Showing papers on "Vortex lift published in 1972"

On the mechanism of dynamic stall.

Abstract: The sequence of events comprising dynamic stall of an airfoil is discussed, with emphasis on the role of the leading edge laminar separation bubble and shed vortex. A simple bubble model, based on a combination of theoretical and experimental investigations, is used to discuss the events prior to the shedding of the vortex, and provides the basis for a heuristic estimate of the delay in the occurrence of dynamic stall on a pitching airfoil. The evidence for the existence and dominant effect of the leading edge vortex on the dynamic stall required (but in most cases not presently available) for the prediction of the effects of stall on helicopter rotor blades are discussed. It is the intention of this paper to focus attention on the laminar separation bubble and the shed leading edge vortex as the dominant features of the dynamic stall mechanism in the hope of stimulating greater emphasis on these features in future dynamic stall research.

Aircraft wing with vortex generation

Abstract: Tail-less airplanes, such as dsclosed in U. S. Pat. No. 3,438,597 as a stall occurs or is about to occur, may be flown at very high angles of attack to generate favorable spanwise vorte, flows which augment the swept wing profiles creating resultant wing profiles having better lift characteristics, i.e., the vortexes created are lift generating. However, the aircraft is then uncomfortable to be in during such flights at very high angles of attack. Therefore to achieve the benefits of this lift generating vortex flow, without maneuvering such aircraft into a very high angle of attack, the swept wing is equipped with airfoil structures and accessories therefor, which are extended beyond the cruising speed contour of the swept wing, at lower speeds, to create spanwise vortex air flows which selectably enlarge the effective overall airfoil contours as sensed by the passing major air flows. Sustaining aerodynamic lift forces are thereby created at angles of attack well beyond the stall angles of the cruising airfoil contour, and by timely use of such airfoil structures, the vortex air flows are created soon enough for lift generating, so stable flight conditions may be created without so extensively altering the pitch of the landing and/or slow flying tailless aircraft.

Application of the leading edge suction analogy to prediction of longitudinal load distribution and pitching moments for sharp edged delta wings

Abstract: The leading-edge-suction analogy of Polhamus has been used to develop the longitudinal load distribution of the vortex lift for delta wings. This distribution is shown to be similar in shape to that of the potential-flow longitudinal loading for delta wings having aspect ratios of 2 or less. The totals of the two theoretical distributions for delta wings with an aspect ratio near 1 are in good agreement with the experimentally determined loadings over the angle-of-attack range from 0 to 30 deg. The corresponding predicted pitching moments show slightly more stability than those measured, because of loss of lift near the wing tips.

Low-speed aerodynamic characteristics of an airfoil optimized for maximum lift coefficient

Abstract: An investigation has been conducted in the Langley low-turbulence pressure tunnel to determine the two-dimensional characteristics of an airfoil optimized for maximum lift coefficient. The design maximum lift coefficient was 2.1 at a Reynolds number of 9.7 million. The airfoil with a smooth surface and with surface roughness was tested at angles of attack from 6 deg to 26 deg, Reynolds numbers (based on airfoil chord) from 2.0 million to 12.9 million, and Mach numbers from 0.10 to 0.35. The experimental results are compared with values predicted by theory. The experimental pressure distributions observed at angles of attack up to at least 12 deg were similar to the theoretical values except for a slight increase in the experimental upper-surface pressure coefficients forward of 26 percent chord and a more severe gradient just behind the minimum-pressure-coefficient location. The maximum lift coefficients were measured with the model surface smooth and, depending on test conditions, varied from 1.5 to 1.6 whereas the design value was 2.1.

Two-Dimensional Subsonic Wind Tunnel Investigations of a Cambered 30- Percent Thick Circulation Control Airfoil

Abstract: : A relatively thick Circulation Control (CC) elliptic airfoil section with thickness-to-chord ratio of 030 and a circular arc camber of 15 percent at the midchord was tested subsonically to determine its aerodynamic properties as a midspan blade section on a blown helicopter rotor The two-dimensional tests established the section's ability to generate the required lift at low and negative incidence Lift coefficients up to 65 were produced at moderate momentum coefficient C sub mu equal to or less than 024 High drag of the unblown bluff ellipse was greatly reduced by the application of very moderate blowing, and equivalent efficiencies of 47 (including power required for blowing) were generated at C sub l approximately equal to 19 The section's performance was found to be heavily influnced by upper and lower aft surface flow separations, especially at the larger positive and negative angles of attack In addition, both low Reynolds number and an increase in slot height were detrimental to section lift capability Nevertheless, the ability to operate at high lift coefficients essentially independent of angle of attack, and with large lift augmentation for relatively low blowing, promises to provide an effective blade section for heavy lift application

Wind Tunnel Tests of a Trapped Vortex-High Lift Airfoil.

Abstract: : Two dimensional low speed wind-tunnel tests have demonstrated the feasibility of employing trapped vortex diffusion on a high lift airfoil configuration. All 11 percent supercritical airfoil was selected as the base contour employing both leading and trailing edge flaps. The leading edge flap was 25 percent chord, drooped 30 degrees with the rear flap, 27 percent chord, deflected 15 degrees and 30 degrees. Trapped vortex cavities, affording boundary layer control with rapid diffusion, are located at each of the flap knees. Because of the high trapped vortex efficiency, blowing requirements are minimal. Section lift coefficients, in excess of 6, were achieved in the test. (Modified author abstract)

Supersonic lifting systems

Abstract: The present invention relates to a method and apparatus for reducing the drag, and thereby increasing the ratio of lift/drag of supersonic wings above that achievable with conventional swept wings. The invention is characterized by the fact that the drag reduction is achieved as a result of a mutual interaction among multiple lifting surfaces in a manner which results in lower drag than is achieved by a single lifting surface at the same lift.

Calculation of the lift on airfoils with separated boundary layers

Abstract: : The report contains a method for calculating the sectional lift coefficient Cl as a function of its angle of attack alpha and free stream Reynolds number Re even at large angles of attack beyond the maximum Cl. The results of the present theory are in agreement with all known experimental data, that is, (1) the theory shows that Cl increases with alpha until stall and then decreases as alpha is increased still further (keeping Re constant), (2) the theory shows that the Reynolds number effect on Cl versus alpha is insignificant at the lower angles of attack but becomes more pronounced near and beyond the stall where viscous effects and flow separation are more important, (3) the theory shows that Cl increases with Re at constant (high) values of alpha. (Author)