An experimental study was performed to measure surface Nusselt number and film cooling effectiveness on a film cooled first stage nozzle guide vane using a transient thin film gauge (TFG) technique. The information presented attempts to further characterize the performance of shaped hole film cooling by taking measurements on a row of shaped holes downstream of leading edge showerhead injection on both the pressure and suction surfaces (hereafter PS and SS) of a 1st stage NGV. Tests were performed at engine representative Mach and Reynolds numbers and high inlet turbulence intensity and large length scale at the Virginia Tech Transonic Cascade facility. Three exit Mach/Reynolds number conditions were tested: 1.0/1,400,000; 0.85/1,150,000; and 0.60/850,000 where Reynolds number is based on exit conditions and vane chord. At Mach/Reynolds numbers of 1.0/1,450,000 and 0.85/1,150,000 three blowing ratio conditions were tested: BR = 1.0, 1.5, and 2.0. At a Mach/Reynolds number of 0.60/850,000, two blowing ratio conditions were tested: BR = 1.5 and 2.0. All tests were performed at inlet turbulence intensity of 12% and length scale normalized by leading edge diameter of 0.28. Film cooling effectiveness and heat transfer results compared well with previously published data, showing a marked effectiveness improvement (up to 2.5x) over the showerhead only NGV and agreement with published showerhead-shaped hole data. NHFR was shown to increase substantially (average 2.6x increase) with the addition of shaped holes, with only a small increase (average 1.6x increase) in required coolant mass flow. Heat transfer and effectiveness augmentation with increasing blowing ratio was shown on the pressure side, however the suction side was shown to be less sensitive to changing blowing ratio. Boundary layer transition location was shown to be within a consistent region on the suction side regardless of blowing ratio and exit Mach number. / Master of Science

Turbine airfoils are exposed to the hottest temperatures in the gas turbine with temperatures typically exceeding the melting point of the blade material. Cooling methods investigated in this computational study included parasitic cooling flow losses, which are inherent to engines, and microcircuit channels. Parasitic losses included dirt purge holes, located along the blade tip, and platform leakage flow, which result from gaps between various turbine components. Microcircuits are a novel cooling technique involving small air passages placed near the airfoil surface to enhance internal cooling. This study evaluated the benefit of external film-cooling flow exhausted from strategically placed microcircuits. Along the blade tip, predictions showed mid-chord cooling was independent of the blowing from microcircuit exits. The formation of a pressure side vortex was found to develop for most microcircuit film-cooling cases. Significant leading edge cooling was obtained from coolant exiting from dirt purge holes with a small tip gap while little cooling was seen with a large tip gap. Along the blade platform, the migration of coolant from the front leakage was shown to cool a considerable part of the platform. Several hot spots were predicted along the platform, which were circumvented through the placement of microcircuit channels. Ingestion of hot mainstream gas was predicted along the aft portion of the gutter and agreed with distress exhibited by actual gas turbine engines. / Master of Science

This study describes a detailed investigation on the effects that upstream step misalignment and upstream purge film cooling have on the endwall heat transfer for nozzle guide vanes in a land based power generation gas turbine at transonic conditions. Endwall Nusselt Number and adiabatic film cooling effectiveness distributions were experimentally calculated and compared with qualitative data gathered via oil paint flow visualization which also depicts endwall flow physics. Tests were conducted in a transonic linear cascade blowdown facility. Data were gathered at an exit Mach number of 0.85 with a freestream turbulence intensity of 16% at a Re = 1.5 x 106 based on axial chord. Varied upstream purge blowing ratios and a no blowing case were tested for 3 different upstream step geometries, one of which was the baseline (no step). The other two geometries are a backward step geometry and a forward step geometry, which comprised of a span-wise upstream step of +4.86% span and -4.86% span respectively. Experimentation shows that the addition of upstream purge film cooling increases the Nusselt Number at injection upwards of 50% but lowers it in the throat of the passage by approximately 20%. The addition of a backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow, thus aiding to help keep the film attached to the endwall at higher blowing ratios. Increasing the blowing ratio increases film cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff. Near endwall streamlines without purge cooling generated by Li et al. [1] for the same geometries were compared to the experimental data. It was shown that even with the addition of upstream purge cooling, the near endwall streamlines as they moved downstream matched strikingly well with the experimental data. This discovery indicates that while the coolant flow will likely affect the flow streamlines three dimensionally, they are minimally effected by the coolant flow near the endwall as the flow moves downstream. / Master of Science

No description available.

