The modern trend of gas turbines design is towards lighter, highly efficient,and more compact engines. Such situation imposes on engineers to continuouslysearch for improved and optimum designs. The thesis presented aims at researching possible performance improvements regarding axial gapdistance in transonic compressors. Decreasing the axial gap would result inlighter engines and achieve design goals. The influence of decreasing the axialgap on performance and structure integrity should be throughly analyzed. This thesis work includes numerical investigations on the axial gap distance effect on performance efficiency and related unsteady aerodynamics phenomena. The first one and a half compressor stages of a Siemens Gas Turbine are modeled in ANSYS CFX. Different axial gap models are simulated for differentconfigurations. The steady state solution is obtained to be initialized for transient time marching calculations. Furthermore, the computational cost of transient calculations is reduced through a geometry scaling technique. The unsteady behavior is further analyzed by a Harmonic Balance solver implemented in STAR-CCM+ software, and compared to a reference case transient calculations. The results obtained supports the presence of an optimalaxial gap distance for maximum efficiency in transonic compressors. Further, the harmonic balance method shows good possibilities for cost and time reductions in transonic compressors performance calculations.

<p>Highly loaded aerodynamic devices for propulsion and power generation are emerging to increase power output in a more compact machine are emerging. These devices can experience increased losses due to separation, as in the low-pressure turbine, which arise due to the operation at conditions that increases the adverse pressure gradients ore decrease the Reynolds number of the flow through the device. Therefore, flow control strategies become appealing to reduce losses at these conditions. This work aims to validate flow injection as an effective flow control strategy in the transonic regime.</p> <p>A test facility which was used to study boundary layer separation in a quasi-2d test article was modified to include flow injection and conditions were modified so that the facility was operated in the transonic regime. Valves were chosen which could achieve a wide range of excitation frequencies and the flow control ports were designed to accommodate their nominal flow rate. A preliminary test matrix was built while considering the limitations of the test facility.</p> <p>A numerical study was conducted to identify flow structures of interest and determine a preliminary understanding of the test article. The flow control was then added to the numerical study to guide the experimental set points for injected flow. The response of the flow to continuous slot blowing was characterized, and a 3D simulation with discrete injection ports was done to ensure the set-points determined from the 2D study were viable for discrete injection.</p> <p>Blow-down experiments were then conducted to study the behavior of bulk separation in a transonic regime for a quasi-2D geometry. Once behavior of the separation was understood, steady injection and then pulsated injection were applied in attempts to mitigate the separation. Steady injection was utilized to find the required pressure of injection relative to the total pressure at the inlet of the test article, while the pulsated injection served to identify a frequency at which the time averaged mitigation of separation was greatest.</p> <p>The experiments show that both steady and pulsated flow injection are viable techniques in flow control. It is also shown that pulsation does not allow for a lower pressure injection, but instead allows for the same effect with a lower mass flow requirement. Two-dimensional computational simulations are shown to be effective in determining injection frequencies but not the extent of separation or required injection pressures.</p>

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No description available.

