A rapid and low noise switch from RANS to WMLES on curvilinear grids with compressible flow solvers

International audienceA turbulent inflow for a rapid and low noise switch from RANS to Wall-Modelled LES on curvilinear grids with compressible flow solvers is presented. It can be embedded within the computational domain in practical applications with WMLES grids around three-dimensional geometries in a flexible zonal hybrid RANS/LES modelling context. It relies on a physics-motivated combination of Zonal Detached Eddy Simulation (ZDES) as the WMLES technique together with a Dynamic Forcing method processing the fluctuations caused by a Zonal Immersed Boundary Condition describing roughness elements. The performance in generating a physically-sound turbulent flow field with the proper mean skin friction and turbulent profiles after a short relaxation length is equivalent to more common inflow methods thanks to the generation of large-scale streamwise vorticity by the roughness elements. Comparisons in a low Mach-number zeropressure-gradient flat-plate turbulent boundary layer up to Reθ = 6 100 reveal that the pressure field is dominated by the spurious noise caused by the synthetic turbulence methods (Synthetic Eddy Method and White Noise injection), contrary to the new low-noise approach which may be used to obtain the low-frequency component of wall pressure and reproduce its intermittent nature. The robustness of the method is tested in the flow around a three-element airfoil with WMLES in the upper boundary layer near the trailing edge of the main element. In spite of the very short relaxation distance allowed, self-sustainable resolved turbulence is generated in the outer layer with significantly less spurious noise than with the approach involving White Noise. The ZDES grid count for this latter test case is more than two orders of magnitude lower than the Wall-Resolved LES requirement and a unique mesh is involved, which is much simpler than some multiple-mesh strategies devised for WMLES or turbulent inflow

Amélioration de la méthodologie de calcul de bruit de jet à l'aide de la ZDES mode 3 et d'une méthode de génération de turbulence silencieuse

International audienceThis article presents Zonal Detached Eddy Simulations (ZDES) of a single stream jet. Focus is put on the improvement of the turbulence modelling methodology aimed at taking into account the nozzle internal boundary layer dynamics to better simulate the initial stage of the mixing layer development. To this effect, ZDES mode 3 (acting as a Wall Modelled Large Eddy Simulation, WMLES) is used inside the nozzle and turbulence is triggered with tripping dots implemented in the computation using the Immersed Boundary Condition method. Results are compared to experimental data and a standard ZDES mode 2 simulation (running in URANS mode for the nozzle boundary layer). The novel methodology involving ZDES mode 3 is successful in removing the spurious RANS-to-LES transition noise in the early stages of the mixing layer observed in the ZDES mode 2 simulation which opens the way to the accurate prediction of jet noise on complex aircraft configurations with ZDES

Flow control for road vehicle drag reduction

This thesis covers topics that span bluff-body aerodynamics, hybrid RANS-LES CFD methods, flow control and model-order reduction. These topics arise from investigating the flow past three geometries: the bullet shaped D-body, the canonical squareback Ahmed body and the commerical Nissan NDP. The study on the D-body was aimed at transitioning the research group from the restrictive block-structured formulated StreamLES solver to the more flexible OpenFOAM code that can use unstructured meshes. Linear feedback control for base pressure increase was applied as was done in the work by Dalla Longa et al. (2017). Identification of the plant, G(s), that represents the wake's response to forcing was completed and correlated well with the results from Dalla Longa et al. (2017). The same can also be said of the sensitivity based designed feedback control law, K(s). When applied in simulation, an attenuation of the base pressure fluctuations was, as desired, achieved, although the base pressure increased by 24.5% as opposed to the 38% achieved by Dalla Longa et al. (2017). In the study on the squareback Ahmed body, wall-resolving (WRLES) and wall-modelled (WMLES) large eddy simulation were successfully applied. First, a simulation setup that is both able to resolve wake bimodality, while remaining reasonable in computational resource use, was created. Subsequently, variants of this setup were used to identify a flow feature that plays a critical role in forcing wake bimodality events. More specifically, a heavily under-resolved WMLES simulation in which both the near-wall and part of the outer-region of the turbulent boundary layer are Reynolds-averaged did not capture the front recirculation bubble near the Ahmed body nose; neither did it resolve a bimodal wake switching event. Meanwhile, the simulations with a more refined near-wall mesh did capture the front separation bubble as well as bimodal switching events of the wake. This front separation bubble sends out powerful hairpin vortices that interact with the rear wake. Specifically, these vortices go on to produce significant amounts of TKE, which, upon convection to the rear of the Ahmed body, ultimately help trigger a bimodal event. The Ahmed body study also involved the application of linear feedback control for drag reduction as was done in the D-body study. In the short term, mean blowing did lead to a base pressure increase, but as the zero-net-mass-flux (ZNMF) jet settled, it oscillated around zero making its effects indiscernible. The final geometry analyzed was the Nissan NDP. This was done by performing benchmark wall-resolving LES (WRLES). First, the benefit of appending a rear cavity to an otherwise "squareback" geometry was assessed. It was concluded that the cavity allows the wake to move more freely about the rear base. Specifically, the wake is freed from its more restricted motion that is present with the "squareback" Nissan NDP. In doing so, the drag reduction achieved with the cavity appendage is about 13.6%. Work on the Nissan NDP also involved an assessment of a moving ground in the simulation. It was concluded that, in the stationary ground simulation, flow detachment at the ground where the flow exits from the underbody has an adverse drag effect. In other words, although moving ground simulations better replicate the real-world conditions, the stationary ground variant is in this case more conservative, as it returns slightly higher drag values.Open Acces