In this study, the cooling of a simulated blade under increasing pitch between holes was examined. The change in non-dimensional surface temperature, phi, was measured experimentally to quantify this performance loss. This critical quantification of the sensitivity of cooling to pitch between holes has not been studied previously. A range of blowing ratios and angles of attack were tested. Data are presented in terms of the laterally averaged phi, and in terms of the minimum phi, arguably more important from a design perspective. Increasing the pitch 13% produced no measureable change using either parameter. An increase of 26% in pitch produced only a 4% loss in lateral averages, while some hot points dropped by 10%. These small changes are due to compensating effects of increased internal and through-hole convective cooling. A limit to these effects was shown when increasing pitch 53%. While performance loss in the average was still relatively small at 15%, the minimum phi decreased by 27%. Heat flux gauges were used to gather data on the internal and external surface. The internal impingement used in this study represents a more accurate representation of internal cooling for an actual engine part than has been previously studied, providing a starting point for exploring the differences between engine configurations and those generally investigated in the literature. External heat flux measurements were used to measure the ratio of heat flux with and without film cooling. These results call into question the use of the net heat flux reduction parameter, which is commonly used to quantify overall film cooling performance. / text

The efficiency of natural gas turbines is directly linked to the turbine inlet temperature, or the combustor exit temperature. Further increasing the turbine inlet temperature damages the turbine components and limits their durability. Advances in turbine vane cooling schemes protect the turbine components. This thesis studies the conjugate effects of internal cooling, film cooling and thermal barrier coatings (TBC) on turbine vane metal temperatures. Two-dimensional thermal profiles were experimentally measured downstream of a single row of film cooling holes on both an adiabatic and a matched Biot number model turbine vane. The measurements were taken as a comparison to computational simulations of the same model and flow conditions. To improve computational models of the evolution of a film cooling jet as it propagates downstream, the thermal field above the vane, not just the footprint on the vane surface must be analyzed. This study expands these data to include 2-D thermal fields above the vane at 0, 5 and 10 hole diameters downstream of the film cooling holes. In each case the computational jets remained colder than the experimental jets because they did not disperse into the mainstream as quickly. Finally, in comparing results above adiabatic and matched Biot number models, these thermal field measurements allow for an accurate analysis of whether or not the adiabatic wall temperature was a reasonable estimate of the driving temperature for heat transfer. In some cases the adiabatic wall temperature did give a good indication of the driving temperature for heat transfer while in other cases it did not. Previous tests simulating the effects of TBC on an internally and film cooled model turbine vane showed that the insulating effects of TBC dominate over variations in film cooling geometry and blowing ratio. In this study overall and external effectiveness were measured using a matched Biot number model vane simulating a TBC of thickness 0.6d, where d is the film cooing hole diameter. This new model was a 35% reduction in thermal resistance from previous tests. Overall effectiveness measurements were taken for an internal cooling only configuration, as well as for three rows of showerhead holes with a single row of holes on the pressure side of the vane. This pressure side row of holes was tested both as round holes and as round holes embedded in a realistic trench with a depth of 0.6 hole diameters. Even in the case of this thinner TBC, the insulating effects dominate over film cooling. In addition, using measurements of the convective heat transfer coefficient above the vane surface, and the thermal conductivities of the vane wall and simulated TBC material, a prediction technique of the overall effectiveness with TBC was evaluated. / text

Film cooling is a cooling technique widely used in high-performance gas turbines to protect turbine airfoils from being damaged by hot flue gases. Film injection holes are placed in the body of the airfoil to allow coolant to pass from the internal cavity to the external surface. The ejection of coolant gas results in a layer or “film” of coolant gas flowing along the external surface of the airfoil. In this study, a new cooling scheme, mist/air film cooling is proposed and investigated through experiments. Small amount of tiny water droplets with an average diameter about 7 μm (mist) is injected into the cooling air to enhance the cooling performance. A wind tunnel system and test facilities were build. A Phase Doppler Particle Analyzer (PDPA) system is employed to measure droplet size, velocity and turbulence. Infrared camera and thermocouples are both used for temperature measurements. Mist film cooling performance is evaluated and compared against air-only film cooling in terms of adiabatic film cooling effectiveness and film coverage. Experimental results show that for blowing ratio M=0.6, net enhancement in adiabatic cooling effectiveness can reach 190% locally and 128% overall along the centerline. The general pattern of adiabatic cooling effectiveness distribution of the mist case is similar to that of the air-only case with the peak at about the same location. The concept of Film Decay Length (FDL) is proposed to quantitatively evaluate how well the coolant film covers the blade surface. Application of mist in the M=0.6 condition is apparently superior to the M=1.0 and 1.4 cases due to the higher overall cooling enhancement, the much longer FDL, and wider and longer film cooling coverage area. Based on droplet measurements through PDPA, a profile describing how the airmist coolant jet flow spreads and eventually blends into the hot main flow is proposed. A sketch based on the proposed profile is provided. This profile is found to be well supported by the measurement results of Turbulent Reynolds Stress. The location where a higher magnitude of Turbulent Reynolds Stress exists, which indicates higher strength of turbulent mixing effect, is found to be in the close neighborhood of the edge of the coolant film envelope. Also the separation between the mist droplets layer and the coolant air film is identified through the measurements. In other words, large droplets penetrate through the air coolant film layer and travel further over into the main flow. Based on the proposed air-mist film profile, the heat transfer results are reexamined. It is found that the location of optimum cooling effect is coincident with the starting point where the air-mist coolant starts to bend towards the surface. Thus the data suggests that the “bending back” film pattern is critical in keeping the mist droplets close to the surface which improves the cooling effectiveness for mist cooling.

