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The current study investigates a range of issues relating to the use of a vane-recessed tubular-passage casing treatment as a passive stall control technique in a multistage axial-flow compressor. The focus of the research was to determine whether such a treatment could delay the initiation of stall at lower mass flow rates as well as providing the most beneficial improvement in flow characteristics without sacrificing compressor efficiencies. Specific objectives of this study were to examine possible improvements or deterioration in the flow characteristics including stall margin, peak pressure rise coefficients and maximum efficiency in a multistage axial flow compressor. A casing treatment in addition to several spacer rings was developed from two initial designs and tested on the first stage of a low speed three-stage axial-flow compressor with a (0.7) hub to tip diameter ratio. The treatment configuration consisted of three parts: an outer casing ring, with a tubular shaped passage on the inside diameter, a set of 120 evenly spaced curved vanes, and a shroud or inner ring. The casing treatment was positioned following the inlet guide vanes upstream and partly covering the tip of the rotor blades. The main parts of the casing treatment including the recessed vanes in addition to some of the spacer rings were manufactured from high quality acrylic. Eight additional spacer rings of various shapes and geometry were added. The first ring held and partly covered the IGVs, in front of the casing treatment. The rotor tip exposure ratio was thought to have a significant impact on the effectiveness of the casing treatment. Therefore the other seven rings were used to provide the desired uncovered region of the rotor tip axial chord of about 10% in order to provide a range of exposures of (23.2%, 33.3%, 43.4%, 53.5%, 63.6%, 73.7%, and 83.8%). The results showed significant improvements in stall margin in all treated casing configurations along with insignificant efficiency sacrifices in some compressor builds. About (28.56%) of stall margin improvement in terms of corrected mass flow rate was achieved using a casing treatment with a (33.3%) rotor tip exposure. The compressor build with (0.535) rotor exposure ratios was the best configuration in terms of efficiency gain and loss characteristics. This build was able to provide the highest values of the maximum efficiencies in comparison with the performance achieved from the solid casing. An improvement of (1.81%) in the maximum efficiency in terms of the overall total-total pressure ratio, in association with a (22.54%) stall margin improvement in terms of the corrected mass flow rates were achieved by the application of this treatment configuration. The improvement in the peak pressure rise coefficients in terms of the overall total-total pressure ratio, obtained from this build was (2.33%). The compressor configuration using a casing treatment with a (0.636) rotor exposure ratio was the best build in terms of the pressure rise coefficients. This configuration was able to provide highest value of the peak pressure rise in comparison with the characteristics achieved from the datum build. An improvement of (2.65%) in the peak pressure rise coefficient in terms of the overall total-total pressure ratio, in association with a (22.49%) improvement in stall margin in terms of the corrected mass flow rates was achieved from this casing treatment build. The improvement in maximum efficiency in terms of the overall total-total pressure ratio, obtained from this build was (1.03%). The results suggest that the vane-recessed tubular-passage casing treatment designed as part of this investigation achieved the objectives, which were established for the research. In the majority of instances it not only produced gains in flow range, pressure rise coefficients and efficiencies, but also enabled the rotating stall, which developed at much lower mass flow rates in the compressor, to become progressive rather than abrupt.