Kompressible CFD-Simulationen von Aeroakustik für Automobilanwendungen

In this work, a direct noise computation method based on a low-Mach number flow solver is investigated. The new solver is implemented in the finite volume framework of the software OpenFOAM, accompanied with a new acoustic damping model for reducing spurious noise. The new solver is utilised to calculate noise generation and propagation for automotive applications. In order to validate the applicability of the low-Mach flow solver, a benchmark consisting of two-struts is calculated. The simulated aerodynamic near field as well as aeroacoustic far field are compared to wind tunnel measurements. The acoustic far field is computed using the direct method as well as a hybrid method. Both methods are evaluated based on comparing far field spectra and directivity patterns with experimental results. After validating the applicability of the low-Mach number solver, the topic of spurious noise generation in direct noise computation is addressed. Different spurious noise sources are presented and their generation mechanisms are investigated. Afterwards, two different strategies for spurious noise reduction, namely selective acoustic damping and numerical grid stretching, are discussed and validated. The acoustic damping model can substantially damp out spurious noise generated at grid interfaces without affecting the turbulence. It is also observed that the direction of grid refinement determines the direction of propagation of spurious noise. The strategies for spurious noise reduction are then applied on a side-mirror test case. For this, a new algorithm for automated and directional grid stretching is implemented. Spurious noise generation in the vicinity of the mirror’s surface as well as in the mirror’s wake could be substantially reduced and a quantitative analysis based on frequency-wavenumber spectra in the wake of the mirror is performed. Finally, the proposed flow solver, along with the strategies for spurious noise reduction, is used to directly compute noise generation on a generic vehicle model. Two different variants are calculated and the effect of the A-pillar and the side-mirror regarding their contribution to the acoustic waves on the side-window is investigated. Aerodynamic as well as aero- and vibroacoustic spectra on the side-window are calculated and compared to wind-tunnel measurements. For both variants, the results calculated using the direct method show good agreement with experimental data.In dieser Arbeit wird eine Methode zur direkten Berechnung von Aeroakustik basierend auf einen Strömungslöser für kleine Mach-Zahlen untersucht. Der Strömungslöser wird mit einem neuen Dämpfungsmodell für die Reduktion numerischer Schallwellen im finite Volumen Code OpenFOAM implementiert, und für die Berechnung der Entstehung und Ausbreitung von Schallwellen im Automobilbereich angewandt. Zur Validierung des neuen Strömungslösers, wird ein Benchmark, der aus zwei parallelen Streben besteht, berechnet. Das simulierte aerodynamische Nahfeld sowie das aeroakustische Fernfeld werden mit Windkanalmessungen verglichen. Das akustische Fernfeld wird mit der direkten sowie mit einer hybriden Methode berechnet. Beide Methoden werden anhand der Fernfeldspektren sowie der Richtcharakteristiken mit experimentellen Daten bewertet. Nach der Validierung des Strömungslösers wird die Entstehung von numerischen Störungen in der direkten Methode analysiert. Es werden verschiedene Quellen numerischer Störungen sowie deren Entstehungsmechanismen dargestellt. Anschließend werden zwei verschiedene Strategien zur Reduktion von Störungen diskutiert und validiert. Das Dämpfungsmodell zeigt sein Potenzial bei der Reduktion von numerischen Schallwellen ohne Beeinflussung der Turbulenz. Es wird außerdem gezeigt, dass die Richtung einer Verfeinerung des numerischen Gitters die Richtung der Ausbreitung numerischer Schallwellen bestimmt. Die Strategien zur Reduktion numerischer Störungen werden weiterhin an einem einzelnen Seitenspiegel angewandt. Dafür wird ein neuer Algorithmus für eine automatisierte und richtungsdefinierte Gitterexpansion implementiert. Die Amplitude numerischer Störungen, die im Spiegelnachlauf entstehen, werden mit Hilfe einer Frequenz-Wellenzahl Analyse quantitativ untersucht. Es zeigt sich, dass das Dämpfungsmodell diese Störungen deutlich reduziert. Abschließend wird der Strömungslöser zusammen mit den vorgeschlagenen Strategien in einer direkten Aeroakustikberechnung eines generischen Fahrzeugmodells angewandt. Es werden zwei unterschiedliche Varianten berechnet und der Einfluss der A-Säule und des Seitenspiegels bezüglich ihres akustischen Beitrags auf der Seitenscheibe untersucht. Sowohl aerodynamische als auch aero- und vibroakustische Spektren werden auf der Seitenscheibe berechnet und mit Windkanalmessungen verglichen. Für beide Varianten zeigen die Ergebnisse der direkten Methode gute Übereinstimmung mit den experimentellen Daten

