Wind Tunnel using IR thermography and transient data reduction techniques. Oil paint flow visualization techniques were used to gather quantitative information regarding the alteration of endwall flow physics due two different blowing rates of high-density coolant. High resolution endwall adiabatic film cooling effectiveness, Nusselt number, and Net Heat Flux Reduction contour plots were used to analyze the thermal effects. The effect of density is dependent on the coolant blowing rate and varies greatly from the high to low blowing condition. At the low blowing condition better near-hole film cooling performance and heat transfer reduction is facilitated with increasing density. However, high density coolant at low blowing rates isn't adequately equipped to penetrate and suppress secondary flows, leaving the SS and PS largely exposed to high velocity and temperature mainstream gases. Conversely, it is observed that density ratio only marginally affects the high blowing condition, as the momentum effects become increasingly dominant. Overall it is concluded density ratio has a first order impact on the secondary flow alterations and subsequent heat transfer distributions that occur as a result of coolant injection and should be accounted for in purge jet cooling scheme design and analysis. Additionally, the effect of increasing high density coolant blowing rate was analyzed. Oil paint flow visualization indicated that significant secondary flow suppression occurs as a result of increasing the blowing rate of high-density coolant. Endwall adiabatic film cooling effectiveness, Nusselt number, and NHFR comparisons confirm this. Low blowing rate coolant has a more favorable thermal impact in the upstream region of the passage, especially near injection. The low momentum of the coolant is eventually dominated and entrained by secondary flows, providing less effectiveness near PS, near SS, and into the throat of the passage. The high momentum present for the high blowing rate, high-density coolant suppresses these secondary flows and provides enhanced cooling in the throat and in high secondary flow regions. However, the increased turbulence impartation due to lift off has an adverse effect on the heat load in the upstream region of the passage. It is concluded that only marginal gains near the throat of the passage are observed with an increase in high density coolant blowing rate, but severe thermal penalty is observed near the passage onset. / Master of Science / Gas turbine technology is used frequently in the burning of natural gas for power production. Increases in engine efficiency are observed with increasing firing temperatures, however this leads to the potential of overheating in the stages following. To prevent failure or melting of components, cooler air is extracted from the upstream compressor section and used to cool these components through various highly complex cooling schemes. The design and operational adequacy of these schemes is highly subject to the mainstream and coolant flow conditions, which are hard to represent in a laboratory setting. This experimental study explores the effects of various coolant conditions, and their respective response, for a purge jet cooling scheme commonly found in engine. This scheme utilizes two rows of staggered cylindrical holes to inject air into the mainstream from platform, upstream of the nozzle guide vane. It is the hope that this air forms a protective layer, effectively shielding the platform from the hostile mainstream conditions. Currently, little research has been done to quantify these effects of purge flow cooling scheme while mimicking engine geometry, mainstream and coolant conditions. For this study, an endwall geometry like that found in engine with a purge jet cooling scheme is studied. Commonly, an upstream gap is formed between the combustor lining and first stage vane platform, which is accounted for in this testing. Mainstream and coolant flow conditions can have large impacts on the results gathered, so both were matched to engine conditions. Varying of coolant density and injection rate is studied and quantitative results are gathered. Results indicate coolant fluid density plays a large role in purge jet cooling, and with neglection of this, potential thermal failure points could be overlooked This is exacerbated with less coolant injection. Interestingly, increasing the amount of coolant injected decreases performance across much of the passage, with only marginal gains in regions of complex flow. These results help to better explain the impacts of experimental neglect of coolant density, and aid in the understanding of purge jet coolant injection.