Significant effort in the current technological development of aircraft is aimed at improving engine efficiency, while reducing fuel burn, emissions, and noise levels. One way to achieve these is to better integrate airframe and propulsion system. Tighter integration, however, may also cause adverse effects to the flow entering the engines, such as total pressure, total temperature, and swirl distortions. Swirl distortions are angular non-uniformities in the flow that may alter the functioning of specific components of the turbomachinery systems. To investigate the physics involved in the ingestion of swirl, pre-determined swirl distortion profiles were generated through the StreamVane method in a low-speed wind tunnel and in a full-scale turbofan research engine. Stereoscopic particle image velocimetry (PIV) was used to collect three-component velocity fields at discrete planes downstream of the generation of the distortions with two main objectives in mind: identifying the physics behind the axial development of the distorted flow; and describing the generation of the distortion by the StreamVane and its impact to the flow as a distortion generating device. Analyses of the mean velocity, velocity gradients, and Reynolds stress tensor components in these flows provided significant insight into the driving physics. Comparisons between small-scale and full-scale results showed that swirl distortions are Mach number independent in the subsonic regime. Reynolds number independence was also verified for the studied cases. The mean secondary flow and flow angle profiles demonstrated that the axial development of swirl distortions is highly driven by two-dimensional vortex dynamics, when the flow is isolated from fan effects. As the engine fan is approached, the vortices are axially stretched and stabilized by the acceleration of the flow. The flow is highly turbulent immediately downstream of the StreamVane due to the presence of the device, but that vane-induced turbulence mixes with axial distance, so that the device effects are attenuated for distances greater than a diameter downstream, which is further confirmed by the turbulent length scales of the flow. These results provide valuable insight into the generation and development of swirl distortion for ground-testing environments, and establishes PIV as a robust tool for engine inlet investigations. / Ph. D.

This thesis presents a joint computational and experimental investigation into the heat transfer to unshrouded turbine blade tips suitable for use in high bypass ratio, large civil aviation turbofan engines. Both the heat transfer to the blade tip and the over-tip leakage flow over the blade tip are characterised, as each has a profound influence on overall engine efficiency. The study is divided into two sections; in the first, computational simulations of a very large scale, low speed linear cascade with a flat blade tip were conducted. These simulations were validated against experimental data collected by Palafox (2006). A thorough assessment of turbulence models and minimum meshing requirements was performed. The standard k-ω and standard k-ϵ turbulence models significantly overpredicted the turbulence levels within the tip gap. The other models were very similar in performance; the SST k-ω and realisable k-ϵ models were found to be the most suitable for the flow environment. The second section documents the development and testing of a novel hybrid blade tip design, the squealet tip, which seeks to combine the known benefits of winglet and double squealer tips. The development of the external geometry was performed primarily through engine-representative CFD simulations at a range of tip gaps from 0.45% to 1.34% blade chord. The squealet tip was found to have a similar aerodynamic sensitivity to tip clearance as a baseline double squealer tip, with a tip gap efficiency exchange rate of 2.03, although this was 18% greater than the alternative winglet tip. The squealet tip displayed higher predicted stage efficiency than the winglet tip over the majority of the range of tip clearances investigated, however. The overall heat load was reduced by 14% compared with the winglet tip but increased by 28% over the double squealer tip, primarily due to the change in wetted surface area. The predicted local heat transfer coefficients were similar across all geometries. A realistic internal cooling plenum and an array of blade tip cooling holes were subsequently added to the squealet tip geometry and the cooling configuration refined by the selective sealing of cooling holes. Film cooling performance was largely assessed by the predicted adiabatic wall temperature distributions. A viable cooling scheme which reduced the cooling air requirement by 38% was achieved, compared to the initial case which had all cooling holes open. This was associated with just a 7% increase in blade tip heat flux and no penalty in peak temperature on the blade tip. Film cooling air ejected from holes on the blade suction side was swept away from the blade tip region, making the squealet rim at the crown of the blade particularly challenging to cool. It was demonstrated that this region could be cooled effectively by ballistic cooling from holes located on the blade tip cavity floor, although this was expensive in terms of the mass flow rate of cooling air required. The computational results were reinforced with experimental data collected in a transonic linear cascade. Downstream aerodynamic loss measurements were taken for a linearised version of the squealet tip design without cooling at nominal tip gaps of 0.45%, 0.89% and 1.34% blade chord, which was compared to similar data taken by O’Dowd (2010) for flat and winglet tips. The squealet was seen to have a similar aerodynamic loss to the flat tip and a reduced loss compared with the winglet tip. Full surface heat transfer measurements were taken for the uncooled squealet tip, at tip gaps of 0.89% and 1.34% blade chord, and for two configurations of the cooled squealet tip, at a tip clearance of 0.89% blade chord. The qualitative similarity between the measured heat transfer distributions and the those predicted by the engine-representative CFD simulations was good. A CFD simulation of the uncooled linear cascade environment at the 1.34% blade chord tip clearance was performed using a single blade with translationally periodic boundary conditions. The predicted size of the over-tip leakage vortex was smaller than had been measured, resulting in a large underprediction in the magnitude of the downstream area-averaged aerodynamic loss. The magnitudes of the predicted blade tip Nusselt number distribution were similar to those produced by the engine-representative CFD simulations and lower than that measured experimentally. Differences in the shape of the Nusselt number distribution were observed in the vicinity of regions of separated and reattaching flow, but other salient features were replicated in the computational data. The squealet tip has been shown to be a promising, viable unshrouded blade tip design with an aerodynamic performance similar to the double squealer tip but is more amenable to film cooling. It is significantly lighter than a winglet tip and incurs a reduced thermal load. The squealet tip design can now be developed into a blade tip geometry for use in real engines to provide an alternative to shrouded turbine blades and current unshrouded blade tip designs. A commercial CFD solver, Fluent 6.3, was shown to capture blade tip heat transfer and over-tip leakage flow sufficiently well to be a useful design guide. However, the sensitivity of the flow structure (and hence, heat transfer) in the forward part of the blade tip cavity suggests that physical testing cannot be eliminated from the design process entirely.

