Aerodynamic Characteristics Of Wing Research Articles

Asymptotic theory of flow separation from wings of low aspect ratio

Numerous methods have been developed to calculate the aerodynamic characteristics of wings of low aspect ratio in the case when there is flow separation from the wing edges. Among the methods based on direct solution of the three-dimensional Euler equations there are the method of discrete vortices [1, 2] and the panel method [3]. In addition, numerical and asymptotic methods [4, 5] based on the theory of slender bodies [6] are used. One of the most important shortcomings of this theory is the dependence of the flow pattern at a given section of the wing on only the upstream flow. The obtained solutions therefore contain no information about the influence of the trailing edge of the wing, on which, as is well known, the Chaplygin-Zhukovskii condition is satisfied. The aim of the present paper is to construct an asymptotic theory of higher approximation and a corresponding numerical method for calculating flow separation from wings of low aspect ratio in which this shortcoming is absent.

Low-speed aerodynamic characteristics of wings with sweep discontinuities

It has been shown that, using the concept of a discrete vortex filament shedding, Weissinger’s method can be modified to determine the span-wise loading for wings with a sweep discontinuity. The sweep discontinuity is modelled by the introduction of a discontinuity in the equivalent vortex pattern chosen to replace the wing. The total lift, induced drag, pitching moment coefficients and span-wise location of the centre of pressure have been found for a few wing geometries using De Young and Harper’s procedure. An increase in the outboard sweep shifts the maximum load position outboard besides decreasing the overall lift and induced drag coefficients. High suction peaks appear at the pivot.

Aircraft configuration optimization for ground attack mission

The optimum aircraft wing configuration for a low-level ground attack mission is investigated. Wing planform variables, wing thickness ratio and wing loading were optimized to achieve the maximum radius of action for a given amount of fuel and armament, assuming fuselage and empennage geometry and engine to be frozen, using a modular interactive computer program. The iterative program is designed to estimate wing geometrical quantities from configuration variables, evaluate wing aerodynamic characteristics, estimate structural weight, compute point performance parameters and adjust design parameters until an optimum is reached. For a given starting configuration, the program is found to provide an optimal configuration which increases the radius of action by 5 nautical miles without violating preset constraints on design variables, take-off distance or all-up weight. The inclusion of design variables characterizing the fuselage and empennage and further constraints into the procedure are under investigation

Effect of Forward Acceleration on Aerodynamic Characteristics of Wings

Effect of Forward Acceleration on Aerodynamic Characteristics of Wings

On the Calculation of Laminar and Turbulent Boundary Layers on Longitudinally Curved Surfaces

Chakrabartty, S. K., Aerodynamic Characteristics of Wings of Arbitrary Planform, To appear in Journal of Aeronautical Society of India. Corsiglia, V. R. and Koenig, D. G., Large Scale Wind Tunnel Tests on a Low Aspect Ratio Delta-Winged Model Equipped with Sharp Edged Strakes, NASA TN D-3621. Wentz, W. H. and McMahon, M. C., Experimental Investigations of the Flowfield about Delta and Double Delta Wings at Low Speeds, NASACR-521. Sacks, A. H. and Burnell, J. A., On the Use of Impact Theory for Slender Configurations Exhibiting Flow Separation, Vidya 92, Itek Corp., March 1963. with

Calculation of nonlinear aerodynamic characteristics of wing of complex planform allowing for nose vortex sheet

Calculation of nonlinear aerodynamic characteristics of wing of complex planform allowing for nose vortex sheet

Calculation of the unsteady-state aerodynamic characteristics of three-dimensional bearing systems

At the present times effective numerical methods have been developed for solving both steady-state and unsteady-state linear problems on the determination of the aerodynamic characteristics of wings of complex form, in a plan view [1–3]. The transition to three-dimensional systems requires the development of new methods. There are a number of known investigations in this direction, for simple three-dimensional configurations [4–7]. The present article proposes a method for solving supersonic problems of the determination of the steady-state and unsteady-state three-dimensional bearing systems. It is a development of known methods for calculating wings of complex form in a plan view. The most effective route is one based on a solution of the problem for stepwise dependences on the time. After this, the transition to any other laws is effected using a “packet” integral.

Nonlinear characteristics of a thin wing in a shock-free flow around the nose

An effective means of controlling wing leading-edge stall at high angles of attack is deflection of the nose in order to assure shock-free entrance of the stream. A numerical method of computing the angles of nose deflection and the aerodynamic characteristics of a thin wing of arbitrary planform for a shock-free entrance of the steady ideal incompressible fluid stream is elucidated in this paper on the basis of nonlinear wing theory [1]. The problem is solved by the method of discrete vortices. In the computations, the wing and its wake, replaced by a vortex sheet, are modeled by a system of discrete vortices which are nonlinear segments with constant circulation along the length. The angles of deflection of the nose and the aerodynamic characteristics of the wing, including shunting of the free vortices shed from the side and trailing edges, are determined during the computation. Examples of an electronic digital computer are presented.

Comment on " Convergence Proof of Discrete-Panel Wing Loading Theories"

maximum tangential (swirl) velocities in the flow. Based on their own study of mass injection and the results of other researchers, Mason and Marchman2 concluded that the introduction of turbulence into the vortex by virtually any means is the dominant factor in tangential velocity reduction and the spreading of the core. The present research seems to substantiate this conclusion. These results show that vortex dissipation effects similar to those produced by mass injection can be realized without either bleeding power from the engines or using auxiliary power. In fact, using the spinning blade concept, the vortex dissipator may actually be used to produce power. Of course, such results are never free and the effects of these modifications on the wing aerodynamic characteristics as well as the vortex must be studied. The present study seems to indicate that such modifications can significantly reduce the vortex Lazard to trailing aircraft without the use of auxiliary or other power and with few, if any, detrimental effects on the wing aerodynamics.

Method for calculating aerodynamic characteristics of wings of complex planform with subsonic leading and trailing edges

Method for calculating aerodynamic characteristics of wings of complex planform with subsonic leading and trailing edges

A case of analogy between the unsteady aerodynamic characteristics of wings

The problems of harmonic oscillations and aperiodic wing motions are considered in the linear formulation. On the basis of the representation of the aerodynamic characteristics in terms of the rotary derivative coefficients it is shown that these coefficients are the same for the case of very low oscillation frequencies and the case of a linear variation of the parameters for aperiodic motion.

Rapid estimation of wing aerodynamic characteristics for minimum induced drag.

Rapid estimation of wing aerodynamic characteristics for minimum induced drag.

Determination of Elastic Wing Aerodynamic Characteristics

The basic aeroelastic integral equation is formulated by treating the wing as a tapered cantilever beam mounted normal to its straight elastic axis. By means of an iterative process, a series solution is obtained for the aeroelastic change in angle-of-attack distribution. For sweptback wings, this Neumann solution fails to converge at dynamic pressures that exceed the absolute value of the divergence dynamic pressure, even though the latter is negative and the wing physically stable. This eigenvalue problem is handled by expanding the series about a new point and thus effectively increasing the radius of convergence to agree with the physical case. I t is shown that the new point, expressed by a convergence factor, can be chosen so as to improve the rate of convergence of the series. The aeroelastic effects on the stability derivatives, on the hinge moments, and on the effectiveness of the control surfaces are investigated and shown to be expressible as simple functions of the dynamic pressure. Wind-tunnel data for a full-scale sweptback wing are presented and agree well with results calculated using the present method.