A hazardous and not fully understood aeroelastic phenomenon is the transonic dip,the decrease in flutter dynamic pressure that occurs for most aircraft configurationsin transonic flows. The difficulty of predicting this phenomenon forces aircraft manufacturersto run long and costly flight test campaigns to demonstrate flutter-free behaviourof their aircraft at transonic Mach numbers.In this project, subsonic and transonic flutter calculations for the KTH-NASA genericfighter research model have been performed and compared to existing experimentalflutter data from wind tunnel tests performed at NASA Langley in 2016. For the fluttercalculations, industry-standard linear panel methods have been used together with afinite element model from NASTRAN.Further, an alternative approach for more accurate transonic flutter predictions usingthe full-potential solver Phi has been investigated. To predict flutter using this newmethodology a simplified structural model has been used together with aerodynamicmeshes of the main wing. The purpose of the approach was to see if it was possibleto find a method that was more accurate than panel methods in the transonic regimewhilst still being suitable for use during iterative design processes.The results of this project demonstrated that industry-standard linear panel methodssignificantly over-predict the flutter boundary in the transonic regime. It was alsoseen that the flutter predictions using Phi showed potential, being close to the linearresults for the same configuration as tested in Phi. For improved transonic accuracy inPhi, an improved transonic flow finite element formulation could possibly help .Another challenge with Phi is the requirement of an explicit wake from all liftingsurfaces in the aerodynamic mesh. Therefore, a method for meshing external storeswith blunt trailing edges needs to be developed. One concept suggested in this projectis to model external stores in "2.5D", representing external stores using airfoils withsharp trailing edges. / Ett farligt och inte helt utrett aeroelastiskt fenomen är den transoniska dippen, minskningeni dynamiska trycket vid fladder som inträffar för de flesta flygplan i transoniskaflöden. Svårigheten i att prediktera detta fenomen tvingar flygplanstillverkare attbedriva tidskrävande och kostsam flygprovsverksamhet för att demonstrera att derasflygplan ej uppvisar fladderbeteende i transonik inom det tilltänkta användningsområdet.I detta projekt har fladderberäkningar genomförts i både underljud och transonikför en generisk stridsflygplansmodell i skala 1:4 ämnad för forskning, byggd som ettsamarbete mellan KTH och NASA. Beräkningarna har också jämförts med fladderresultatfrån vindtunnelprov genomförda vid NASA Langley under sommaren 2016. Förfladderberäkningarna har industri-standarden linjära panelmetoder används tillsammansmed en befintlig finit element modell för användning i NASTRAN.Vidare har ett alternativt tillvägagångssätt för att förbättra precisionen i transoniskafladderresultat genom att använda potentiallösaren Phi undersökts. En förenkladstrukturmodell har använts tillsammans med aerodynamiska nät av huvudvingen föratt prediktera fladder. Syftet med denna metodik var att undersöka om det var möjligtatt hitta en metod som i transoniska flöden var mer exakt än panelmetoder men somfortfarande kunde användas i iterativa design processer.Resultaten från detta projekt visade att linjära panelmetoder, som de som används iindustrin, är signifikant icke-konservativa gällande fladdergränsen i transonik. Resultatenfrån Phi visade potential genom att vara nära de linjära resultaten som räknadesfram med hjälp av panelmetoder för samma konfiguration som i Phi. För ökad transonisknoggrannhet i Phi kan möjligen en förbättrad transonisk element-formuleringhjälpa.En annan utmaning med Phi är kravet på en explicit vak från alla bärande ytor idet aerodynamiska nätet. Därför behöver det utvecklas en metodik för nätgenereringav yttre laster med trubbiga bakkanter. Ett koncept som föreslås i denna rapport är attmodellera yttre laster i "2.5D", där alla yttre laster beskrivs genom att använda vingprofilermed skarpa bakkanter.

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This dissertation presents a detailed experimental and numerical analysis of the aerothermal characteristics of the turbine extremity regions i.e. the blade tip and endwall regions. The heat transfer and secondary flow characteristics were analyzed for different engine relevant configurations and exit Mach/Reynolds number conditions. The experiments were conducted in a linear blowdown cascade at transonic high turbulence conditions of Mexit ~ 0.85, 0.60 and 1.0, with an inlet turbulence intensity of 16% and 12% for the vane and blade cascade respectively. Transient infrared (IR) thermography technique and surface pressure measurement were used to map out the surface heat transfer coefficient and aerodynamic characteristics. The experiments were complemented with computational modeling using the commercial RANS equation solver ANSYS Fluent. The CFD results provided further insight into the local flow characteristics in order to elucidate the flow physics which govern the measured heat transfer characteristics. The results reveal that the highest heat transfer exists in regions with local flow reattachment and new-boundary layer formation. Conversely, the lowest heat transfer occurs in regions with boundary layer thickening and separation/lift-off flow. However, boundary layer separation results in additional secondary flow vortices, such as the squealer cavity vortices and endwall auxiliary vortex system, which significantly increase the stage aerodynamic losses. Furthermore, these vortices result in a low film-cooling effectiveness as was observed on a squealer tip cavity with purge flow. Finally, the importance of transonic experiments in analyzing the turbine section heat transfer and flow characteristics was underlined by the significant shock-boundary layer interactions that occur at high exit Mach number conditions. / Ph. D.

