A coupled density based solver in the framework of foam-extend is used to perform simulations of transonic flows. The solver is based on an explicit and time-accurate algorithm and is coupled to a compressible Unsteady Reynolds-Averaged Navier-Stokes (URANS) and a Large Eddy Simulation (LES) module. The solver is first attested on canonical compressible flow scenarios such as a 1-D shock tube and the transonic flow through a 2-D channel. Following this, a 2-D URANS simulation of the flow within the passages of a High Pressure Turbine Nozzle Guide Vane (HPT-NGV) is performed and compared against experimental data. Finally, preliminary results of a 3-D LES on a simplified geometry of the HPT-NGV are presented. In the future, this numerical setup will be used to study indirect combustion noise in aircraft engines.

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

Efficiency improvement is one of the most important aspects of engineering and especially important in the field of energy production. In the past decades, energy was mostly produced by fossil based technologies involving turbomachines, and the efficiency of these machines nearly quadrupled since the introduction of the first economically viable gas turbines. The progress continues, as there are still areas where improvement can be made. Such area is the High Pressure Turbine stage (HPT), which influences the flow characteristics and losses downstream, which this thesis will examine in more detail. In the open literature it can be found that one of the areas with potential for progress is the external cooling of the nozzle guide vanes (NGV) of the HPT stage. However not many studies go towards supersonic exit velocities even though that is the most common trend followed by the industry these days. The external cooling allows the turbine entry temperature (TET) to go beyond the melting point of the blade material thus increase Carnot efficiency but in the meantime influences the flow characteristics and losses. To understand these influences of the cooling, experiments in an annular sector cascade (ASC) were conducted with exit velocities from Mach 0.95 to 1.2 without and with cooling applied. The findings of the experiments are believed to help the more detailed understanding of the flow behaviour at high exit velocities. When comparing the corresponding runs in the two cases it became obvious that with cooling applied the deviation of the exit flow angle is generally smaller than in the uncooled case. This might be a highly important design feature for designers to work with. From the available data it was concluded that the total pressure distribution across the span is not significantly affected with the introduction of cooling.

Over the last two decades, most developed countries have reached a consensus that greener energy production is necessary for the world, due to the climate changes and limited fossil fuel resources. More efficient turbine is desirable and can be archived by higher turbine-inlet temperature (TIT). However, it is difficult for nozzle guide vane (NGV), which is the first stage after combustion chamber, to withstand a very high temperature. Thus, cooling methods such as film cooling have to be implemented. Film-cooled NGV of an annular sector cascade (ASC) is studied in this thesis, for getting comprehensive calculation of vorticity, and analyzing applicability of existing loss models, namely Hartsel model and Young &amp; Wilcock model. The flow-field calculation methods from previously published studies are reviewed. Literatures focusing on Hartsel model and Young &amp; Wilcock model are studied. Measurement data from previously published studies are analyzed and compared with the loss models. In order to get experience of how measurements take place, participation of a test run experiment is involved. Calculation of flow vector has been evaluated and modified. Actual flow angle is introduced when calculating velocity components. Thus, more exact results are obtained from the new method. Calculation of vorticity has been evaluated and made more comprehensive. Vorticity components as well as magnitude of total streamwise vorticity are calculated and visualized. Vorticity is higher and more extensive for fully cooled case than uncooled case. Highest vorticity is found at regions near the hub, tip and TE. Axial and circumferential vorticities show similar patterns, while the radial vorticity is relatively simpler. Compressibility is introduced as a new method when calculating circumferential and radial vorticities, resulting more extensive and higher vorticities than results from incompressible solutions. Hartsel model and Young &amp; Wilcock model have been evaluated and compared to the ASC to see the applicability of the models. In general, Hartsel model cannot agree with the ASC to a satisfactory level and thus cannot be applied. Coolant velocity is found to be the dominant factor of Hartsel model. Young &amp; Wilcock model may match SS1 and SS2 cases, or even PS and SH4 cases, but cannot match TE case. The applicability of Young &amp; Wilcock model is much dependent on the location of cooling rows.