Predictive Capabilities of Laminar-Turbulent Transition Models for Aerodynamics Applications

Laminar-turbulent boundary-layer transition has a demonstrable impact on the performance of aerospace vehicles. The ability to accurately predict transition is integral to properly capturing relevant flow physics. Traditionally, computational fluid dynamics simulations are performed fully turbulent, meaning that laminar flow is neglected. This, however, can result in errant predictions of vehicle performance as quantities such as skin-friction drag may be overpredicted. Resultingly, development of Reynolds-averaged Navier-Stokes transition models has seen significant attention over the last decades in order to model transition and realize the performance improvements of laminar flow. In this work, the behaviors of several different transition-prediction methods are analyzed both for their ability to predict transition and vehicle performance. The popular local-correlation transition model is assessed analytically and numerically. It is found that there exists a singularity near the wall after transition to turbulence. This results in singular-like behavior of the destruction term of the turbulent kinetic energy equation and prevents the model from ever achieving asymptotic grid convergence. A turbulence index was developed to more robustly and accurately detect transition relative to other quantities such as turbulent intermittency. The behavior of the amplification factor transport (AFT) model was examined for a four-bladed helicopter rotor undergoing dynamic pitching conditions. The transition front predicted by the AFT model agreed well with experiments, with exception to that during the upstroke of the pitch cycle. The effect of freestream turbulence intensity on transition was examined by varying the critical $N$-factor, finding that as turbulence intensity increased, the transition front moved increasingly further upstream throughout the pitch cycle. Additionally, large eddy simulations were performed for a rotorcraft airfoil undergoing dynamic pitching conditions. A laminar separation bubble was found to be the primary mechanism of transition, finding also that the length of the separation bubble decreased as the pitch angle increased. A Kelvin-Helmholtz-like instability was identified near the aft end of the separation bubble which drives transient bursting of the bubble and is partially responsible for the behavior of the transition front. An additional investigation using the AFT model found that predicted transition agreed well, but predicted a shorter separation bubble