Correcting wind tunnel data for wall interference is a critical part of relating the acquired data to a free-air condition. Accurately determining and correcting for the interference caused by the presence of boundaries in wind tunnels can be difficult especially for facilities employing ventilated boundaries. In this work, three varying levels of ventilation at the National Transonic Facility (NTF) were modeled and calibrated with a general slotted wall (GSW) linear boundary condition to validate the computational model used to determine wall interference corrections. Free-air lift, drag, and pitching moment coefficient predictions were compared for a range of lift production and Mach conditions to determine the uncertainty in the corrections process and the expected domain of applicability. Exploiting a previously designed statistical validation method, this effort accomplishes the extension of a calibration and validation for a boundary pressure wall interference corrections method. The foundational calibration and validation work was based on blockage interference only, while this present work extends the assessment of the method to encompass blockage and lift interference production. The validation method involves the establishment of independent cases that are then compared to rigorously determine the degree to which the correction method can converge free-air solutions for differing interference fields. The process involved first establishing an empty-tunnel calibration to gain both a centerline Mach profile of the facility at various ventilation settings, and to gain a baseline wall pressure signature undisturbed by a test article. The wall boundary condition parameters were then calibrated with a blockage and lift interference producing test article, and final corrected performance coefficients were compared for varying test section ventilated configurations to validate the corrections process and assess its domain of applicability. During the validation process discrimination between homogeneous and discrete implementations of the boundary condition was accomplished and final results indicated comparative strength in the discrete implementation's ability to capture experimental flow physics. Final results indicate that a discrete implementation of the General Slotted Wall boundary condition is effective in significantly reducing variations caused by differing interference fields. Corrections performed with the discrete implementation of the boundary condition collapse differing measurements of lift coefficient to within 0.0027, drag coefficient to within 0.0002, and pitching moment coefficient to within 0.0020. / Doctor of Philosophy / The purpose of conducting experimental tests in wind tunnels is often to acquire a quantitative measure of test article aerodynamic characteristics in such a way that those specific characteristics can be accurately translated into performance characteristics of the real vehicle that the test article intends to simulate. The difficulty in accurately simulating the real flow problem may not be readily apparent, but scientists and engineers have been working to improve this desired equivalence for the better part of the last half-century. The primary aspects of experimental aerodynamics simulation that present difficulty in attaining equivalence are: geometric fidelity, accurate scaling, and accounting for the presence of walls. The problem of scaling has been largely addressed by adequately matching conditions of similarity like compressibility (Mach number), and viscous effects (Reynolds number). However, accounting for the presence of walls in the experimental setup has presented ongoing challenges for ventilated boundaries; these challenges include difficulties in the correction process, but also extend into the determination of correction uncertainties. Exploiting a previously designed statistical validation method, this effort accomplishes the extension of a calibration and validation effort for a boundary pressure wall interference corrections method. The foundational calibration and validation work was based on blockage interference only, while this present work extends the assessment of the method to encompass blockage and lift interference production. The validation method involves the establishment of independent cases that are then compared to rigorously determine the degree to with the correction method can converge free-air solutions for differing interference scenarios. The process involved first establishing an empty-tunnel calibration to gain both a centerline Mach profile of the facility at various ventilation settings, and to gain a baseline wall pressure signature undisturbed by a test article. The wall boundary condition parameters were then calibrated with a blockage and lift interference producing test article, and final corrected performance coefficients were compared for varying test section ventilated configurations to validate the corrections process and assess its domain of applicability. During the validation process discrimination between homogeneous and discrete implementations of the boundary condition was accomplished and final results indicated comparative strength in the discrete implementation's ability to capture experimental flow physics. Final results indicate that a discrete implementation of the General Slotted Wall boundary condition is effective in significantly reducing variations caused by differing interference fields. Corrections performed with the discrete implementation of the boundary condition collapse differing measurements of lift coefficient to within 0.0027, drag coefficient to within 0.0002, and pitching moment coefficient to within 0.0020.