L’étude effectuée au cours de cette thèse a permis la caractérisation expérimentale des instabilités aérodynamiques se développant dans un compresseur centrifuge et une première évaluation de l’efficacité d’une stratégie de contrôle par aspiration de couche limite. Le compresseur, développé par Safran Helicopter Engines et dénommé Turbocel, est composé d’une roue directrice d’entrée, d’un rouet centrifuge splitté, d’un diffuseur radial aubé et splitté et d’un redresseur axial. Des travaux numériques antérieurs réalisés au Laboratoire de Mécanique des Fluides et d’Acoustique ont montré, aux bas régimes de rotation, un comportement singulier caractérisé par une structure d’écoulement dite « alternée » impliquant deux canaux adjacents du diffuseur radial. L’étude stationnaire réalisée sur l’ensemble des régimes de rotation du compresseur a conduit à une ségrégation des vitesses de rotation suite à l’établissement d’une variable– le taux d’asymétrie - caractérisant l’asymétrie de l’aérodynamique du diffuseur. Ce taux, quasi nul à très basse vitesse de rotation, croît jusqu’à atteindre un maximum à vitesse de rotation intermédiaire, puis s’effondre pour ré-augmenter légèrement. Une analyse fine des données instationnaires acquises à bas régimes a permis la description de deux modes de fonctionnement du compresseur associés à des structures de décollements différentes dans le diffuseur. Le premier mode est caractérisé par l’oscillation à une fréquence de l’ordre de 42 Hz d’un décollement localisé sur la face en dépression des aubes principales du diffuseur. Le second mode, à 12Hz, associé au pompage modéré du compresseur, correspond à la mise en place d’un schéma alterné et à son oscillation sur deux canaux adjacents du diffuseur.Les origines probables de ces différents modes de fonctionnement sont discutées à partir de considérations • aérodynamiques -- la mise en place d’une recirculation en tête de rouet est suspectée d’influer sur le taux d’asymétrie en modifiant l’incidence en entrée de diffuseur,• géométriques -- le nombre et le calage des aubes du diffuseur radial ainsi que la distance inter-roue indiquent une prédisposition du diffuseur à fonctionner en régime alterné sous certaines conditions d’incidence,• aéro-acoustiques -- un accrochage des fréquences aérodynamiques avec les fréquences des ondes acoustiques du banc d’essai semble se produire. Enfin, les résultats sur le contrôle d’écoulement par aspiration de couche limite à régime partiel sont présentés. Une amélioration du rendement est observée à certains points de fonctionnement, mais aucune extension de la plage de fonctionnement du compresseur n’est mesurée. Sans l’atténuer, l’aspiration permet de contrôler sur quels canaux s’établit le régime alterné. / This thesis presents an experimental characterization of the evolution of aerodynamic instabilities in a centrifugal compressor, and a first evaluation of the effectiveness of boundary layer suction as a control strategy. The compressor used in this study is Turbocel, a centrifugal compressor developed by Safran Helicopter Engines, featuring inlet guide vanes, a backswept splittered unshrouded impeller, a splittered vaned radial diffuser and axial outlet guide vanes.Previous numerical work, conducted at the Laboratoire de Mécanique des Fluides et d’Acoustique de Lyon, revealed a unusual behaviour of the compressor at low rotational speeds characterized by a distinctive alternate flow structure in the radial diffuser that develops across two adjacent blade channels. The steady analysis, which was conducted over the full operating range of rotational speeds, led to the distinction of different operating zones, following the establishment of a new indicator variable - the asymmetry rate - characterizing the asymmetry of the diffuser aerodynamics. This rate, which is close to zero at very low rotation speed, increases until it reaches a maximum value at intermediate rotational speed, before collapsing and slightly increasing again near the nominal rotational speed.Analysis of the unsteady data acquired at low speeds allowed for the characterization of two compressor operating modes, associated with different flow phenomena in the stalled diffuser. The first mode is characterized by the oscillation of a separation at 42 Hz, on the suction side of the main blades in the diffuser. The second mode, at 12Hz, associated with mild surge of the compressor, corresponds to the emergence of an alternate pattern of unsteady flow separation that occurs across two adjacent channels of the diffuser.The probable causes for these different operating modes are discussed in the context of different considerations:• aerodynamic -- the formation of a recirculation near the tip of the impeller is suspected to affect the asymmetry rate by changing the incidence angle at the diffuser inlet.• geometric -- the number and the stagger angle of the radial diffuser blades as well as the distance between the impeller and the diffuser may result in a predisposition of the diffuser to operate in an alternating mode, under certain conditions of incidence.• aero-acoustic -- as there is evidence of a lock-in of the aerodynamic frequencies with the acoustic modes of the test rig.Finally, boundary layer suction is explored as a means of flow control at partial rotational speed. Improvements in performance were observed for some operating points, however no extension of the compressor operating range was measured. Although boundary layer suction did not allow for the intensity of the oscillating separation pattern in the diffuser to be reduced, it was found to be an effective means of controlling the location of the alternate flow structure in the diffuser.

