The resolution of differential equations of diverse degree of complexity is necessary to simulate the phenomena present in the complex turbomachinery flows and in particular, requires accounting for unsteady effects that may have a preponderant role. Today, only the LES (Large Eddy Simulation) fully compressible approach has the required accuracy to predict the physics associated to reactive and turbulent flows in such complex geometries. This work covers the numerical modelling of physics in the near-wall region of a high-pressure turbine blade with special focus on thermal predictions. This work was supported by the European project COPA-GT, dedicated to the numerical multi-physics simulation of a complete gas turbine.

This investigation studies the leakage flows over the high pressure turbine blade tip at high speed flow conditions. There is an unavoidable gap between the un-shrouded blade tip and the engine casing in a turbine stage, where the pressure difference between the pressure and the suction surfaces of the blade gives rise to the development of leakage flows through this gap. These flows contribute to about one third of the aerodynamic losses in a turbine stage. In addition they expose the blade tip to a very high temperature and result in thermal damages which reduce the blade‟s operational life. Therefore any improvement on the tip design to reduce these flows has a significant impact on the engine‟s efficiency and turbine blade‟s operational life. At the engine operational condition, the leakage flows over the high pressure turbine blade tip are mostly transonic. On the other hand literature survey has shown that most of the studies on the tip leakage flows have been performed at low speed conditions and there are only a few experimental works on the transonic tip flows. This project aims to explore the tip leakage flows at high speed condition which is the real engine condition, both experimentally and computationally and establish a comprehensive understanding of these flows on different tip geometries. The effect of tip geometry was studied using the flat tip and the cavity tip models and the effect of in-service burnout on these two tip models was established using the radius-edge flat tip and the radius-edge cavity tip models. The experimental work was carried out in the transonic wind tunnel of Queen Mary University of London and the computational simulations were performed using RANS and URANS. As the flow approached each tip model it turned and accelerated around its leading edge in the same way as the flow turns around the leading edge of an aerofoil. In the case of the tip models with sharp edges the tip flow separated at the inlet to the tip gap. For the flat tip model the flow reattachment occurred further downstream whereas in the case of the cavity tip model the length of the pressure side rim was not sufficient for the reattachment to occur and the separated flow left the rim as a free shear layer. The cavity tip model was found to have a smaller effective tip gap and hence smaller discharge coefficient in comparison to the flat tip model. For the radius-edge tip models, no separation occurred at the inlet to the tip gap and the effective tip gap was found to be the same as the geometrical tip gap. Therefore it was concluded that the tip model with radius-edges had a larger effective tip gap and hence a greater discharge coefficient than the tip geometry with sharp edges. It was observed that in the case of the supersonic tip leakage flows, decreasing the pressure ratio PR (i.e. the ratio of the static pressure at the tip gap exit to the stagnation pressure at the inlet to the tip gap) increased the discharge coefficient Cd for the tip models with sharp edges but it decreased the Cd value in the case of the tip models with radius edges. The cavity tip model with sharp edges was found to have the smallest discharge coefficient and thus the best performance in reducing the tip leakage flows as compared to all the other tip models studied in this investigation.

In modern gas turbines, with the increase of inlet gas temperature to raise thework output, the importance of accurate aero-thermal analysis has become of vitalimportance. These analysis are required for temperature prediction throughoutthe turbine and to predict the thermal stresses and to estimate the cooling requiredfor each component.In the past 20 years, computational fluid dynamics (CFD) methods have becomea powerfool tool aero-thermal analysis. Due to reasons including numericallimitation, flow complications caused by blade row interactions and the effect offilm cooling, using simple steady state CFD methods may result in inaccuratepredictions. Even though employing transient simulations can improve the accuracyof the simulations, it will also greatly increase the simulation time and cost.Therefore, new methods are constantly being developed to increase the accuracywhile keeping the computational costs relatively low. Investigating some of thesedeveloped methods is one of the main purposes of this study.A simplification that has long been applied in gas turbine simulations hasbeen the absence of cooling cavities. Another part of this thesis will focus onthe effect of cooling cavities and the importance of including them in the domain.Therefore, all transient and steady state simulations have been examined for twocases; a simplified case and a detailed case. The results are then compared tothe experimental measurements to evaluate the importance of their presence inthe model. The software used to perform all simulations is the commercial codeANSYS CFX 15.The findings suggest that even though including cooling cavities would improvethe results, the simulations should be run in transient. One important finding wasthat when performing transient simulations, especially the Time Transformationmethod, not only is the pitch ratio between every subsequent blade row important,but also the pitch ratio between the stators is highly influential on the accuracyof the results.