In the modern advanced gas turbines, the turbine inlet temperature may exceed 1500°C as a requirement to increase power output and thermal efficiency. Therefore, it is imperative that the blades and vanes are cooled so they can withstand these extreme temperatures. Film cooling is a cooling technique widely used in high-performance gas turbines. However, the film cooling effectiveness has almost reached plateau, resulting in a bottleneck for continuous improvement of gas turbines' efficiency. In this study, an innovative cooling scheme, mist film cooling is investigated through experiments. A small amount of tiny water droplets with an average diameter about 10-15 µm (mist) is injected into the cooling air to enhance the cooling performance. A Phase Doppler Particle Analyzer (PDPA) system is used for droplet measurements. Mist film cooling performance is evaluated and compared against air-only film cooling. This study continues the previous work by (a) adding fan-shaped holes and comparing their cooling performance with the round holes, (b) extending the length of the test section to study the performance farther downstream the injection holds, and (c) using computational simulation to investigate the feasibility of transporting mist to the film cooling holes through gas turbine inside passages. The results show that, with an appropriate blowing ratio, the fan-shaped holes performs about 200% better than round holes in cooling effectiveness and adding 10% (wt.) mist can further enhance cooling effectiveness 170% in average. Farther downstream away from the injection holes (X/D> 50), mist cooling enhancement prevails and actually increases significantly. PDPA measurements have shed lights to the fundamental physics of droplet dynamics and their interactions with thermo-flow fields. These experimental results lead to either using lower amount of cooling air or use fewer number of cooing holes rows. This means higher gas turbine power output, higher thermal efficiency, and longer components life which will reflect as a cheaper electricity bill. Computational Fluid Dynamics (CFD) showed that it is feasible to transport the water mist, with initial diameters ranging from 30 µm-50 µm and mist ratio of 10-15%, to the cooling holes on the surface of the turbine vanes and rotors to provide the desired film cooling. Key words: Gas Turbines, Heat Transfer, Film / mist Cooling, Experimental Study, Mist Transport, CFD, PDPA.

Approximately 10% of all energy consumed in the United States is derived from high temperature gas turbine engines. As a result, a 1% increase in engine efficiency would yield enough energy to satisfy the demands of approximately 1 million homes and savings of over $800 million in fuel costs per year. Efficiency of gas turbine engines can be improved by increasing the combustor temperature. Modern engines now operate at temperatures that far exceed the material limitations of the metals they are comprised of in the pursuit of increased thermal efficiency. Various techniques to thermally protect the turbine components are used to allow for safe operation of the engines despite the extreme environments: film cooling, internal convective cooling, and thermal barrier coatings. Historically, these thermal protection techniques have been studied separately without account for any conjugate effects. The end goal of this work is to provide a greater understanding of how the conjugate effects might alter the predictions of thermal behavior and consequently improve engine designs to pursue increased efficiency. The primary focus of this study was to complete the first open literature, high resolution experiments of a modeled first stage turbine vane with both active film cooling and a simulated thermal barrier coating (TBC). This was accomplished by scaling the thermal behavior of a real engine component to the model vane using the matched Biot number method. Various film cooling configurations were tested on both the suction and pressure side of the model vane including: round holes, craters, traditional trenches and a novel modified trench. IR thermography and ribbon thermocouples were used to measure the surface temperature of the TBC and the temperature at the interface of the TBC and vane wall, respectively. This work found that the presence of a TBC significantly dampens the effect of altering film cooling conditions when measuring the TBC interface temperature. This work also found that in certain conditions adiabatic effectiveness does not provide an accurate assessment of how a film cooling design may perform in a real engine. An additional focus of this work was to understand how contaminant deposition alters the cooling performance of a vane with a TBC. This work focused on quantifying the detrimental effects of active deposition by seeding the mainstream flow of the test facility with simulated molten coal ash. It was found that in most cases, except for round holes operating at relatively high blowing ratios, the performance of film cooling was negatively altered by the presence of contaminant deposition. However, the cooling performance at the interface of the TBC and vane wall actually improved with deposition due to the additional thermal resistance that was added to the exterior surface of the model vane. / text