Due to the requirements for high blade loading, modern turbo‐machine blades operate very close to the stall regime. This can lead to flow separation with periodic shedding of vortices, which could lead to self induced oscillations or stall flutter of the blades. Previous studies on stall flutter have focused on flows at high Reynolds number (Re ~ 106). The Reynolds numbers for fans/propellers of Micro Aerial Vehicles (MAVs), high altitude turbofans and small wind turbines are substantially lower (Re < 105). Aerodynamic characteristics of flows at such low Re is significantly different from those at high Re, due in part to the early separation of the flow and possible formation of laminar separation bubbles (LSB). The present study is targeted towards study of stall flutter in a cascade of blades at low Re. We experimentally study stall flutter of a cascade of symmetric NACA 0012 blades at low Reynolds number (Re ~ 30, 000) through forced sinusoidal pitching of the blades about mean angles of incidences close to stall. The experimental arrangement permits variations of the inter‐blade phase (σ) in addition to the oscillation frequency (f) and amplitude; the inter‐blade phase angle (σ) being the phase difference between the motions of adjacent blades in the cascade. The unsteady moments on the central blade in the cascade are directly measured, and used to calculate the energy transfer from the flow to the blade. This energy transfer is used to predict the propensity of the blades to undergo self‐induced oscillations or stall flutter. Experiments are also conducted on an isolated blade in addition to the cascade. A variety of parameters can influence stall flutter in a cascade, namely the oscillation frequency (f), the mean angle of incidence, and the inter‐blade phase angle (σ). The measurements show that there exists a range of reduced frequencies, k (=πfc/U, c being the chord length of the blade and U being the free stream velocity), where the energy transfer from the flow to the blade is positive, which indicates that the flow can excite the blade. Above and below this range, the energy transfer is negative indicating that blade excitations, if any, will get damped. This range of excitation is found to depend upon the mean angle of incidence, with shifts towards higher values of k as the mean angle of incidence increases. An important parameter for cascades, which is absent in the isolated blade case is the inter‐blade phase angle (σ). An excitation regime is observed only for σ values between ‐450 and 900, with the value of excitation being maximum for σ of 900. Time traces of the measured moment were found to be non‐sinusoidal in the excitation regime, whereas they appear to be sinusoidal in the damping regime. Stall flutter in a cascade has differences when compared with an isolated blade. For the cascade, the maximum value of excitation (positive energy transfer) is found to be an order of magnitude lower compared to the isolated blade case. Further, for similar values of mean incidence angle, the range of excitation is at lower reduced frequencies for a cascade when compared with an isolated blade. A comparison with un‐stalled or classical flutter in a cascade at high Re, shows that the inter‐blade phase angle is a major factor governing flutter in both cases. Some differences are observed as well, which appear to be due to stalled flow and low Re.

