Michel Costes

Improved Two-Dimensional Dynamic Stall Prediction with Structured and Hybrid Numerical Methods

A joint comprehensive validation activity on the structured numerical method elsA and the hybrid numerical method TAU was conducted with respect to dynamic stall applications. In order to improve two-dimensional prediction, the influence of several factors on the dynamic stall prediction were investigated. The validation was performed for three deep dynamic stall test cases of the rotor blade airfoil OA209 against experimental data from two-dimensional pitching airfoil experiments, covering low speed and high speed conditions. The requirements for spatial discretization and for temporal resolution in elsA and TAU are shown. The impact of turbulence modeling is discussed for a variety of turbulence models ranging from one-equation Spalart-Allmaras-type models to state-of-the-art seven-equation Reynolds stress models. The influence of the prediction of laminar/turbulent boundary layer transition on the numerical dynamic stall simulation is described. Results of both numerical methods are compared to allow conclusions to be drawn with respect to an improved prediction of dynamic stall.

Dynamic Stall Control Using Deployable Leading-Edge Vortex Generators

A new concept of active dynamic stall control is proposed, designed and experimentally tested on a OA209 airfoil model. The active control principle is based on leading-edge vortex generation in order to alleviate the dynamic stall vortex formed and convected at the leading-edge of an airfoil operating at a helicopter blade in fast forward flight. The active device aims to beused only during retreating blade side for dynamic stallflight conditions in order to avoid drag penalties on the advancing blade side. The designed actuator is a row of deployable vortex generators (DVGs) located at the leading-edge of the airfoil that fit the airfoil shape when retracted. Deployment is possible for different heights as well as different phases and frequencies with respect to the airfoil oscillation. The paper addresses the validation of the effectiveness of the devices to delay static stall and alleviate dynamic stall penalties. Results show a delay in static stall angle of attack of 3 deg and a reduction of negative pitchingmoment peak up to 60% for dynamic stall. The analysis of the experimental database indicates that different compromises between lift and pitchingmoment can be achieved depending on the phase actuation of the DVGs.

Zonal Detached-Eddy Simulation of Separated Flow Around a Finite-Span Wing

The unsteady turbulent flow around a three-dimensional finite-span wing at stall is investigated by means of a zonal detached-eddy simulation. The computation captures a nondelayed development of unsteady structures at the beginning of the separation and accurately resolves the energy cascade in the fully turbulent mixing layer over a large frequency range. A comparison with the experimental data shows that the zonal detached-eddy simulation approach clearly improves the prediction of the mean flow, whereas classical Reynolds-averaged Navier–Stokes methods usually fail. A spectral analysis of pressure signals highlights the origin and the characteristics of the various unsteady mechanisms involved in this complex flow.

Numerical simulation of a pitching airfoil under dynamic stall conditions including laminar/turbulent transition .

This paper describes CFD activities achieved by ONERA in the frame of an internal multidisciplinary project on dynamic stall. These activities concern the physical analysis of the dynamic stall. The comparisons between a detailed experimental flow field and the CFD solution s yield to a very good validation of the modeling and to a better physical understanding of the stall phenomenon. The simulations have been performed for a OA209 airfoil in dynamic s tall conditions by solving the RANS equations for unsteady flows . The influence of the laminar/turbulent trans ition process is also discussed. T o ensure the good quality of the solution s all the computations are performed by taking into account the conclusions given by previous work on the efficiency of numerical methods, the accuracy of turbulence models and the influence of the grid resolution .

Low-Mach-Number Preconditioning Applied to Turbulent Helicopter Fuselage Flowfield Computation

HIS paper deals with the problems encountered in simulating low-Mach-number e ows encountered in applied aerodynamics. Typical examples are e ows over external aerodynamic bodies such as helicopter fuselages, road vehicles, submarines, and multielement airfoils at high angles of attack. Low-speed e ows may be compressible because of surface heat transfer or volumetric heat addition. Classical numerical methods for the simulation of compressible viscous e ows using the Reynolds averaged Navier‐ Stokes (RANS)equationsperformsatisfactorilyinmoste owregimes(from mediumsubsonice owtohypersonice ow).However,inthelowsubsonic e ow regime, that is, for Mach numbers lower than about 0.1, these methods give poor results in terms of convergence rate and solution accuracy. The dife culty in solving the compressible equations for low Mach numbers is associated with the large ratios that exist between the acoustic and convective wave speeds. Two approaches are currently considered to solve this problem. The e rst is a perturbation method. It is based on a perturbed form of the equations where the physical acoustic waves are replaced