Efficiency improvement in turbomachines is an important aspect in reducing the use of fossil-based fuel and thereby reducing carbon dioxide emissions in order to achieve a sustainable future. Gas turbines are mainly fossil-based turbomachines powering aviation and land-based power plants. In line with the present situation and the vision for the future, gas turbine engines will retain their central importance in coming decades. Though the world has made significant advancements in gas turbine technology development over past few decades, there are yet many design features remaining unexplored or worth further improvement. These features might have a great potential to increase efficiency. The high pressure turbine (HPT) stage is one of the most important elements of the engine where the increased efficiency has a significant influence on the overall efficiency as downstream losses are substantially affected by the prehistory. The overall objective of the thesis is to contribute to the development of gas turbine efficiency improvements in relation to the HPT stage.   Hence, this study has been incorporated into a research project that investigates leading edge contouring near endwall by fillet and external cooling on a nozzle guide vane with a common goal to contribute to the development of the HPT stage. In the search for HPT stage efficiency gains, leading edge contouring near the endwall is one of the methods found in the published literature that showed a potential to increase the efficiency by decreasing the amount of secondary losses. However, more attention is necessary regarding the realistic use of the leading edge fillet. On the other hand, external cooling has a significant influence on the HPT stage efficiency and more attention is needed regarding the aerodynamic implication of the external cooling. Therefore, the aerodynamic influence of a leading edge fillet and external cooling, here film cooling at profile and endwall as well as TE cooling, on losses and flow field have been investigated in the present work. The keystone of this research project has been an experimental investigation of a modern nozzle guide vane using a transonic annular sector cascade. Detailed investigations of the annular sector cascade have been presented using a geometric replica of a three dimensional gas turbine nozzle guide vane. Results from this investigation have led to a number of new important findings and also confirmed some conclusions established in previous investigations to enhance the understanding of complex turbine flows and associated losses.   The experimental investigations of the leading edge contouring by fillet indicate a unique outcome which is that the leading edge fillet has no significant effect on the flow and secondary losses of the investigated nozzle guide vane. The reason why the leading edge fillet does not affect the losses is due to the use of a three-dimensional vane with an existing typical fillet over the full hub and tip profile. Findings also reveal that the complex secondary flow depends heavily on the incoming boundary layer. The investigation of the external cooling indicates that a coolant discharge leads to an increase of profile losses compared to the uncooled case. Discharges on the profile suction side and through the trailing edge slot are most prone to the increase in profile losses. Results also reveal that individual film cooling rows have a weak mutual effect. A superposition principle of these influences is followed in the midspan region. An important finding is that the discharge through the trailing edge leads to an increase in the exit flow angle in line with an increase of losses and a mixture mass flow. Results also indicate that secondary losses can be reduced by the influence of the coolant discharge. In general, the exit flow angle increases considerably in the secondary flow zone compared to the midspan zone in all cases. Regarding the cooling influence, the distinct change in exit flow angle in the area of secondary flows is not noticeable at any cooling configuration compared to the uncooled case. This interesting zone requires an additional, accurate study. The investigation of a cooled vane, using a tracer gas carbon dioxide (CO2), reveals that the upstream platform film coolant is concentrated along the suction surfaces and does not reach the pressure side of the hub surface, leaving it less protected from the hot gas. This indicates a strong interaction of the secondary flow and cooling showing that the influence of the secondary flow cannot be easily influenced.   