Approved for public release, distribution is unlimited / The flow around second-generation controlled-diffusion blades in cascade at stall was examined experimentally through the use of a two-component laser-Doppler velocimeter. Blade surface pressure measurements were also preformed at mid span on the blades at various Reynolds numbers. Flow visualization techniques were used to observe and record the flow on the surface of the blade. A correlation between the experimental results and computational fluid dynamic predictions was attempted in order to determine the exact nature of the flow as the blades approached stall, to further assist in the development of advanced blade design. The blade surface pressure measurements showed that the mid-span section of the blade was at a lower loading than previously measured at a smaller inlet flow angle. This indicated that the blade section was at stall. The flow visualization highlighted the extent of the three-dimensional flow over the blades. The LDV measurements documented the mid-span boundary layer and wake profiles. / Ensign, United States Navy

La conception des groupes moto-ventilateurs au sein de Valeo Systèmes Thermiques et la prédiction de leurs performances aérauliques reposent majoritairement sur les méthodes de développement virtuel, i.e. la conception assistée par ordinateur et la simulation numérique de la mécanique des fluides. Dans ce cadre, le présent travail propose une méthodologie de prédiction et de minimisation de la composante tonale du bruit d'un groupe moto-ventilateur. L'approche adoptée est hybride et dissocie la génération et la propagation du bruit. La propagation en champ libre est calculée avec une formulation intégrale de l'analogie de Ffowcs-Williams et Hawkings. Dans un premier temps, les termes-sources à la surface du rotor et du stator sont calculés par une simulation numérique instationnaire. La compacité de la pale ainsi que l'influence du maillage acoustique sur la prédiction sont ensuite investiguées. Finalement, les résultats sont comparés aux mesures expérimentales. Dans un deuxième temps, les sources acoustiques à la surface du stator sont calculées avec le modèle de Sears enrichi avec des données extraites d'une simulation stationnaire du rotor complet. Avant de procéder à la prédiction acoustique, l'influence du modèle de turbulence sur les résultats finaux est évaluée à travers une comparaison entre LES et RANS pour l'écoulement autour de profils extrudés. Enfin, la problématique de minimisation du bruit tonal est traitée en tant que problème d'optimisation où la géométrie d'une aube est paramétrée et où la recherche de l'optimum est conduite par un algorithme génétique. Cette optimisation a permis de concevoir un stator moins bruyant et adapté à l'écoulement en aval du rotor étudié. / The design of fan systems at Valeo Thermal Systems and the prediction of their aerodynamic performances rely mainly on virtual development methods, i.e. computer-aided-design and computational fluid dynamics. Within this context, this dissertation develops a methodology for predicting and minimizing the tonal noise of a fan system. The hybrid approach is used, thus separating noise generation and propagation. The free-field propagation is computed via an integral formulation of the Ffowcs-Williams and Hawkings analogy. In a first step, the source terms located at the surfaces of the rotor and the stator are extracted from an unsteady numerical simulation. The compactness of the blade and the influence of the acoustic mesh on the prediction are then investigated. Finally, the computational results are compared to the experimental ones. In a second step, the acoustic sources at the surface of the stator are computed with Sears' model. Its inputs are extracted from a steady simulation of the whole rotor. Before proceeding to the acoustic prediction, the influence of the turbulence model on the final results is assessed via a comparison between LES and RANS simulations of the flow around airfoils. Finally, minimizing tonal noise is formulated as an optimization problem. The shape of a stator-blade is parametrized and the optimization is conducted with a genetic algorithm. The resulting stator is less noisy and adapted to the flow downstream of the studied rotor.