Computational aeroacoustic study of aircraft slat tracks and cut-outs

As one of the major contributors to aircraft noise, the noise generated by high-lift devices has been explored for many years. However, the noise related to the slat track system, which includes all the extrusive components connecting the slat and the main element, is still generally studied through experimentation due to the complex geometry. In this project, the aerodynamics and aeroacoustics of the slat track and cut-out, especially the main element cut-out, were investigated through numerical simulations for the first time. Two methods were employed in this work. Noise propagation is first studied via a compact source model to evaluate the contribution of each source individually and to investigate the influence of the slat track system on the noise propagation in the slat region. The APE-IV system was employed but modified by using a more accurate expression of enthalpy perturbation to calculate the acoustic fields. The results show that both the slat track system and the background flow modify the sound propagation path. The energy radiated towards the ground is increased due to the interaction of sound waves with the slat track system and the background flow. Detached eddy simulations were run to investigate the mechanisms of the slat track and cut-out noise generation. Major noise sources in the slat cove region are identified and a noise generation feedback loop is proposed. The results show that the increment of noise levels due to the existence of the slat track system is two-fold. The slat track and the cut-out generate noise individually and they also amplify the noise generated within the slat region when the sound waves propagate though the slat cove area and interact with the slat track and cut-out. The dominant frequencies of the spectrum are seen to shift towards the high frequencies due to these added on components. In this work, two kinds of possible noise attenuation approaches were proposed and studied. Geometries based on replacing the sharp cut-out on the main element leading edge with an edge-rounded or a sealed cut-out have been proved to be able to reduce the cut-out noise significantly. Application of acoustic bulk absorbing material can also attenuate the cut-out noise efficiently for a certain range of frequencies

