Double wake model for separated flows over airfoils

34th Aerospace Sciences Meeting and Exhibit

Comparisons are made between numerically and experimentally produced wake structures behind airfoils undergoing rapid, oscillatory plunging motions. Numerical simulations are performed using an unsteady panel code. Inviscid, incompressible ows about arbitrary moving airfoils are computed with the unsteady wake approximated by discrete point v ortices, tracked using a Lagrangian mesh scheme. Numerically computed results are visualized using an interactive, graphical-animation interface. Experimental data are obtained from a low-speed water tunnel. Two-color dye injection is used to visualize unsteady wake structures, and velocity data are acquired using laser doppler velocimetry. Comparisons of vortex location agree well with linear theory for low amplitude motions. For large amplitude, high frequency motions, results diverge from linear theory, but wakes from the two a pproaches compare well with each other, including highly non-linear, non-symmetric wakes obtained for high amplitude, high frequency motions. Computed velocity pro les and integrated thrust coe cients for both approaches agree well. Nomenclature c = c hord length C d = drag coe cient per unit span, D=(q 1 c) C l = lift coe cient per unit span, L=(q 1 c) C t = thrust coe cient per unit span, T= (q 1 c) D = drag per unit span h = plunge amplitude in terms of c k = reduced frequency, !c=V 1

University Of Khartoum Engineering Journal, 2011

Fluid flow around a NACA 4412 airfoil in a wind tunnel test section at Reynolds number of 3 x 106, based on the chord of the airfoil (150 mm) and free stream velocity (30 m/s), is considered. The study covers the boundary layers around the airfoil and the wake region at different angles of attack. Different turbulence models are used to predict separated flows over the airfoil. Two-equation turbulence models, k-ω and k-ε, and Reynolds Stress Model are tested for the ability to predict boundary layer separation on the airfoil. Reynolds Stress model captured the physics of separated flow favourably, and gave a very realistic evolution of the vortex formed due to separation. Statistics of the flow which is generated by RSM are in good agreement with an existing wind tunnel experimental data. The flow field which is generated by the two-equation turbulence models poorly predicted flow separation and vortex dynamics and consequently overestimated the lift coefficient for angles of attack larger than the critical angle of attack.

Wind Energy, 2009

Direct numerical simulations of the Navier-Stokes equations are performed to achieve a better understanding of the behaviour of wakes generated by wind turbines. The simulations are performed by combining the in-house developed computer code EllipSys3D with the actuator-line methodology. In the actuator-line method, the blades are represented by lines along which body forces representing the loading are introduced. The body forces are determined by computing local angles of attack and using tabulated aerofoil coefficients. The advantage of using the actuator-line technique is that it is not needed to resolve blade boundary layers and instead the computational resources are devoted to simulating the dynamics of the flow structures. In the present study, approximately 5 million mesh points are used to resolve the wake structure in a 120-degree domain behind the turbine. The results from the computational fluid dynamics (CFD) simulations are evaluated and the downstream evolution of the velocity field is depicted. Special interest is given to the structure and position of the tip vortices. Further, the circulation from the wake flow field is computed and compared to the distribution of circulation on the blades.

Journal of aircraft, 2011

The presented investigation includes a combined experimental-numerical approach to quantify the wake vortex system of a high-agility aircraft from the near field up to the far field. Detailed near-field data are obtained by low-speed wind-tunnel tests on a delta-canard configuration of 1:15 scale. The measurements are performed at several angles of attack applying advanced hot-wire anemometry. For a wake distance of up to 16 wing spans, mean and turbulent velocity fields are measured. The upstream data are used to initialize implicit large-eddy simulations aimed to compute the velocity fields of the wake vortex system over a wake distance of up to 50 spans. Here, a validation case is shown comparing measured and calculated wake data over a distance from 4 to 16 spans, with the implicit large-eddy simulations initialized by the measured quantities at a position of two wing spans. Compared with the experimental data, the numerical results show the expected lateral and vertical movement of the wake vortex system due to the interaction of the single vortices. The distributions of axial vorticity, crossflow velocities, and turbulence intensities match well with the experimental data. In addition, the dissipation process can be observed, resulting in a reduction of circulation. In the context of this study, the measured and computed velocity fields will be used to determine unsteady aerodynamic loads acting on a fighter aircraft encountering the wake. This is of great importance as wake induction may result in critical structural dynamic loads.

2007

Abstract The two-dimensional Unsteady Reynolds Averaged Navier Stokes (URANS) equations have been solved in order to analyse the wake of a blunt airfoil. Turbulence closure was obtained by using the standard twoequation k��� �� model and the computed results were compared with experimental data. Only the flow over the trailing edge region was simulated in this study along with the resulting wake flow. A simplified flow domain was constructed using an empirical relation for boundary layer growth on airfoils.

2008

An experimental measurements of unsteady wakes behind a sinusoidally plunging airfoil was surveyed with a Hot-wire rake containing seven I-wires. The aim of these measurements was to study the velocity profiles behind the oscillating airfoil trailing edge. Influences of reduced frequency and angle of attack were studied in details. It was shown that the angle of attack and reduced frequency are the most important parameters which influence on the velocity profiles. The momentum deficit was relative to reduced frequency and amplitude of oscillation and its behaviours were changed after static stall angles. At the angle of attack beyond static stall angle flow separation causes increasing the momentum deficit. The velocity profile has approximately symmetric shape profile. Data were taken at mean incidence angles of 0, 8 degrees at reduced frequencies of 0.09 to 0.56. The nondimensional amplitude of oscillation was 0.27. The corresponding Reynolds numbers, based on the chord length, w...