L’objectif de cette thèse est de faire progresser la modélisation de la transition de couche limite sur des aubes de turbines basse-pression fortement chargées. Cette modélisation repose sur l’utilisation du modèle de transition de Menter et Langtry utilisé pour des calculs RANS dans le code elsA. Une fois les limitations du modèle de transition clairement identifiées par une étude sur la mise en données des calculs, nous avons entrepris de modifier ce dernier. Pour cela, un processus d’optimisation a été développé afin de permettre la recalibration des fonctions de corrélation internes au modèle de transition. Cette nouvelle version du modèle nous permet d’obtenir des gains sur la modélisation d’environ 20 % sur les cas T106C du VKI en capturant mieux la transition au sein du bulbe de décollement. Ces précédents calculs correspondent à des cas idéaux, où l’on peut considérer les surfaces comme étant lisses. Cependant, nous avons aussi un besoin de se rapprocher de surfaces plus réalistes pour lesquelles les rugosités peuvent avoir un impact sur l’écoule- ment. En effet, les rugosités de surface peuvent notamment avoir un effet sur la transition. En particulier, si les rugosités entraînent le déclenchement de la transition en amont du point de décollement laminaire théorique en surface lisse, ce décollement sera supprimé. Vu nos efforts pour améliorer la prévision de la transition par bulbe de décollement par le modèle γ-Rθt, il parait intéressant que celui-ci puisse prendre en compte l’état des surfaces. Pour cela, nous avons implanté une méthode de prévision de la transition sur surfaces rugueuses développée par Stripf et al. au sein du modèle γ-Rθt. Enfin, l’utilisation du modèle de transition γ-Rθt a été étendue au modèle de turbulence k-l de Smith. / The goal of this thesis is to enhance laminar-turbulent transition modeling on high-lift low- pressure turbine blades. The presented transition modeling method relies on the Menter and Langtry transition model used in a RANS framework in the elsA solver. Once the model’s limits were clearly identified through a parametric study, we moved on to modification of the model. To do so, an optimization method was developed that allows recalibration of the model’s inner correlation functions. This new version of the model allows us to obtain modeling gains of about 20% on the VKI T106C cases through better capture of the separation-induced transition process. These previous computations correspond to ideal cases, for which surfaces may be considered as being smooth. However, we also have the need to consider more realistic surfaces for which roughness may influence the flow. Indeed, among those effects, is the potential influence of surface roughness on transition. In particular, if surface roughness induces transition up-stream of the smooth separation point, the separation bubble will be suppressed. Considering our efforts on modeling separation-induced transition with the γ-Rθt model, it seemed natural to add roughness-induced transition modeling capacities to it. To do so, we implemented in the γ-Rθt model a method developed by Stripf et al. to take into account surface roughness. Finally, the use of the γ-Rθt transition model was extended to the k-l of Smith tur- bulence model. Indeed, this turbulence model is widely used in turbomachinery. In order that our works on transition modeling over turbine blades be more widely usable, we have completed this thesis by proposing an evolution of the transition model so that it may be used alongside the k-l model.

Assessing the impact of inlet flow distortion in turbomachinery is desired early in the design cycle. This thesis introduces and validates the use of methods based on the Proper Orthogonal Decomposition (POD) to analyze clean and 1/rev static pressure distortion simulation results at design and near stall operating condition. The value of POD comes in its ability to efficiently extract both quantitative and qualitative information about dominant spatial flow structures as well as information about temporal fluctuations in flow properties. Observation of the modes allowed qualitative identification of shock waves as well as quantification of their location and range of motion. Modal coefficients revealed the location of the passage shock at a given angular location. Distortion amplification and attenuation between rotors was also identified. A relationship was identified between how distortion manifests itself based on downstream conditions. POD provides an efficient means for extracting the most meaningful information from large CFD simulation data. Static pressure and axial velocity were analyzed to explore the flow physics of 3 rotors of a compressor with a distorted inlet. Based on the results of the analysis of static pressure using the POD modes, it was concluded that there was a decreased range of motion in passage shock oscillation. Analysis of axial velocity POD modes revealed the presence of a separated region on the low pressure surface of the blade which was most dynamic in rotor 1. The thickness of this structure decreased in the near stall operating condition. The general conclusion is made that as the fan approaches stall the apparent effects of distortion are lessened which leads to less variation in the operating condition. This is due to the change in operating condition placing the fan at a different position on the speedline such that distortion effects are less pronounced. POD modes of entropy flux were used to identify three distinct levels of entropy flux in the blade row passage. The separated region was the region with the highest entropy due to the irreversibilities associated with separation.