Towards Understanding Stall Cells and Transonic Buffet Cells

L’enveloppe de vol des avions civils est contrainte par plusieurs phénomènes aérodynamiques. Parmi ceux-ci figurent le décrochage et le tremblement. Le décrochage est défini comme une perte de portance causée par la séparation de l’écoulement sur une voilure. Il survient le plus souvent à faible vitesse et à un angle d’attaque élevé, influençant donc les performances au décollage et à l’atterrissage. Le tremblement est quant à lui une vibration de la structure de l’appareil causée par des efforts aérodynamiques instationnaires. Plus spécifiquement, le tremblement transsonique est défini comme une instabilité aérodynamique causant une oscillation de la position de l’onde de choc sur une aile. Il se produit à haute vitesse, cause un chargement instationnaire et des vibrations, et limite la vitesse maximale de l’aéronef. Ces deux phénomènes nuisent au contrôle de l’appareil et doivent être évités. Cependant, le tremblement et le décrochage demeurent mal compris et difficiles à prédire. Pour ces raisons, les avionneurs utilisent des lois empiriques pour prédire ces limites de l’enveloppe de vol, qui sont par la suite vérifiées lors des tests en vol. Une meilleure compréhension de ces phénomènes permettrait donc de réduire les incertitudes sur leur prédiction et de réduire les marges par rapport aux limites de décrochage et de tremblement lors de la conception, et ainsi d’améliorer les performances des prochaines générations d’avions. Le décrochage basse vitesse implique des dynamiques complexes, parmi lesquelles les cellules de décrochage. Ces dernières sont une variation de la position de la séparation d’une section en envergure de l’aile à l’autre. Elles sont associées à une structure en forme de champignon formée de tourbillons contrarotatifs. Des expériences réalisées sur des modèles s’étendant sur la largeur d’une soufflerie et des simulations sur des ailes infinies ont montré leur présence. Cela indique la présence d’une instabilité qui transforme l’écoulement bidimensionnel en un écoulement tridimensionnel. Dans le régime transsonique, le tremblement est caractérisé par une oscillation de la position de l’onde de choc le long de la corde causée par une interaction entre l’onde de choc et une couche limite décollée. La visualisation de l’écoulement sur des cas tridimensionnels montre la présence de structures formées par une variation de la position du choc le long de l’envergure. Ce phénomène a été nommé cellule de tremblement et présente une structure en forme de champignon similaire à celle qui est observée en condition basse vitesse. Cette observation mène à l’idée d’étudier les deux phénomènes avec les mêmes méthodes. Cette thèse vise l’amélioration de la compréhension du décrochage et du tremblement en étudiant ces cellules. L’hypothèse soutenant ce travail est que les cellules de décrochage et de tremblement sont en fait le même phénomène sous deux régimes d’écoulement distincts. Pour examiner cette hypothèse, les cellules sont d’abord caractérisées avec des calculs instationnaires des équations de Navier-Stokes en moyenne de Reynolds (URANS) et des simulations zonales des tourbillons décollés (ZDES). Des configurations formées par des ailes infinies sont sélectionnées afin d’exclure l’effet de la géométrie et des conditions aux limites sur la présence d’un écoulement tridimensionnel. Ces configurations sont obtenues en fermant le domaine de calcul avec des conditions de périodicité dans la direction de l’envergure. Ces simulations montrent la présence de cellules de décrochage et de tremblement. Ces dernières sont stationnaires pour le cas des ailes sans flèche et les cellules sont convectées dans la direction de l’envergure lorsqu’un angle de flèche est ajouté. La vitesse de cette convection est proportionnelle à la vitesse de l’écoulement en champ lointain, projetée dans la direction du bord d’attaque de l’aile. Le nombre de cellules augmente avec l’envergure. De plus, la visualisation des lignes de frottement pariétal montre une topologie d’écoulement en forme de champignon. Les simulations du décrochage sur les ailes sans flèche permettent d’obtenir des cellules de décrochage stationnaires. Dans le cas du tremblement, une instabilité bidimensionnelle est présente ce qui induit un écoulement instationnaire pour tous les angles de flèche. Cependant, cette étude n’exclut pas la possibilité d’observer des cellules de tremblement stationnaire. Ces résultats permettent de distinguer deux effets dominant le tremblement tridimensionnel. Le premier est l’instabilité de tremblement bidimensionnel classique qui survient à une fréquence constante, peu importe l’angle de flèche. Le second est une convection des cellules de tremblement avec une fréquence proportionnelle à l’angle de flèche. Cela explique la fréquence de tremblement plus élevé sur les avions civils que les cas bidimensionnels, puisque ces ailes ont un angle de flèche élevé. Pour supporter l’hypothèse de base et proposer un modèle expliquant l’origine des cellules de décrochage et de tremblement, elles ont aussi été étudiées à l’aide d’analyses de stabilité globale. Ces analyses permettent de caractériser la transition entre une solution bidimensionnelle stationnaire vers un état instationnaire et/ou tridimensionnel dans le régime linéaire. La méthodologie utilisée dans cette étude suppose que les perturbations autour de l’état de base sont périodiques dans la direction de l’envergure afin de réduire le coût de calcul. Ces analyses montrent qu’un mode globalement instable est responsable de la transition d’un état bidimensionnel vers une solution tridimensionnelle. Ce mode a une fréquence nulle pour le cas de l’aile sans flèche, et a une fréquence d’oscillation pour le cas des ailes en flèche. Cette fréquence est cohérente avec celle des structures convectées aux vitesses trouvées dans les simulations URANS. Dans le cas du décrochage, la longueur d’onde de ces cellules est aussi bien prédite. Ce faisant, la fréquence est aussi bien prédite. Cependant, les longueurs d’onde trouvées