Effect of Leading-edge Suction on Vortices Over Delta Wings PDF Download
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Author: Scott Duncan McCormick
Publisher:
ISBN:
Category :
Languages : en
Pages : 114
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Book Description
Author: Scott Duncan McCormick
Publisher:
ISBN:
Category :
Languages : en
Pages : 114
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Book Description
Author: Edward C. Polhamus
Publisher:
ISBN:
Category : Delta wing airplanes
Languages : en
Pages : 18
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Book Description
Author: Brant H. Maines
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Book Description
Author: Edward Charles Polhamus
Publisher:
ISBN:
Category : Aerodynamics
Languages : en
Pages : 28
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Book Description
A concept for the calculation of the vortex lift of sharp-edge delta wings is presented and compared with experimental data. The concept is based on an analogy between the vortex lift and the leading-edge suction associated with the potential flow about the leading-edge. This concept, when combined with potential-flow theory modified to include the nonlinearities associated with the exact boundary condition and the loss of the lift component of the leading-edge suction, provides excellent prediction of the total lift for a wide range of delta wings up to angles of attack 20[degrees] or greater.
Author: Charles H. Fox
Publisher:
ISBN:
Category : Delta wing airplanes
Languages : en
Pages : 44
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Book Description
Author: William P. Henderson
Publisher:
ISBN:
Category : Airplanes
Languages : en
Pages : 60
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Author: H. Q. Yang
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Author: Q. Deng
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Book Description
Author: Melvin H. Snyder
Publisher:
ISBN:
Category : Aerodynamic load
Languages : en
Pages : 28
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Book Description
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.
Author: National Aeronautics and Space Adm Nasa
Publisher:
ISBN: 9781731351593
Category :
Languages : en
Pages : 162
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Book Description
The vortical flow over a delta wing contributes an important part of the lift - the so called nonlinear lift. Controlling this vortical flow with its favorable influence would enhance aircraft maneuverability at high angle of attack. Several previous studies have shown that control of the vortical flow field is possible through the use of blowing jets. The present experimental research studies vortical flow control by applying a new blowing scheme to the rounded leading edge of a delta wing; this blowing scheme is called Tangential Leading Edge Blowing (TLEB). Vortical flow response both to steady blowing and to unsteady blowing is investigated. It is found that TLEB can redevelop stable, strong vortices even in the post-stall angle of attack regime. Analysis of the steady data shows that the effect of leading edge blowing can be interpreted as an effective change in angle of attack. The examination of the fundamental time scales for vortical flow re-organization after the application of blowing for different initial states of the flow field is studied. Different time scales for flow re-organization are shown to depend upon the effective angle of attack. A faster response time can be achieved at angles of attack beyond stall by a suitable choice of the initial blowing momentum strength. Consequently, TLEB shows the potential of controlling the vortical flow over a wide range of angles of attack; i.e., in both for pre-stall and post-stall conditions. Lee, K. T. and Roberts, Leonard BLOWING; DELTA WINGS; FLOW DISTRIBUTION; LEADING EDGES; VORTICES; AIRCRAFT MANEUVERS; ANGLE OF ATTACK; MOMENTUM; NONLINEARITY; STABILITY...