La résolution d’équations différentielles de divers degrés de complexité est nécessaire afin de simuler tous les phénomènes présents dans les écoulements complexes de turbomachine et en particulier les effets hors équilibre qui peuvent y jouer un rôle prépondérant. Aujourd’hui, seule l’approche LES (Large Eddy Simulation) sous forme totalement compressible permet d’obtenir avec une précision satisfaisante la physique associée aux écoulements réactifs et turbulents en géométrie complexe. Le travail porte sur la modélisation numérique et physique des échanges thermiques en proche paroi. Ce travail de thèse s'est appuyé sur le projet Européen COPA-GT dédié à la simulation numérique et multi-physique d'un moteur complet. / The resolution of differential equations of diverse degree of complexity is necessary to simulate the phenomena present in the complex turbomachinery flows and in particular, requires accounting for unsteady effects that may have a preponderant role. Today, only the LES (Large Eddy Simulation) fully compressible approach has the required accuracy to predict the physics associated to reactive and turbulent flows in such complex geometries. This work covers the numerical modelling of physics in the near-wall region of a high-pressure turbine blade with special focus on thermal predictions. This work was supported by the European project COPA-GT, dedicated to the numerical multi-physics simulation of a complete gas turbine.

<p>Improvements in modern jet engines to obtain lower specific fuel consumption and emission levels have pushed towards the optimization of every component in these machines. One current approach in minimizing engine weight is using high lift blading to reduce blade count and therefore the overall turbine weight while maintaining similar levels of efficiency. The goal of this work was to evaluate the performance of six different high pressure turbine geometries to assess the viability of using high lift designs. To do so, this study begins through the development of measurement techniques to aid the aerodynamic characterization of the flow field in a turbine environment: five-hole probe calibration maps were produced, hot wire sensors were calibrated, and an oil visualization technique was developed. The airfoils in this work were tested in a rainbow annular cascade in a stationary rig and a flow conditioning gauze was used to replicate the rotor inlet profile in the blade reference frame. The next step taken was characterizing and validating the flow field behind the gauze. Subsequently, the six turbine geometries were tested at two transonic Mach number conditions and at different Reynolds numbers. Continuous total pressure and flow angle traverses were carried out to evaluate the pressure and kinetic loss coefficients across the turbines; blade loading measurements and oil visualizations were used to quantify the power extraction of the different geometries and to study the flow structures within a turbine passage. </p>

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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

<p>To increase overall aircraft engine efficiency, the diameter of the high-pressure turbine is reduced, leading to low aspect ratio airfoils. Secondary flow dominates in these low aspect ratio turbines, and the small airfoil geometry inhibits flush-mounted, full-spatial dynamic pressure measurements with pressure transducers. Airfoil surface pressure measurements are vital to understanding the inherently unsteady flow phenomena in turbines. Additionally, aerodynamic performance data derived from high-resolution surface pressure measurements provide invaluable data for validating computational fluid dynamics codes used for prediction. Non-intrusive measurement techniques such as fast-responding Pressure Sensitive Paint (PSP) offer a potential solution of a full-field optical measurement of surface pressure fluctuation, with each camera pixel representing a sensor. The porous binder improves the dynamic response of PSP, making it suitable for unsteady flow environments such as turbomachinery applications. In this view, the overall objective of the current doctoral research is to develop a lifetime PSP method using laser-based excitation for surface pressure measurement on a new class of high-pressure turbines. </p> <p>The overall research goal was subdivided into three main strategies. (1) A pulse lifetime calibration procedure of a porous polymer/ceramic binder PSP was developed in a pressure-controlled chamber to assess the correlation between pressure and time-resolved luminescent lifetime, pressure sensitivity, and signal-to-noise ratio. (2) The lifetime technique was implemented for surface pressure measurements in a linear test section to measure high spatial pressure gradients and resolve unsteady flow features. A data reduction routine and an optimal binning bundle of pixels were proposed for calibration analysis to reduce the overall pressure uncertainty. Uncertainty quantification and sensitivity analysis were also completed to determine the parameters with a substantial effect on the pressure uncertainty. (3) The pulse lifetime method was demonstrated on a high-pressure turbine vane suction surface at engine representative conditions. The surface pressure data were corroborated with static pressure tappings and computational simulations. This research effort provided new insights into time-resolved luminescent lifetime PSP techniques. Steady and unsteady flow features from surface pressure measurements were identified using a precise calibration method. The lifetime pulse method was effective in a high-pressure turbine flow field, paving the way for back-to-back PSP experiments with different turbine geometries. </p>