Surface pressure measurements on a pitching swept airfoil

The unsteady flow development over a pitching, backward-swept airfoil is described for a range of sweep angles and dimensionless pitch rates. Pressure measurements at six spanwise locations, taken at different angles of attack during the pitch-up motion of the airfoil, were analyzed to understand the effects of cross flow on the unsteady flow development. Vortexpropagation patterns over the suction surface were identified, together with changes in these patterns with sweep angle. Results show that for 0° sweep, detachment of the dynamic-stall vortex first occurs near the root, and that the vortex convection velocity is higher in this region. Pinning of the stall vortex delays unsteady flow events and detachment processes in the tip region. Vortex detachment and propagation progress in the chordwise direction in an oblique fashion with respect to the leading edge of the airfoil, with the largest vortex-propagation velocity near the root and the smallest near the tip. Propagation patterns of the stall vortex differ significantly as the sweep angle increases. At 15° sweep, vortex detachment first occurs in the mid-span region. The vortex propagation velocity is higher in the mid-span region than the root and tip regions. With increase in the sweep angle to 30°, vortex detachment first occurs near the wing tip and the vortex front progresses obliquely, with the slowest convection velocity near the root. The vortex detachment and propagation patterns at 30° sweep are the reverse of those at 0° sweep. As the sweep angle increases, the vortex-propagation velocity near the root decreases. Vorticity-flux calculations show that the vorticity source which feeds the dynamic-stall vortex is weaker in regions where vortex detachment occurs first, and the vorticity flux reaches a maximum magnitude at an earlier angle of attack compared to other spanwise locations. Pitch rate effects on the flow patterns are also described. c CL C,M Nomenclature airfoil chord length sectional lift coefficient sectional moment coefficient Cp pressure coefficient, 2(p pm)/(p U») p surface pressure p,,., reference static pressure Rec chord Reynolds number, U«,(cosA)c/v tp time period for pitch-up motion through an angle Act U^ free-stream velocity x chordwise distance along airfoil z spanwise distance along airfoil a angle of attack Act change in angle of attack during pitch up ct dimensionless pitch rate, (Aotc)/( tp UMcosA) p fluid density v fluid kinematic viscosity A sweep angle Introduction and Background The study of unsteady flow over pitching airfoils has been motivated largely by the need to understand helicopter rotor aerodynamics and by interest in supermaneuverable aircraft. The flow is dominated by a coherent vortical structure, the dynamic-stall vortex (DSV). Understanding the development of this vortex, especially in the incipient stages, is essential for the control of this unsteady flow. Detachment of the DSV from the airfoil surface leads to dynamic stall. Central to the flow development are a complex viscous-inviscid interaction, unsteady flow separation, and vortex dynamics. The difference in viscous and inviscid time scales due to the airfoil motion results in a delay of flow separation and in a transient augmentation of aerodynamic lift. Recent studies at IIT and elsewhere have revealed important information about the mechanisms involved in the initiation, development, and shedding of this vortex over two-dimensional airfoils. However, unsteady three-dimensional separation is not completely understood, and the aerodynamic theories and methods available to predict the development of these flows are few, and applicable only for a particular body shape rather than arbitrary three-dimensional configurations. Most of the published studies are for flows over delta wings and * Graduate Research Assistant, Fluid Dynamics Research Center. Student Member AIAA Professor, Fluid Dynamics Research Center. Associate Fellow, AIAA Copyright © by M. Ahsanul Karim and Mukund Acharya. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc. conical bodies. Some experimental investigations have been carried out recently to study the flow associated with pitching or oscillating rectangular wings of finite aspect-ratio'. Lorber used a model that pitched/oscillated about an axis connecting the root and tip quarter-chord, to simulate helicopter rotor blades. Patterson et al. examined the unsteady flow evolution over a pitching airfoil model swept back at an angle of 15". It is useful to conduct experiments that simulate the pitch-up maneuver of a fixed-wing aircraft for applications where the objective is to achieve post-stall flight maneuver and side slipping. With this in mind, the present set up was designed to pitch the airfoil model about a spanwise axis perpendicular to the oncoming flow, passing through the quarter-chord location at the airfoil root. The motion chosen was a large-amplitude pitch up of the airfoil, which is the appropriate airfoil motion to examine for such applications. Objectives The objective of this work was to develop an understanding of the physics of the flow over pitching swept wings, and provide information that would be useful in the design of a flow management or control system. A primary goal was to understand the effects of cross flow induced by wing sweep on the unsteady flow development through an analysis of surface-pressure and vorticity flux distributions, and to examine the effects of pitch rate on the flow. A description of the change in DSV propagation patterns at three different sweep angles, and the effect of cross flow on lift and moment coefficients are presented as part of the results. Another objective was to examine the effects of suction in the leading-edge region on the evolution of the flow. This will be discussed in a future paper. Experimental Arrangement Wind Tunnel and Airfoil Model The experiments were conducted in the Andrew Fejer Unsteady Wind Tunnel at IIT's Fluid Dynamics Research Center. This is a closed-circuit low-speed facility in which a flow velocity up to 40 m/s can be reached in the test-section with a turbulence level of 0.03%. The wind tunnel test section is 0.61m x 0.61m in cross-section and 3.1m in length. The hardware needed to impart a controlled motion to a model in the flow is available. For more details and a complete description of this facility, refer to Karim. The airfoil model used for this investigation had a symmetrical NACA 0012 profile with a chord length of 10 in. (25.4 cm) and an aspect ratio of 1.77. Experiments were carried out at sweep angles, A, of 0°, 15° and 30° (Fig. 1). The sweep angle was changed by using different mounting inserts. A shield was placed at the junction of the airfoil and the mounting attachment to prevent disturbances to the flow over the airfoil due to end-wall boundary layers. A cap rounded into an approximate half body of revolution was fitted to the wing tip. The airfoil was instrumented with pressure taps at six spanwise locations measured from wing root (z/c=22, 44.6, 88.6, 132.9, 154.2, and 168.2 %).