Current lean-premixed fuel injector designs have shown great potential in terms of reducing emissions of pollutants, but such designs are susceptible to combustion instabilities in which aerodynamic instability plays a major role and also has an effect on mixing of air and fuel. In comparison to prototype testing with combustors running in operating conditions, computational approaches such as Large Eddy Simulations (LES) offer a much more cost-effective alternative in the design stage. However, computational models employed by LES require validation by experimental data. This is one of the main motivations behind the present experimental study. Combined particle image velocimetry (PIV) and planar laser induced fluorescence (PLIF) instrumentation allowed simultaneous measurements of velocity vector and a conserved scalar introduced into the fuel stream. The results show that the inner swirl shear layer features two pairs of vortices, which draw high concentration fuel mixture from the central jet into the swirl stream and causes it to rotate in their wakes. Such periodic entrainment also occurs with the characteristic frequencies of the vortices. This has clear implications for temporal variations in fuel/air ratio in a combusting flow; these bursts of mixing, and hence heat release, could be a possible cause of mixing-induced pressure oscillation in combusting tests. For the first time in such a flow, all 3 components of the turbulent scalar flux were available for validation of LES-based predictions. A careful assessment of experimental errors, particularly the error associated with spatial filtering, was carried out. Comparison of LES predictions with experimental data showed very good agreement for both 1st and 2nd moment statistics, as well as spectra and scalar pdfs. It is particularly noteworthy that comparison between LES computed and measured scalar fluxes was very good; this represents successful validation of the simple (constant Schmidt number) SGS model used for this complex and practically important fuel injector flow. In addition to providing benchmark data for the validation of LES predictions, a new experimental technique has been developed that is capable of providing spatially resolved residence time data. Residence times of combustors have commonly been used to help understand NOx emissions and can also contribute to combustion instabilities. Both the time mean velocity and turbulence fields are important to the residence time, but determining the residence time via analysis of a measured velocity field is difficult due to the inherent unsteadiness and the three dimensional nature of a high-Re swirling flow. A more direct approach to measure residence time is reported here that examines the dynamic response of fuel concentration to a sudden cutoff in the fuel injection. Residence time measurement was mainly taken using a time-resolved PLIF technique, but a second camera for PIV was added to check that the step change does not alter the velocity field and the spectral content of the coherent structures. Characteristic timescales evaluated from the measurements are referred to as convection and half-life times: The former describes the time delay from a fuel injector exit reference point to a downstream point of interest, and the latter describes the rate of decay once the effect of the reduced scalar concentration at the injection source has been transported to the point of interest. Residence time is often defined as the time taken for a conserved scalar to reduce to half its initial value after injection is stopped: this is equivalent to the sum of the convection time and the half-life values. The technique was applied to a high-swirl fuel injector typical of that found in combustor applications. Two test cases have been studied: with central jet (with-jet) and without central jet (no-jet). It was found that the relatively unstable central recirculation zone of the no-jet case resulted in increased transport of fuel into the central region that is dominated by a precessing vortex core, where long half-life times are also found. Based on this, it was inferred that the no-jet case may be more prone to NOx production. The technique is described here for a single-phase isothermal flow field, but with consideration, it could be extended to studying reacting flows to provide more insight into important mixing phenomena and relevant timescales.