Improved Unsteady Far-Field Drag Breakdown Method and Application to Complex Cases

Far-field drag extraction has the advantage over near-field integration of providing a phenomenological breakdown of drag. A decomposition into components linked to shock waves, viscous interactions, and lift-induced vortices is straightforward for steady flows. A formulation based on thermodynamic considerations is used at ONERA–The French Aerospace Lab and by its partners, but it is restricted to steady cases. A generalization to unsteady flows of the Van der Vooren formulation has been developed and tested on three unsteady cases previously. The proposed method allowed the breakdown of drag into the three usual components only; however, the induced drag coefficient remained ill-defined. This unsteady formulation is here modified to better express the induced drag. A new drag component is identified as a propagation and acoustics contribution. The new formulation is then applied to complex cases: two-dimensional and three-dimensional pitching cases, and an OAT15A profile subject to buffet simulated by z...

Numerical Investigation of Three-Dimensional Static and Dynamic Stall on a Finite Wing

Three-dimensional numerical computations using ONERA’s structured elsA code and the unstructured DLR-TAU code are compared with the OA209 finite wing experiments in static stall and dynamic stall conditions at a Mach number of 0.16 and a Reynolds number of 1 × 10^6 . The DLR-TAU computations were run with the Spalart–Allmaras and Menter shear stress transport (SST) turbulence models, and the elsA computations were carried out using the Spalart–Allmaras and the k–ω Kok + SST turbulence models. Although comparable grids were used, the static simulations show large discrepancies in the stall region between the structured and unstructured approaches. Large differences for the three-dimensional dynamic stall case are obtained with the computations using the Spalart–Allmaras turbulence model showing trailing edge separation only in contrast to the leading edge stall in the experiment. The three-dimensional dynamic stall computations with the two- equation turbulence models are in good agreement with the unsteady pressure measurements and flow field visualizations of the experiment, but also show a shift in the stall angle compared to the experiment. The analysis of the flow field around the finite wing using the numerical simulations reveals the evolution of the -shaped vortex, generated by the interaction of the blade tip vortex.

Investigation of the unsteady flow development over a pitching airfoil by means of TR-PIV

The flow over an OA209 airfoil subjected to a sinusoidal pitching motion under dynamic stall conditions is investigated experimentally by means of time resolved particle image velocimetry (TR-PIV) and surface pressure measurements. Dynamic stall is distinguished by the formation and convection of large scale coherent structures and a delay in massive flow separation. A vortex detection scheme based on an identification function derived directly from the velocity fields is adopted to identify vortex cores. The combination of global time resolved imaging and an automated vortex identification algorithm allows for the investigation of the spatial and temporal evolution of vortical structures within a single oscillation. Furthermore, the mechanisms associated with the dynamic stall delay are considered.

Zonal Detached-Eddy Simulation (ZDES) of the Three-dimensional Stalled Flow around a Finite Span Wing

A Zonal Detached-Eddy Simulation of the ow around a 3D nite span wing at stall is investigated. The time-averaged ow is in good agreement with wall pressure and Particule Image Velocimetry measurements. In particular, the massive ow separation on the suction side of the wing and the ow reattachement induced by the tip vortex is correctly captured while classical Reynolds-Averaged Navier-Stokes method fails. A more detailed analysis of the unsteady ow structures highlights the large scale range involved in this complex ow.

Comparison Between Two-Dimensional and Three-Dimensional Dynamic Stall

Numerical computations using the DLR-TAU code investigate the differences and similarities between dynamic stall on the two-dimensional OA209 airfoil and the three-dimensional OA209 finite wing. The mean angle of attack in the two-dimensional computations is reduced to match the effective angle of attack at the spanwise position where in the finite wing computations the dynamic stall vortex starts to evolve. Small variations of the mean angle of attack in the two-dimensional numerical simulations show a change from trailing edge separation only to deep dynamic stall. The analysis of the three-dimensional flow field reveals that after the evolution of the dynamic stall vortex the flow field is split into two parts: 1. High spanwise velocities towards the wing’s root in the region between the plane of the first occurrence of stall and the wing’s root. 2. High spanwise velocities towards the wing’s tip in the region between the plane of the first occurrence of stall and wing tip.