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Modern gas turbine designs' ever-increasing turbine inlet temperature raises challenges for the nozzle guide vane cooling. Two typical endwall cooling schemes, jump cooling and louver cooling, result in different interactions between the injected coolant and the mainstream, leading to different cooling effects. This study investigates these two cooling schemes on the endwall cooling experimentally and numerically. Wind tunnel tests and the CFD simulations are carried out with engine-representative conditions of an exit Mach number of 0.85, an exit Reynolds number of 1.5×10^6, and an inlet Turbulence intensity of 16%. The jump cooling scheme experiments investigate two blowing ratios, 2.5 and 3.5, two density ratios, 1.2 and 1.95, and three endwall profiles with different NGV-turbine alignments. Four coolant mass flow ratios from 1.0% to 4.0% are tested for the louver cooling. The results show that the cavity vortex, the horseshoe vortex, and the passage vortex are the main factors that prevent the upstream coolant from reaching the NGV passage. The jump cooling scheme generally provides high momentum to the cooling jets. As a result, the coolant at the design case density ratio of 1.95 and blowing ratio of 2.5 is sufficiently energized to penetrate the horseshoe vortex. It then forms a relatively uniform coolant film near the NGV passage inlet, leading to a minimum adiabatic cooling effectiveness of 0.4 throughout the passage. Reducing the coolant density or increasing the blowing ratio leads to higher coolant momentum, so the coolant jets can further suppress the horseshoe vortex. However, high momentum may cause coolant lift-off, mitigating the coolant reattachment. Therefore, the density ratio needs to be carefully balanced with the blowing ratio to optimize the cooling effect. This balance is also affected by the combustor-NGV misalignment, as a higher step height requires higher coolant momentum to overcome the step-induced vortices. On the contrary, the louver cooling scheme provides less momentum to the coolant. The results showed that only by exceeding a coolant mass flow rate of 1~2% can the coolant form a uniform film which provides good coverage upstream of the NGV passage inlet. As for the cooling of the NGV passage, the mass flow rate ratio of the range investigated is not sufficient for desirable cooling performance. The pressure side endwall proves most difficult for the coolant to reach. In addition, the fishmouth cavity at the combustor-NGV passage causes a three-dimensional cavity vortex that transports the coolant in the pitch-wise direction. Moreover, the coolant transport pattern is dependent on the coolant blow rate. Overall, the more-energized coolant film generated by the jump cooling tends to survive longer, but it is also more prone to lift-off. At the same time, the less-energized coolant film caused by the louver cooling is more susceptible to vortices and the discontinuity of the endwall geometry. However, it develops faster, especially in the lateral direction. The two schemes could be applied simultaneously for an ideal cooling system. The jump cooling can provide enough momentum for the coolant to persist in the NGV passage. Meanwhile, the louver cooling covers the upstream region before the jump cooling coolant reattaches to the endwall. / Doctor of Philosophy / Gas turbines, sometimes called combustion turbines, are widely used to generate power or propulsion for various applications. The three main components of a gas turbine are compressor, combustor, and turbine. Modern gas turbines run at a high turbine inlet temperature that exceeds the current metal limits to increase efficiency. However, this brings significant challenges to the cooling of the first stage of the turbine, the nozzle guide vane. In this research, two commonly used endwall cooling methods, jump cooling and louver cooling, are investigated under engine-representative conditions experimentally and numerically. In addition, flow physics is demonstrated to explain the endwall cooling performance, mainly the upstream boundary layer caused by the interaction between the mainstream and the coolant flow. The results show that the cavity vortex, the horseshoe vortex, and the passage vortex are the main factors that prevent the upstream coolant from reaching the NGV passage. The jump cooling scheme provides high momentum to the cooling jets. As a result, the coolant in the design case is sufficiently energized to penetrate the horseshoe vortex, providing a desirable cooling effect in the NGV passage. Reducing the density ratio or increasing the blowing ratio can help the coolant jets further suppress the horseshoe vortex but also causes more lift-off, which adversely affects the cooling performance. On the contrary, the louver cooling scheme provides less momentum to the coolant, forming a less energized coolant film. The lack of coolant causes the louver coolant film to provide good coverage immediately downstream of the louver scheme exit. However, due to unfavorable interaction with vortices and endwall discontinuity, the cooling effect decays quickly downstream. Overall, the more-energized coolant film generated by the jump cooling tends to survive longer, but it is also more prone to lift-off. At the same time, the less-energized coolant film caused by the louver cooling is more susceptible to vortices and the discontinuity of the endwall geometry. However, it develops faster, especially in the lateral direction. The two schemes could be applied simultaneously for an ideal cooling system to work mutually beneficially.