Dans le contexte économique et environnemental actuel, la prochaine génération de moteurs d’avion devra offrir opérabilité, compacité et hauts rendements. Les compresseurs demeurent une des pièces critiques de ces moteurs, et leur conception un challenge. À débit réduit, leur plage de fonctionnement est contrainte par la limite de pompage, phénomène hautement instable et dangereux. À ce jour, peu d’études expérimentales sur un compresseur en situation de pompage ont été réalisées, étant donné le danger inhérent pour les installations. Dans ce cadre, la simulation numérique peut apporter des informations sur le développement des instabilités aérodynamiques et aider à la prévision de la limite de pompage. L’objectif du travail présenté dans cette thèse est de mettre en place une méthode afin de simuler numériquement l’entrée en pompage et un cycle complet de l’instabilité avec le code elsA. Le cas test retenu est le compresseur de recherche axial multi-étage CREATE dessiné par Snecma, et étudié expérimentalement par le LMFA. Des études antérieures ont montré le rôle joué par les volumes entourant le compresseur ; l’originalité de cette étude réside donc dans l’inclusion des volumes du banc d’essai dans la simulation du compresseur. Une des difficultés inhérentes à la simulation de ces instabilités est leur temps caractéristique, qui représente plus d’une centaine de rotations de la machine. Le calcul a donc nécessité le recours à une approche massivement parallèle ; environ un million d’heures CPU ont été utilisées pour décrire le cycle. Enfin, compte tenu du retournement de l’écoulement dans le compresseur, les conditions aux limites ont été modifiées pour pouvoir s’adapter aux changements de sens de l’écoulement. La simulation a permis de décrire l’entrée en pompage et un cycle complet de l’instabilité. La comparaison avec les données expérimentales montre que les caractéristiques du cycle sont correctement prédites (phénomènes physiques précurseurs de l’instabilité, durée du cycle..). En parallèle, une étude acoustique a été menée afin de mettre en évidence les modes propres du banc d’essai. L’analyse de ces résultats a notamment montré le rôle de l’acoustique dans le déclenchement du pompage. Les différentes phases du cycle de pompage sont ensuite étudiées, et caractérisées (déclenchement, débit inversé, récupération et recompression). Ce travail a généré une base de données qui permet de mieux comprendre les instabilités qui se développent dans ce type de machine. À terme, ces résultats pourront être utilisés pour élaborer et valider des modélisations du phénomène de pompage moins coûteuses, pouvant intervenir dans un cycle de conception. / In order to deal with the current economical and environmental context, the next engine generation will need to offer great operability, compactness and high efficiency. In aircraft engines, the compressor remains one of the critical components, and its design is still a challenging task. At low massflow rate, their operability is bounded by the surge limit, surge being a highly unstable and dangerous phenomenon. Today, few experimental studies on compressor surge are available because of the inherent threat to the facility. In that context, numerical simulation can bring about information on the onset of aerodynamic instabilities and help to predict the surge limit. The work presented in this PhD thesis aims at setting up a method to perform the numerical simulation of surge inception and of an entire cycle of the instability with the CFD code elsA. The chosen test case is the axial multistage research compressor CREATE designed and built by Snecma, and experimentally studied at LMFA. Previous studies have pointed out the role of the volumes adjacent to the compressor ; the originality of this work is thus the inclusion of the volumes of the test-rig in the simulation of the compressor. One of the difficulties inherent to the simulation of those instabilities is their characteristic time of at least one hundred revolutions of the machine. Hence the computation has required a massively parallel approach and about one million CPU hours. Finally, given that the flow reverses during a surge cycle, the boundary conditions have been modified to be able to cope with the flow inversions. The simulation was able to capture surge inception and the entire cycle of the instability. The comparison with the experimental data showed that the main patterns of the cycle are correctly predicted (precursor phenomena of surge, duration of the cycle...). In the meantime, an acoustic study has been performed in order to isolate the eigenmodes of the test-rig. The analysis of the results pointed out the role of acoustic phenomena in surge inception. The different phases of the cycle are then studied and characterized (surge inception, reversed-flow phase, recovery and repressurization). This work has incremented a database that allows a better understanding of the instabilities that develop in this kind of machine. From now on, those results may help to elaborate and validate cheaper models of the surge phenomenon to be used in the design process.