2009

A research on studying the effect ofthe wake generated by an aircraft wing section on a following aircraft wing section is performed. NACA 2412 is chosen as the airfoil model in this project. The airfoil model is fabricated by using CNC machining. Aluminium is chosen as the material for the airfoil model because it provides a good surface finish. Three experiments are conducted by using the open-circuit wind tunnel in Universiti Sains Malaysia (USM). The wind tunnel tests were carried out at the velocity of 5m/s to 30m/s in a test section of the size 0.30m (W), 0.30m (H) and 0.60m (L), at the Reynolds number of4.10 x 104 to 2.54 x 105. Experiment 1 isthe testing of single airfoil model to define the coefficientof lift, coefficient of drag and Reynolds number at various angles of attack. Experiment 1 is functioning as references for comparison for Experiment 2 and Experiment 3. Experiment 2 and Experiment 3 are the testing of two airfoil models at a separating distance of 1 chord length (13cm) and 2 Chord lengths (26cm) respectively. The main objective of Experiment 2 and Experiment 3 are to study the effects of wake turbulence on the characteristic of coefficient of lift, coefficient of drag and Reynolds number of a following airfoil model when an airfoil model is placed in front of it at a specific distance. The characteristics of the wake generated by an airfoil model on a following airfoil model are observed and studied during the testing in wind tunnel. Further investigations, discussions and conclusions are carried out throughout completing this research project. The results show that the separating distance between the two airfoils affects the coefficient of lift, coefficient of drag and the stall angle of the following airfoil at various angle of attack and free stream velocity. ACKNOWLEDGEMENTS First and foremost, I would liketo express my utmost gratitude to Dr. Ahmed Maher Said AH, the project supervisor for my Final Year Project. He has been very supportive during the entire project period. He spent very much time and energy to guide me throughout these two semesters despite his other commitments and packed schedule as a lecturer in Universiti Teknologi Petronas. Under his constant supervision, I managed to complete my project according to the timeframe scheduled.

This paper carries out numerical study of the flow separations around NACA 0012 airfoil at large angles of attack. Flow separation introduces two major effects: sudden loss of lift and generation of aerodynamic noise. These two factors are highly concerned on the aircraft designation. This study gives a detail picture of flow separation. As the problems caused by flow separation are complicated, the spatial and temporal complexity makes it difficult to access by conventional experiment methods. In the presented work, the numerical investigation results from solving the time-dependent Navier-Stockes equations in the generalized curvilinear coordinates. Using a fourth order centered compact scheme for spatial discretization facilitates high resolution of the flow field, which will be neglected if using low-order numerical schemes. To avoid possible non-physical wave reflection, the non-reflecting boundary conditions are used at far-field and outlet boundary. Complex flow separation, vortex shedding, vortex merging, and vortex paring are observed in the computational results. The main purpose of this paper is to provide more detailed information of the flow separation.

European Journal of Mechanics B-fluids, 2016

The objective of this work is to investigate numerically the different physical mechanisms of the transition to turbulence of a separated boundary-layer flow over an airfoil at low angle of attack. In this study, the spectral elements code Nek5000 is used to simulate the flow over a SD7003 wing section at an angle of attack of α = 4 •. Several laminar cases are first studied from Re = 2000 to Re = 10000, and a gradual increase of the Reynolds number is then performed in order to investigate one transitional case at Re = 20000. Computations are compared with measurements where the instability mechanisms in the separated zone and near wake zone have been analyzed. The mechanism of transition is investigated, where the DMD (Dynamic Mode Decomposition) is used in order to extract the main physical modes of the flow and to highlight the interaction between the transition and the wake flow. The results suggest that the transition process appears to be physically independent of the wake flow, while the LSB shedding process is locked-in with the von Kármán instability and acts as a sub-harmonic.

Experiments in Fluids, 2018

Time-resolved particle image velocimetry (TR-PIV) and hydrogen bubble visualization are employed to study the effects of Reynolds number on the wake/shear layer interactions over multi-element airfoil (30P30N). The Reynolds number based on the stowed chord length (Re c) ranges from 9.3 × 10 3 to 3.05 × 10 4. According to the variation of dominated flow structures, a critical Re c interval from 1.27 × 10 4 to 1.38 × 10 4 is found, which is novel for the low-Reynolds-number flow over multielement airfoil. The slat wakes can be divided into two types by this critical interval. When Re c is smaller than this critical interval, no roll-up occurs to the shear layer of slat cusp. Görtler vortices generated by a virtual curved wall dominate the slat wake. When Re c is larger than this critical interval, roll-ups occur to the shear layer of slat cusp, which is similar to the cases at high Reynolds number (Re c ~ 10 6). These roll-ups and their evolution result in the coexistence of spanwise vortices and streamwise vortices in the slat wake. Different kinds of slat wake result in different kinds of wake/shear layer interactions above the main element. The flow physics behind these complex interactions, especially the novel flow structures and their evolution, is analyzed in detail to contribute to the fundamental research of wake/shear layer interactions. When Görtler vortices dominate the slat wake, they could trigger streaky structures within the leading-edge separated shear layer of the main element. When spanwise vortices and streamwise vortices co-exist in the slat wake, novel spanwise "double secondary vortices" are triggered above the main element by the spanwise vortices of slat cusp shear layer.