Showing papers by "M. S. Chandrasekhara published in 1994"

Quantitative study of unsteady compressible flow on an oscillating airfoil

Abstract: Detailed interferometric measurements of the flow near the leading edge of an oscillating airfoil offer the first detailed experimental quantification of the locally compressible flow field that surrounds an oscillating airfoil at moderate subsonic Mach numbers. Interferograms obtained by a specially adapted real-time point-diffraction interferometry technique have revealed significant characteristics of this complex, and very rapidly varying, locally supersonic flow. Instantaneous pressure distributions determined from these interferograms document the effect of unsteadiness on the leading-edge flow environment.

Reattachment studies of an oscillating airfoil dynamic stall flowfield

Abstract: The reattaching flow over an oscillating airfoil executing large amplitude sinusoidal motion around a mean angle of attack of 10 degrees has been studied using the techniques of stroboscopic schlieren, two component laser Doppler velocimetry and point diffraction interferometry, for a free stream Mach number of 0.3 and a reduced frequency of 0.05. The results show that the dynamically stalled flow reattaches in a process that begins when the airfoil is very close to the static stall angle on its downward stroke and progresses over the airfoil through a large range of angles of attack as the airfoil decreases to about 6 degrees. The airfoil suction peak shows a dramatic rise as the static stall angle is approached and the velocity profiles develop such that the flow near the surface is accelerated. The process completes through the disappearance of a separation bubble that forms over the airfoil.

A Study of Dynamic Stall Using Real Time Interferometry

Abstract: Dynamic stall over an oscillating airfoil in compressible flow was studied using a real-time interferometry technique. Instantaneous flow field data was obtained for various unsteady as well as steady flow conditions. Comparison of steady flow interferograms with those taken in unsteady flow reveal a significant delay in the development of leading edge suction peaks in the unsteady case. The interferograms permit detailed analysis of the leading edge pressure field; as many as 13 pressure values have been obtained around the leading edge in the first 1 percent of the airfoil chord. The results offer a significant new insight into the character of the dynamic stall vortex, and the stall delay that is observed during dynamic motions.

An Assessment of the Effect of Compressibility on Dynamic Stall

Abstract: Compressibility plays a significant role in the development of separation on airfoils experiencing unsteady motion, even at moderately compressible free-stream flow velocities. This effect can result in completely changed stall characteristics compared to those observed at incompressible speed, and can dramatically affect techniques used to control separation. There has been a significant effort in recent years directed toward better understanding; of this process, and its impact on possible techniques for control of separation in this complex environment. A review of existing research in this area will be presented, with emphasis on the physical mechanisms that play such an important role in the development of separation on airfoils. The increasing impact of compressibility on the stall process will be discussed as a function of free-stream Mach number, and an analysis of the changing flow physics will be presented. Examples of the effect of compressibility on dynamic stall will be selected from both recent and historical efforts by members of the aerospace community, as well as from the ongoing research program of the present authors. This will include a presentation of a sample of high speed filming of compressible dynamic stall which has recently been created using real-time interferometry.

Analysis of low Reynolds number airfoil flows

Abstract: 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 boundary- layer 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. HE prediction of steady, inviscid flows over aerodynamic configurations is performed routinely nowadays. The computation of flows with separation bubbles or of fully sep- arated flows, on the other hand, is still a very challenging problem. For many practical applications, the assumption of fully developed turbulent flow yields good predictions of the flowfield. In other circumstances, such as the leading-edge dynamic stall flow, this assumption is not valid and compu- tations of such flows need improved methods. A characteristic feature of the dynamic stall flow is the onset of compressibility effects at a very low freestream Mach num- ber of O.2. 1-2 In addition, at transitional Reynolds numbers, it has been shown3-4 that the dynamic stall events are closely governed by the formation of a laminar separation bubble and its subsequent bursting. In fact, the above-cited studies demonstrate that the failure of the separated shear layer to reattach initiates dynamic stall, leading to the formation of the dynamic stall vortex. Further complications arise when the locally supersonic flow forms shocks that interact with the local boundary layer. It is important to recognize that the scales of the flow are very small here and the flow physics is not very clear. It is obvious that an accurate computational study of the problem demands a proper modeling of the phys- ics of the local compressible flow. In an effort to reach this eventual goal, it is necessary to include the transition physics that plays a key role in the dynamic stall process. The study to be reported represents a step in this direction. It is well known that prediction of the transition point and the transition length in such strongly adverse pressure gradient driven flows with current methods is difficult and involves uncertainties. Although several methods are available, the engineering prediction of transition relies on empirical for- mulation for boundary-layer flows. Whereas these methods have been moderately successful in steady and subsonic flows,

Boundary layer tripping studies of compressible dynamic stall flow

Abstract: The challenging task of properly tripping the boundary layer of a leading-edge-stalling airfoil experiencing compressible dynamic stall at Reynolds numbers between 3.6 X 10 5 and 8.1 X 10 5 has been addressed. Real-time interferometry data of the flow over an oscillating airfoil have been obtained at freestream Mach numbers of 0.3 and 0.45. The airfoil was tripped by separately placing five different trips of varying lengths near the leading edge. The trip heights ranged from 40 to 175 μm. The resulting flow and airfoil performance were evaluated using the criteria of elimination of the laminar separation bubble that otherwise forms, delay of dynamic stall onset to higher angles of attack, and production of consistently higher suction peaks. Quantitative analysis of the interferograms showed that the laminar separation bubble was still present with the smallest trip and premature dynamic stall occurred with the largest trip. The right trip was determined to be a distributed roughness element extending from 0.5 to 3% chord. Its height was found to compare reasonably with the airfoil boundary-layer thickness at the dynamic stall vortex formation angle of attack, at a location slightly upstream of the vortex origin in the adverse pressure gradient region.