The overall outcome enhances the understanding of complex turbine flows, loss behaviour of cooled blade, secondary flow and interaction of cooling and secondary flow and provides recommendations to the turbine designers regarding the leading edge contouring and external cooling. Additionally, this study has provided to a number of new significant results and a vast amount of data, especially on profile and secondary losses and exit flow angles, which are believed to be helpful for the gas turbine community and for the validation of analytical and numerical calculations. / Ökad verkningsgrad i turbomaskiner är en viktig del i strävan att minska användningen av fossila bränslen och därmed minska växthuseffekten för att uppnå en hållbar framtid. Gasturbinen är huvudsakligen fossilbränslebaserad, och driver luftfart samt landbaserad kraftproduktion. Enligt rådande läge och framtidsutsikter bibehåller gasturbinen denna centrala roll under kommande decennier. Trots betydande framsteg inom gasturbinteknik under de senaste årtionden finns fortfarande många designaspekter kvar att utforska och vidareutveckla. Dessa designaspekter kan ha stor potential till ökad verkningsgrad. Högtrycksturbinsteget är en av de viktigaste delarna av gasturbinen, där verkningsgraden har betydande inverkan på den totala verkningsgraden eftersom förluster kraftigt påverkas av tidigare förlopp. Huvudsyftet med denna studie är att bidra till verkningsgradsförbättringar i högtrycksturbinsteget.   Studien är del i ett forskningsprojekt som undersöker ledskenans framkantskontur vid ändväggarna samt extern kylning, i jakten på dessa förbättringar. Den aerodynamiska inverkan av en förändrad geometri vid ledskenans ändväggar har i tidigare studier visat potential för ökad verkningsgrad genom minskade sekundärförluster. Ytterligare fokus krävs dock, med användning av en rimlig hålkälsradie. Samtidigt har extern kylning i form av filmkylning stor inverkan på verkningsgraden hos högtrycksturbinsteget och forskning behövs med fokus på den aerodynamiska inverkan. Av denna anledning studeras här inverkan både av ändrad hålkälsradie vid ledskenans framkant samt extern kylning i form av filmkylning av skovel, ändvägg och bakkant på aerodynamiska förluster och strömningsfält. Huvudpelaren i detta forskningsprojekt har varit en experimentell undersökning av en geometrisk replika av en modern tredimensionell gasturbinstator i en transonisk annulärkaskad. Detaljerade undersökningar i annulärkaskaden har gett betydande resultat, och bekräftat vissa tidigare studier. Detta har lett till ökad förståelsen för de komplexa flöden och förluster som karakteriserar gasturbiner.   De experimentella undersökningarna av förändrad framkantsgeometri leder till den unika slutsatsen att den modifierade hålkälsradien inte har någon betydande inverkan på strömningsfältet eller sekundärförluster av den undersökta ledskenan. Anledningen till att förändringen inte påverkar förlusterna är i detta fall den tredimensionella karaktären hos ledskenan med en redan existerande typisk framkantsgeometri. Undersökningarna visar också att de komplexa sekundärströmningarna är kraftigt beroende av det inkommande gränsskiktet. Undersökning av extern kylning visar att kylflödet leder till en ökad profilförlust. Kylflöde på sugsidan samt bakkanten har störst inverkan på profilförlusten. Resultaten visar också att individuella filmkylningsrader har liten påverkan sinsemellan och kan behandlas genom en superpositionsprincip längs mittsnittet. En viktig slutsats är att kylflöde vid bakkanten leder till ökad utloppsvinkel tillsammans med ökade förluster och massflöde. Resultat tuder på att sekundärströmning kan minskas genom ökad kylning. Generellt ökar utloppsvinkeln markant i den sekundära flödeszonen jämfört med mittsnittet för alla undersökta fall. Den kraftiga förändringen i utloppsvinkel är dock inte märkbar i den sekundära flödeszonen i något av kylfallen jämfört med de okylda referensfallet. Denna zon fordrar ytterligare studier. Spårgasundersökning av ledskenan med koldioxid (CO2) visar att plattformskylning uppströms ledskenan koncentreras till skovelns sugsida, och når inte trycksidan som därmed lämnas mer utsatt för het gas. Detta påvisar den kraftiga interaktionen mellan sekundärströmning och kylflöden, och att inverkan från sekundärströmningen ej enkelt kan påverkas. De generella resultaten från undersökningen ökar förståelsen av komplexa turbinflöden, förlustbeteenden för kylda ledskenor, interaktionen mellan sekundärströmning och kylflöden, och ger rekommendationer för turbinkonstruktörer kring förändrad framkantsgeometri i kombination med extern kylning. Dessutom har studien gett betydande resultat och en stor mängd data, särskilt rörande profil- och sekundärförluster och utloppsvinkel, vilket tros kunna vara till stor hjälp för gasturbinssamfundet vid validering av analytiska och numeriska beräkningar. / <p>QC 20140909</p> / Turbopower, Sector rig