The effect of Gurney flaps on twodimensional airfoils, three-dimensional wings, and reflection plane model was investigated. There have been a number of studies on Gurney flaps in recent years. However, these studies have been limited to two-dimensional airfoil sections. A comprehensive investigation on the effect of Gurney flaps for a wide range of configurations and test conditions was conducted at Wichita State University. A symmetric NACA 0011 and a cambered GA(W)-2 airfoils were used during the single element airfoil part of this investigation. The GA(W)-2 airfoil was also used during the two-element airfoil study with its 25% chord slotted flap deflected at 10, 20, and 30 degrees. Straight and tapered reflection plane wings with Natural Laminar Flow (NLF) airfoil sections were tested for the three-dimensional wing part of this investigation. A fuselage and engine were attached to the tapered NLF wing for the reflection plane model investigation. Compared to the baseline clean configuration, the Gurney flap improved the maximum lift coefficient. However, there was a drag penalty associated with this lift increase. c Cd Q cm Nomenclature Chord length Drag coefficient Lift coefficient Pitching moment coefficient, quarter or half chord as indicated Freestream dynamic pressure, Uw Freestream velocity x,y,z Streamwise, spanwise, normal directions a Angle of attack d Flap deflection p Freestream density Introduction The Gurney flap is a short flat plate attached to the trailing edge perpendicular to the chordline on the pressure side of the airfoil. Race car driver Dan Gurney used this flap to increase the down force and thus the traction generated by the inverted wings on his race cars. Field tests by Gurney found that the flap increased the lift (i.e., traction) while the drag was slightly decreased. Increasing the Gurney flap height beyond 2% of chord continued to Increase the lift, but at the cost of substantially increased drag. Numerous wind tunnel tests on Gurney flaps have been conducted on both single and multi-element airfoils (see Giguere et afi for an extensive list). Liebeck found that the lift was increased when a Gurney flap was attached onto a Newman airfoil. Tuft flow visualization during the experiment indicated a downward turning of the flow behind the Gurney flap. Dye flow visualization on a NACA 0012 airfoil by Neuhart and Pendergraft also showed a downward turning of the flow behind the Gurney flap. Airfoil pressure distribution measurements were taken by Robert McGhee on an advanced technology airfoil. He found that the Gurney flap produced an overall decrease in pressure on the upper surface and an overall increase in 'Assistant Professor, Senior Member AIAA. 1 Associate Professor, Member AIAA. ^Graduate Assistant, Student Member AIAA. Copyright ® 1997 by Roy Myose, Michael Papadakis, and Ismael Heron. American Institute of Aeronautics and Astronautics, Inc. with permission. Published by the
A series of experiments was conducted at Wichita State University to study the effect on the vortex burst position when a von Karman vortex street from a small cylinder was impinged upon a 70-degree sweep-back delta wing. Phase comparisons between different experimental runs were accomplished, to a limited degree, by controlling the circular cylinder’s vortex shedding. A small fin was attached on the cylinder’s downstream side, and the cylinder-fin arrangement was rotated at a frequency equal to the cylinder’s natural shedding frequency. The start of the delta wing’s pitch-up was then synchronized with the fin’s rotational position. Dye flow visualization showed that the vortex burst position appeared to jump forward towards the apex and then moved gradually back toward the trailing edge in sync with the passage of the von Karman vortices. The range of forward to rear-most variation in the burst position was about 10 to 30% of chord at an angle of attack of 35 degrees, and diminished to about 5 to 10% of chord at higher angles of attack.
The effect of a Gurney flap in a compressor cascade model at low Reynolds number was investigated using tuft flow visualization in a water table facility. Although small in scale, water tables have the advantage of low cost and the ease with which test conditions can be varied. In this experiment, tuft flow visualization was used to determine the outgoing flow angle for a NACA 65-(12)10 compressor cascade model with a solidity of 1.5 at a blade chord Reynolds number of 16,000. The baseline (no flap) results were found to be in good agreement compared to results in the literature for tests conducted at Reynolds number in the 250,000 + range. A second set of measurements were then taken for a Gurney flap with a height of 2% of the chord length attached to the trailing edge of the cascade blades. The results suggest that the Gurney flap energizes the flow and delays the stall at large incoming flow angles. Nomenclature c = chord length Re C = Reynolds number based on chord length, Uc/ν U = freestream velocity y = offset distance in the stagger direction βin = incoming flow angle, between the in-flow direction and a line perpendicular to the stagger line βout = outgoing flow angle, between the out-flow direction and a line perpendicular to the stagger line λ = stagger angle, between the chord line and a line perpendicular to the stagger line ν = kinematic viscosity σ = solidity of cascade, c/y
A series of experiments on a 70-degree delta wing was conducted at Wichita State University to study the effect on the vortex burst position when simultaneous dynamic pitch and unsteady freestream velocity were used. The aim was to better understand the relationship between the freestream velocity and the time constants involved in the movement of the vortex burst point. Experiments indicated that a change in the freestream velocity resulted in a momentary pause in the forward progression of the vortex burst of the leading-edge vortex.
A series of experiments was performed at Wichita State University's water tunnel on a 70-degree-sweep delta wing using a towing mount. A video camera captured dye-flow visualization images of the vortex burst that were subsequently analyzed using a computer-assisted image analysis software. The aim was to better understand the relationship between the freestream velocity and the time constants involved in the movement of the vortex-burst point. Experiments indicated that a change in the freestream velocity changed the forward progression of the vortex burst. Under pitch-up conditions, deceleration resulted in a momentary retardation in the forward progression ofthe burst, whereas acceleration resulted in a faster progression toward the apex.
A series of experiments was conducted at Wichita State University to study the effect of support stem on the vortex burst position of a 70-degree sweepback delta wing. It is well known that a fighter aircraft’s performance at high angles of attack is greatly influenced by the development of leading edge vortices on a delta-shaped wing. The present investigation was motivated by a desire to understand how the design of specific model support structures can affect the delta wing vortex burst behavior. Results indicate that there is a slight influence on the burst location even if the support stem is located aft of the wing on the pressure side.
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