The primary requirements for a modern industrial gas turbine consist of a continuous trend of an increasing efficiency combined with very low emissions in a robust, cost-effective manner. To fulfil these tasks a high turbine inlet temperature together with advanced dry low NOX combustion chambers are employed. These dry low NOX combustion chambers generate a rather flat temperature profile compared to previous generation gas turbines, which have a rather parabolic temperature profile before the nozzle guide vane. This means that the nozzle guide vane endwall heat load for modern gas turbines is much higher compared to previous generation gas turbines. Therefore the prediction of the nozzle guide vane flow field and endwall heat transfer is crucial for the engineering task of the design layout of the vane endwall cooling system. The present study is directed towards establishing new in-depth aerodynamic and endwall heat transfer knowledge for an advanced nozzle guide vane of a modern industrial gas turbine. To reach this objective the physical processes and effects which cause the different flow fields and the endwall heat transfer pattern in a baseline configuration, a combustion chamber variant, a heat shield variant without and with additional cooling air and a cavity variant without and with additional cooling air have been investigated. The variants, which differ from the simplified baseline configuration, apply design elements which are commonly used in real modern gas turbines. This research area is crucial for the nozzle guide vane endwall heat transfer, especially for the advanced design of the nozzle guide vane of a modern industrial gas turbine and has so far hardly been investigated in the open literature. For the experimental aerodynamic and endwall heat transfer research of the baseline configuration of the advanced nozzle guide vane geometry a new low pressure, low temperature test facility has been developed, designed and constructed, since no experimental heat transfer data exist in the open literature for this type of vane configuration. The new test rig consists of a linear cascade with the baseline configuration of the advanced nozzle guide vane geometry with four upscaled airfoils and three flow passages. For the aerodynamic tests the two middle airfoils and the hub and the tip endwall are instrumented with pressure taps to monitor the Mach number distribution. For the heat transfer tests the temperature distribution on the hub endwall is measured via thermography. The analysis of these measurements, including comparisons to research in the open literature shows that the new test rig generates accurate and reproducible results which give confidence that it is a reliable tool for the experimental aerodynamic and heat transfer research on the advanced nozzle guide vane of a modern industrial gas turbine. Previous own research work together with the numerical analysis performed in another part of the project as well as conclusions from a detailed literature study lead to the conclusion that advanced Navier-Stokes CFD tools with the v2-f turbulence model are most suitable for the calculation of the flow field and the endwall heat transfer of turbine vanes and blades. Therefore this numerical tool, validated against different vane and blade geometries and for different flow conditions, has been chosen for the numerical aerodynamic and endwall heat transfer research of the advanced nozzle guide vane of a modern industrial gas turbine. The evaluation of the numerical and experimental investigations of the baseline configuration of the advanced design of a nozzle guide vane shows the flow field of an advanced mid-loaded airfoil design with the features to reduce total airfoil losses. For the hub endwall of the baseline configuration of the advanced design of a nozzle guide vane the flow characteristics and heat transfer features of the classical vane endwall secondary flow model can be detected with a very weak intensity and geometric extension compared to the studies of less advanced vane geometries in the open literature. A detailed analysis of the numerical simulations and the experimental data showed very good qualitative and quantitative agreement for the three-dimensional flow field and the endwall heat transfer. These findings, together with the evaluations obtained from the open literature, lead to the conclusions that selected CFD software Fluent together with the applied v2-f turbulence model exhibits a high level of general applicability and is not tuned to a special vane or blade geometry. Therefore the CFD code Fluent with the v2-f turbulence model has been selected for the research of the influence of the several geometric variants of the baseline configuration on the flow field and the hub endwall heat transfer of the advanced nozzle guide vane of a modern industrial gas turbine. Most of the vane endwall heat transfer research in the open literature has been carried out only for baseline configurations of the flow path between combustion chamber and nozzle guide vane. Such a simplified geometry consists of a long, planar undisturbed approach length upstream of the nozzle guide vane. The design of real modern industrial gas turbines however requires often significant variations from this baseline configuration consisting of air-cooled heat shields and purged cavities between the combustion chamber and the nozzle guide vane. A detailed evaluation of the flow field and the endwall heat transfer shows major differences between the baseline and the heat shield configuration. The heat shield in front of the airfoil of the nozzle guide vane influences the secondary flow field and the endwall heat transfer pattern strongly. Additional cooling air, released under the heat shield has a distinctive influence as well. Also the cavity between the combustion chamber and the nozzle guide vane affects the secondary flow field and the endwall heat transfer pattern. Here the influence of additional cavity cooling air is more decisive. The results of the detailed studies of the geometric variants are applied to formulate guidelines for an optimized design of the flow path between the combustion chamber and the nozzle guide vane and the nozzle guide vane endwall cooling configuration of next-generation industrial gas turbines. / QC 20100917

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