Interferometry and Computational Studies of an Oscillating Airfoil Compressible Dynamic Stall

Proc. of the 5th Asian Congress of Fluid Mechanics, Taejon City, Korea, Aug. 1992, Vol. II, pp. 1047 - 1050

Schlieren Studies of Compressibility Effects on Dynamic Stall of Airfoils in Transient Pitching Motion

Compressibility effects on the flowfield of an airfoil executing rapid transient pitching motion from 0 - 60 degrees over a wide range of Mach numbers and pitching rates were studied using a stroboscopic schlieren flow visualization technique. The studies have led to the first direct experiments] documentation of multiple shocks on the airfoil upper surface flow for certain conditions. Also, at low Mach numbers, additional coherent vortical structures were found to be present along with the dynamic stall vortex, whereas at higher Mach numbers, the flow was dominated by a single vortex. The delineating Mach number for significant compressibility effects was 0.3 and the dynamic stall process was accelerated by increasing the Mach number above that value. Increasing the pitch rate monotonically delayed stall to angles of attack as large as 27 degrees.AFOSR-MIPR-87-0029 and 88-0010NAVAIRAR

Computation of unsteady flows over airfoils

Two methods are described for calculating unsteady flows over rapidly pitching airfoils. The first method is based on an interactive scheme in which the inviscid flow is obtained by a panel method. The boundary layer flow is computed by an interactive method that makes use of the Hilbert integral to couple the solutions of the inviscid and viscous flow equations. The second method is based on the solution of the compressible Navier-Stokes equations. The solution of these equations is obtained with an approximately factorized numerical algorithm, and with single block or multiple grids which enable grid embedding to enhance the resolution at isolated flow regions. In addition, the attached flow region can be computed by the numerical solution of compressible boundary layer equations. Unsteady pressure distributions obtained with both methods are compared with available experimental data.Approved for public release; distribution is unlimited

Numerical Investigation of Passive and Active Control of Unsteady Compressible Flow

AIAA Paper 00-4417, Denver, CO, Aug. 2000.The article of record as published may be found at http://dx.doi.org/10.2514/6.2000-4417Numerical simulations of flow controls applied to airfoils undergoing oscillatory motion in compressible flow are presented. Currently available efficient and accurate numerical methods for the full compressible flow equations and advanced turbulence models are used for the numerical predictions. Time-accurate computations at low angles of incidence indicate development of flow unsteadiness for the slatted airfoil. It is found, however, that the leading-edge slat effectively suppresses dynamic stall, it improves dynamic blade response, and reduces flow separation, for compressible flow speeds. It was also found that dynamic stall can be suppressed using control with pulsating jets

A Zonal Method for Unsteady, Viscous, Compressible Airfoil Flows

The analysis and prediction of fluid-structure interaction for viscous, separated flows presents a great challenge to the aeroelastician. In this paper a zonal method for the computation of unsteady, viscous, separated flows over airfoils is presented. The flowfield is divided into a viscous inner zone, where higher grid resolution may be used, and an inviscid outer zone. Zonal grid solutions are presented for subsonic and transonic flows over a NACA-0012 airfoil subject to ramp and oscillatory motions. Transonic shock/boundary layer interaction and dynamic stall effects are encountered during the unsteady motion. The computed solutions are in good agreement with available experimental data

Analysis of Low Reynolds Number Airfoil Flows

(AIAA Paper 94-0534), Journal of Aircraft, Vol. 32, No. 3, May-Jun. 1995, pp. 625-630.Compressible steady and unsteady flowfields over a NACA 0012 airfoil at transitional Reynolds numbers are investigated. Comparisons with recently obtained experimental data are used to evaluate the ability of a numerical solution based on the compressible thin layer Navier-Stokes approximation, augmented with a transition model, to simulate transitional flow features. The discretization is obtained with an upwind-biased, factorized, iterative scheme. Transition onset is estimated using an empirical criterion based on the computed mean flow boundarylayer quantities. The transition length is computed from an empirical formula. The incorporation of transition modeling enables the prediction of the experimentally observed leading-edge separation bubbles. Results for steady airfoil flows at fixed angles of attack and for oscillating airfoils are presented