The onset and evolution of the dynamic stall vortex (DSV) are analysed by means of large eddy simulations of an SD7003 aerofoil undergoing periodic plunging motion in a transitional Reynolds number flow ( $Re =6\times 10^{4}$ ). Interactions between upstream propagating Kelvin–Helmholtz instabilities and a shear layer formed at the leading edge trigger flow separation. The former appear to be related to acoustic waves scattered at the trailing edge due to initial vortex shedding. Two freestream Mach numbers ( $M_{\infty }=0.1$ and $0.4$ ) are employed to examine the flow differences due to compressibility variations. The existence of a common timing for the acoustic perturbations in both flows suggests a possible Mach number invariance for the birth of the Kelvin–Helmholtz instability. Increasing compressibility, however, induces earlier spanwise fluctuations, higher flow three-dimensionality and a weaker and more diffuse DSV, which is formed further downstream of the leading edge and has lower residency time. In order to better characterize the onset of the DSV, two empirical criteria are assessed: the leading edge suction parameter and the chord-normal shear layer height. Results demonstrate a higher robustness of the latter with respect to Mach number variations. Modal decomposition, performed with both the classical dynamic mode decomposition (DMD) and its multi-resolution variant (mrDMD), highlights key trends and demonstrates the capacity of the mrDMD to extract physically meaningful flow structures related to the stall onset. Such detailed characterization of the shear layer can be used for a systematic exploration of flow control strategies for unsteady aerofoils.
This paper presents a study of the leading-edge noise radiated by an airfoil undergoing a turbulent inflow. The noise prediction of generic airfoil profiles subjected to spanwise-varying inflow conditions is performed with the support of Amiet's theory and the inverse strip technique. In the proposed methodology, the aeroacoustic transfer function of a generic airfoil profile is computed by the boundary element method. The effects of the airfoil leading-edge thickness on the inflow turbulence are accounted for by a turbulence spectrum based on the rapid distortion theory. This research shows that the turbulence distortion plays a significant role on the predicted noise levels. Compared with the von Kármán model for isotropic turbulence, the rapid distortion theory predicts reduced noise levels at high frequencies and increased noise levels at low frequencies. This paper also shows that the spanwise-varying inflow, here represented by a linearly changing condition, contributes to raising the noise levels when compared to the similar uniform inflow case. This research confirms that the finite airfoil thickness decreases the airfoil-gust lift response, consequently reducing the noise levels. This observation is more pronounced for microphones positioned downstream of the airfoil and for high frequencies.
Acoustic measurements of turbulent jets in the vicinity of a flat plate, mimicking a neighbouring wing, were compared to results from two wavepacket-based source models previously studied in the literature: the Tailored Green's Function method, which considers the radiation of the turbulent structure in the vicinity of a semi-infinite flat plate, and the Boundary Element Method, which can represent the full geometry of the plate used in the experiments. Particular interest is given to analysing how the angle of attack of the plate (α) affects the sound radiated by an installed jet with trailing edge 6 diameters away from the nozzle and 1 diameter away from the centerline for 0° ≤ α ≤ 45°. The results herein confirm the behaviour identified by the models: the scattered acoustic field follows the rotation of the plate, shifting a silence region with negligible scattered sound, and creating regions with lower noise levels in positions that correspond to the ground for an aircraft with engines under its wings. This is further explored by means of a Mach number analysis for M = 0.5, 0.7, and 0.9, showing that this trend is present whenever trailing-edge scattering of jet disturbances is dominant in the acoustic field.
We propose a numerical framework to compute the airfoil-gust lift response and its subsequent leading-edge noise generation due to an incident compressible turbulent flow. This approach is valid for blades with large aspect ratios, general airfoil geometries, threedimensional supercritical perturbations and compressible subsonic flows. The linearized equation for unsteady potential flow is rewritten as a Helmholtz equation in the transformed Prandtl-Glauert plane, leading to a boundary value problem prescribed by the linearized airfoil theory. The boundary element method is then employed iteratively to solve the Helmholtz equation for realistic airfoil configurations. Results show that non-zero thickness airfoils drastically reduce the pronounced acoustic radiation expected by oblique gusts. However, at moderate Mach numbers, the compressibility effects may increase the noise radiation in the upstream direction compared to Amiet's analytical solution when thickness is addressed into the analysis.
Large-eddy simulations (LES) are employed to investigate the pitch–plunge equivalence of an SD7003 airfoil undergoing constant ramp motions at Reynolds number [Formula: see text]. The equivalence is constructed based on the geometric effective angle of attack according to the quasi-steady thin-airfoil theory. Two rates of descent (or pitch up) are analyzed for different Mach numbers in order to investigate the effects of compressibility on the evolution of the dynamic stall vortex (DSV). During the onset of the DSV and its transport along the airfoil surface, remarkable similarities are found between pitch and plunge in terms of flow topology, aerodynamic loads, and signatures of wall pressure and friction coefficients. However, these flow similarities cease at high-load conditions as the DSV becomes more susceptible to the peculiarities of the airfoil motion, manifested here by different trailing-edge vortices. Employing a correction for the rotation-induced apparent camber effect present in the pitching case, which results from the quasi-steady thin-airfoil theory, improves the agreement between pitch and plunge. However, it is not sufficient to assimilate their disparate trailing-edge systems. Results also demonstrate that the limit angle at which pitch–plunge equivalence remains valid decreases for higher Mach numbers.
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