<p>In the development of turbo machinery components it's desirable to not spend more time than necessary when setting up aero-thermal calculations to investigate uncertainties in the design. This report aims to describe general thoughts used in the development of an ICEM-mesh script and the possible configurations in the script file which enables the user to build mesh-grids with/without clearance gap at the hub and/or shroud for different blade geometries. It also aims to illustrate the performance analysis made on the Vinci LH2 turbine, a next generation upper stage engine to the Ariane 5 rocket, in which the effect of the tip gap size on the efficiency has been studied.</p><p>The calculations made have shown good agreement with experimental data. The efficiency loss due to the mixing of fluid where leakage flow passes the tip gap, which results in growth of a strong vortex, and the fluid passing the blade tip, with almost no work extracted from it, has shown a quite linear efficiency dependence depending on the tip gap size.</p>

In the development of turbo machinery components it's desirable to not spend more time than necessary when setting up aero-thermal calculations to investigate uncertainties in the design. This report aims to describe general thoughts used in the development of an ICEM-mesh script and the possible configurations in the script file which enables the user to build mesh-grids with/without clearance gap at the hub and/or shroud for different blade geometries. It also aims to illustrate the performance analysis made on the Vinci LH2 turbine, a next generation upper stage engine to the Ariane 5 rocket, in which the effect of the tip gap size on the efficiency has been studied. The calculations made have shown good agreement with experimental data. The efficiency loss due to the mixing of fluid where leakage flow passes the tip gap, which results in growth of a strong vortex, and the fluid passing the blade tip, with almost no work extracted from it, has shown a quite linear efficiency dependence depending on the tip gap size.

Dans les turbomachines conventionnelles, l’estimation des performances (rendement, puissance et rapport de pression) se fait en général en admettant l’adiabaticité de l’écoulement. Mais, de nombreuses études ayant montré l’influence négative des échanges thermiques internes et externes sur les performances des petites turbomachines dans les faibles charges et aux bas régimes, cette hypothèse ne peut plus être recevable. L’objectif principal de cette thèse est de contribuer à lever l’hypothèse d’adiabaticité.Une étude préalable de l’état de l’art a permis de relever les différents types de transferts thermiques dans les turbomachines et de circonscrire notre étude.Puis, une analyse exergétique généralisée, ayant pour but la prise en compte des deux principes de la thermodynamique, a été effectuée et l’évolution de l’indice de performance caractérisant le niveau d’énergie récupérable en fonction des échanges thermiques est étudiée.Les performances des turbomachines à fluide compressible sont généralement représentées sous forme graphique dans des systèmes de coordonnées adimensionnelles établies avec l’hypothèse d’adiabaticité. Ces cartographies couramment utilisées par les exploitants et constructeurs ne conviennent pas aux machines fonctionnant avec transferts thermiques. L’étude de la similitude des turbomachines thermiques à fluide compressible présentée dans ce travail, propose de nouvelles coordonnées adimensionnelles pouvant être utilisées aussi bien en adiabatique que dans les écoulements avec transferts thermiques.Enfin, nous proposons un protocole de mesures et un modèle numérique pour l’évaluation des transferts thermiques dans un turbocompresseur.Certains résultats obtenus montrent que les performances calculées avec l’hypothèse d’adiabaticité de l’écoulement du fluide sont surestimées. Les nouvelles lois de la similitude proposées généralisent le théorème de Rateau au fluide compressible fonctionnant dans n’importe quelle condition et permettent de calculer les échanges thermiques à chaud à partir des résultats d’essai à froid. Une donnée supplémentaire (température de refoulement) est néanmoins nécessaire pour la prédiction complète des performances et des échanges thermiques.Le modèle numérique de calcul des échanges thermiques proposé donne des résultats en accord avec ceux attendus, mais nécessite des données réelles issues de mesure sur banc pour une validation complète. / In the conventional turbomachines, calculations are done assuming adiabatic flow. But, the negative influence of external and internal heat exchange on the performance of small turbomachines at low loads and low speeds have been shown by many studies in the literature. Then, this assumption is no longer admissible. The main objective of this thesis is to help remove the assumption of adiabaticity.A study of the state of art has identified the different kinds of heat transfer and defined the limits of our investigations.Afterwards, a generalized exergy analysis whose main goal is to take into account the two principles of thermodynamics has been made and the variation of exergy performance versus heat transfer has been studied.The maps currently used are made with the assumption of adiabaticity. The laws of similarity in turbomachines working with compressible fluid studied propose new dimensionless coordinates that can be used in any operating condition (adiabatic or not).Finally, we present a measurement protocol and a numerical model for calculating heat transfer in a turbocharger.Some results from our work indicate that the performance of thermal turbomachinery announced regardless of thermal heat exchanges are found to be overestimated.The new laws of similarity proposed generalize the Rateau’s theorem to compressible fluid flow in any operating condition and can be used to calculate heat transfer from adiabatic test results. Supplementary information is still required for the complete prediction of performance and heat transfer.The numerical model for calculating heat transfer gives some results that are in agreement with those expected. But actual data from test bench are required for complete validation.