pour les cas de tremblement ne corroborent avec celles obtenues dans les simulations instationnaires. Toutefois, on observe que les simulations instationnaires bifurquent vers un état tridimensionnel en suivant la longueur d’onde présente dans les analyses de stabilité. Ainsi, des effets non linéaires sont probablement responsables du changement de longueur d’onde. On observe aussi que la présence de cellules de décrochage et de tremblement coïncide avec une pente négative de la relation entre le coefficient de portance et l’angle d’attaque des solutions utilisées comme champs de base pour les études de stabilité. Cela est aussi observé dans les calculs URANS où les solutions demeurent invariantes en envergure tant que cette pente est positive. Des analyses de stabilité sont aussi réalisées pour un cas à haut nombre de Reynolds et pour des nombres de Mach allant du régime subsonique au transsonique. Un mode tridimensionnel instable est trouvé pour tous les nombres de Mach et un dimensionnement par l’épaisseur de la séparation est appliqué avec succès pour comparer la longueur d’onde de ces modes. Pour finir, un modèle non visqueux (surface portante) avec correction de viscosité est utilisé pour étudier les cellules de décrochage. Ce modèle est capable de reproduire les principales caractéristiques des cellules de décrochage, comme leur nombre sur une aile ayant un rapport d’aspect donné. Ce modèle prédit la présence de cellules dès que la pente de la portance selon l’angle d’attaque devient négative. Il permet aussi de penser que les cellules ne seront observées que pour des décrochages du bord de fuite, pour lesquels la séparation se propage progressivement du bord de fuite vers le bord d’attaque lorsque l’angle d’attaque augmente. Cette conclusion est faite puisqu’un décrochage de bord d’attaque est caractérisé par une perte soudaine de portance et que la plage d’angle d’attaques où la pente de la portance est négative est infiniment mince. Cette conclusion est pour un cas parfait et le comportement sur une géométrie complexe comme celle d’un avion devrait être plus complexe. Ce modèle basse fidélité permet d’obtenir des solutions aillant le même nombre de cellules que des expériences sur des ailes rectangulaires. Un point intéressant de ce modèle est le fait que les effets visqueux sont traités de façon macroscopique à travers une polaire de portance. Le modèle n’inclut aucune information sur la topologie de l’écoulement décollé. Pour cette raison, on peut conclure qu’une instabilité de la distribution de la circulation permet d’expliquer les cellules de décrochage, et par extension les cellules de tremblement puisqu’elles ont été reliées à une instabilité similaire. Ensemble, ces résultats montrent que le tremblement bidimensionnel et tridimensionnel sont deux phénomènes différents. Ce dernier a plus de similarité avec l’instabilité causant les cellules de décrochage qu’avec une instabilité entre l’onde de choc et la couche limite décollée. Il ne faudrait donc pas l’appeler tremblement tridimensionnel, mais plutôt cellules de décrochage. Ces résultats montrent aussi que les cellules de décrochage/tremblement sont reliées à une instabilité de la circulation sur l’aile (cf. théorie de la ligne portante). Les idées et les méthodes présentées dans cette thèse pourront être utilisées dans un cadre industriel pour améliorer la modélisation de ces phénomènes, et améliorer l’efficacité et la sécurité du transport aérien. ---------- Abstract The flight envelope of civil aircraft is bounded by several complex aerodynamic phenomena, of those the stall and the buffeting of lifting surfaces. Stall is defined as a loss of lift caused by flow separation. This usually occurs at low speed and high angle of attack, thus influencing the take-off and landing performances. Buffeting is a vibration of the airframe caused by unsteady aerodynamic loads. More specifically, the transonic buffet phenomenon causes an oscillation of the shock position on an aircraft wing. It occurs at high speed, induces unsteady loads and vibrations, and limits the maximum speed of aircraft. Both phenomena are detrimental to aircraft handling and must be avoided. However, transonic buffet and stall remain misunderstood and difficult to predict. For these reasons, aircraft manufacturer use empirical relations to predict these flight envelope limits, which are later verified in flight tests. Improving the comprehension of these phenomena will allow to reduce the uncertainty on the buffet and stall margins in the design stages and thus to improve the performances of future aircraft. Low speed stall involves complex dynamics, of which the stall cells. The latter are a variation of the chordwise separation point from one spanwise section to the other. This is associated with a complex vortical flow structure which has a mushroom shape. Experiments on wings spanning from one wind tunnel wall to the other and idealized infinite wing numerical simulations, which are essentially two-dimensional configurations, show their occurrence. This points to a flow instability which causes the two-dimensional separated flow to become three-dimensional. On the other hand, transonic buffet is characterized by an oscillation of the shock wave position caused by an interaction between the shock wave and a shock induced flow separation. Flow visualizations show three-dimensional flow features in the form of a variation of the chordwise shock position between spanwise wing stations (i.e. buffet cells). These buffet cells have a mushroom shape similar to the stall cells, which leads to the idea of studying both phenomena with the same methods. This thesis aims to improve our understanding of stall and buffet by studying these cells. The primary hypothesis on which this work is based is that the stall cells and the buffet cells are in fact the same flow phenomenon in two different flow regimes. To assess this hypothesis, the behavior of the cells is first characterized using Unsteady Reynolds Averaged Navier-Stokes (URANS) simulations and Zonal Detached Eddy Simulations (ZDES). Infinite swept wing configurations are selected to exclude influence of the geometry and boundary conditions on the occurrence of three-dimensional flow features. Such configurations are achieved numerically by closing the computational domain with periodicity boundary conditions in the wing spanwise direction. Results from these simulations exhibit similar behavior between the stall cells and the buffet cells. Both are stationary for unswept wings and convected in the spanwise direction when the wings are swept. The convection speed is proportional to the magnitude of the projection of the free steam velocity along the leading edge direction. The number of buffet and stall cells increases with the span of the computational domain. Moreover, visualizations of the skin friction lines in both flow conditions exhibit a similar mushroom shape flow topology. In the simulation of the subsonic stall, fully converged steady solutions with stall cells are obtained on the unswept wings. In the case of the transonic buffet, the classic two-dimensional buffet instability is still present and causes the flow to be unsteady, even on unswept wings. However, the present study does not exclude the possibility to observe steady buffet cells on unswept wings. These results also allow to clearly distinguish between two dominant features of the three-dimensional buffet phenomenon. The first one is a two-dimensional buffet instability which occurs at a constant frequency, no matter the sweep angle. The second one is the convection of buffet cells at a frequency proportional to the sweep angle. This explains why the buffet of civil aircraft wings has a higher frequency than the two-dimensional buffet instability, since these wings are highly swept. In order to support the primary hypothesis and to work towards a model explaining the stall cell and buffet cell phenomena, they are also studied using global linear stability analyses about steady two-dimensional base-flows. These analyses allow to study the bifurcation from the two-dimensional base-flows towards an unsteady and/or a three-dimensional behavior, in the linear regime. The methodology used in this study assumes the instability will be periodic in the spanwise direction to reduce the computational cost. These analyses show that a globally unstable mode is responsible for a transition from the two-dimensional state towards a three-dimensional one. This mode is non-oscillatory for unswept wing and has a non-zero frequency for swept wing cases. This frequency is consistent with the one of flow structures convected at the speed observed in the URANS simulations. In the case of the subsonic stall, the wavelength of the stall cells is correctly predicted by the stability analyses and so is the frequency. In the case of the transonic buffet, the wavelengths found in the stability analyses do not match the one of the URANS simulations. However, it is found that the URANS simulations indeed diverge from the two-dimensional state with the wavelength of the stability analysis. Hence, nonlinear effects are probably responsible for this shift in the wavelength. Finally, it is observed that stall cells and buffet cells occur when the slope of the lift versus angle of attack curve of the two-dimensional steady base-flows becomes negative (i.e. onset of the stall). This is observed in the results of the stability analyses as well as in the URANS simulations where the flow stays spanwise invariant as long as this slope stays positive. These analyses are also carried out for a case at high Reynolds number and a threedimensional unstable mode is identified for every Mach number from the incompressible to the transonic regime. Moreover, a scaling of the spanwise wavelength of the mode by the separation thickness is successfully used to link the results from all these Mach numbers, at least near the onset of the three-dimensional instability. Finally, the behavior of a simple inviscid model (lifting surface) coupled with a viscous correction is studied to explain the stall cell phenomenon. This model is able to reproduce the main characteristics of the stall cell such as the number of cells over a wing of a given aspect ratio. This model predicts the onset of stall cells as soon as the slope of the lift versus angle of attack curve is negative. This model also allows to infer that the stall cells will only be observed for a trailing edge type of stall where the flow separation progressively moves from the trailing edge to the leading edge of the wing as the angle of attack is increased. This conclusion is made since a leading edge type of stall results in a sharp reduction of the lift in the post stall regime. Hence, the range of angles of attack for which the lift slope is negative is infinitely small and the entire wing will jump from a two-dimensional attached flow to a two-dimensional separated flow. Of course, this conclusion is for an idealized case and the flow behavior is expected to be more complex for real aircraft applications. An interesting point is the fact that the viscous effects are treated in a macroscopic way in the inviscid model. In fact, this model contains no information about the flow topology of the separated flow, it only takes into account the lift versus angle of attack polar of an airfoil. For this reason we can conclude that an instability in the lifting line or lifting surface theory can explain the stall cells and by extension the buffet cells since they have been shown to relate to similar instabilities. Together, these results shed light on the fact that the two-dimensional and three-dimensional buffet are two different phenomena, the latter being more akin to the stall cells than to a shock induced separation instability. They also emphasize that the stall/buffet cells are linked to an instability of the circulation distribution on a wing (i.e. lifting line theory). The ideas and methods put forward in the thesis can be used in the industry towards more accurate modeling of these phenomena, and to increase air transport efficiency and safety