Les écoulements transsoniques rencontrés dans le cadre de la propulsion aéronautique et spatiale sont associés à l'apparition d'ondes de choc. En impactant la couche limite se développant sur une paroi, un gradient de pression adverse est généré qui conduit à l'épaississement ou au décollement de la couche limite. Lors de la vibration de la structure, l'onde de choc oscille et interagit avec la couche limite, générant une fluctuation de la pression statique à la paroi. Il s'ensuit alors un échange d'énergie entre le fluide et la structure qui peut être stabilisant ou au contraire conduire à une instabilité aéroélastique (flottement). La modélisation de la réponse instationnaire de l'interaction onde de choc / couche limite pour l'étude des interactions fluide-structure est l'objet de ce travail de recherche. Il s'appuie sur la résolution des équations de Navier-Stokes moyennées (RANS) et la modélisation de la turbulence. Les méthodes et modèles utilisés ont été validés à partir de résultats expérimentaux issus d'une tuyère transsonique dédiée à l'étude des interactions fluide-structure. Ces travaux sont ensuite appliqués à l'amélioration de la prédiction du flottement en turbomachine. Une méthode linéarisée en temps permettant la résolution des équations RANS dans le domaine fréquentiel est utilisée. Nous confirmons l'importance de la dérivation du modèle de turbulence lors de la prédiction d'une interaction forte entre une onde de choc et une couche limite décollée. Une méthode de régularisation est présentée puis appliquée aux opérateurs non dérivables du modèle de turbulence k-! de Wilcox (2006). La prédiction de la réponse instationnaire de l'interaction onde de choc / couche limite dans une tuyère est évaluée à partir de simulations bidimensionnelles et présente un bon accord avec les données expérimentales. En évaluant l'influence de la fréquence réduite, une instabilité aéroélastique de type flottement transsonique est identifiée. Un dispositif de contrôle, reposant sur la génération d'ondes de pression rétrogrades à l'aval de la tuyère, est proposé puis validé numériquement. Enfin, une méthodologie est proposée pour comprendre les mécanismes aérodynamiques conduisant au flottement. Pour cela, il a été réalisé un dessin provisoire d'une soufflante transsonique à fort taux de dilution. Cette soufflante, l'ECL5, est destinée à l'étude expérimentale des instabilités aérodynamiques et aéroélastiques. La méthodologie proposée repose sur la simulation 2D d'une coupe de tête et met à profit la linéarisation pour analyser la contribution de sources locales en fonction de la fréquence réduite, du diamètre nodal et de la déformée modale / Transonic flows, which are common in aeronautical and spatial propulsion systems, produce shock-waves over solid boundaries. When a shock-wave impacts the boundary layer, an adverse pressure gradient is generated and a thickening or even a separation of the boundary layer is induced. If the solid boundary vibrates, the shock-wave oscillates, interacts with the boundary layer and produce a fluctuation of the static pressure at the wall. This induces an exchange of energy between the fluid and the structure which can be stabilising or lead to an aeroelastic instability (flutter).The main objective of this PhD thesis is the modelling of the unsteady behaviour the simulation of the shock-wave/boundary layer interaction for fluid-structure interaction studies. To this end, simulations have been carried out to solve Reynolds-Averaged Navier-Stokes equations using two equations turbulence model. The method is validated thanks to experimental data obtained on a transonic nozzle dedicated to aeroelastic studies. This method is then use to increase the predictability of flutter events in turbomachinery.A time linearised frequency-domain method is applied to RANS equations. It is shown that the unsteady behaviour of the turbulent boundary-layer contributes to the fluctuating static pressure when the shock-wave boundary layer interaction is strong. Hence, the frozen turbulence assumption is not valid and the turbulence model must be derivated. Thus, the regularisation of the non derivable operators is proposed and applied on k-? Wilcox (2006) turbulence model.The unsteady behaviour of the shock-wave/boundary-layer interaction in a transonic nozzle is evaluated thanks to 2D numerical simulations and shows good agreement with experimental data. When varying the reduced frequency an aeroelastic instability is found, known as transonic flutter. An active control device generating backward travelling pressure waves is then designed and numerically validated.Finally, a methodology is proposed to understand the aerodynamic onsets of transonic flutter. To this end, a preliminary design of a high bypass ratio transonic fan has been carried out. This fan, named ECL5, is dedicated to experimental aerodynamic and aeroelastic studies. The methodology relies on 2D simulations of a tip blade passage and uses linearisation to analyse the contribution of local sources as a function of reduced frequency, nodal diameter and mode shape