Aerodynamic improvement of a delta wing in combination with leading edge flaps

Recently, various studies of micro air vehicle(MAV) and unmanned air vehicle(UAV) have been reported from wide range points of view. The aim of this study is to research the aerodynamic improvement of delta wing in low Reynold's number region to develop an applicative these air vehicle. As an attractive tool in delta wing, leading edge flap(LEF) is employed to directly modify the strength and structure of vortices originating from the separation point along the leading edge. Various configurations of LEF such as drooping apex flap and upward deflected flap are used in combination to enhance the aerodynamic characteristics in the delta wing. The fluid force measurement by six component load cell and particle image velocimetry(PIV) analysis are performed as the experimental method. The relations between the aerodynamic superiority and the vortex behavior around the models are demonstrated.     

Navier–Stokes computation of transonic vortices over a round leading edge delta wing

A 3D Navier–Stokes solver has been developed to simulate laminar compressible flow over quadrilateral wings. The finite volume technique is employed for spatial discretization with a novel variant for the viscous fluxes. An explicit three-stage Runge–Kutta scheme is used for time integration, taking local time steps according to the linear stability condition derived for application to the Navier–Stokes equations. The code is applied to compute primary and secondary separation vortices at transonic speeds over a 65° swept delta wing with round leading edges and cropped tips. The results are compared with experimental data and Euler solutions, and Reynolds number effects are investigated.     

三角翼前缘涡的某些破裂形式及特性研究

基于流动显示和PIV技术测量的实验结果，对三角翼前缘涡破裂的一些形式和破裂特性进行了分析和讨论。通过PIV测量所得到的涡量分布证实了在螺旋破裂的情况下，涡核的螺旋方向与前缘涡的旋转方向相反，及双螺旋破裂形式的存在等。进而对螺旋波的形成机理提出了与有关文献不同的看法。     

Flow past a delta wing with a sinusoidal leading edge: near-surface topology and flow structure

The near-surface flow structure and topology on a delta wing of low sweep angle, having sinusoidal leading edges of varying amplitude and wavelength, are investigated using a stereoscopic technique of high-image-density particle image velocimetry at a Reynolds number of 15,000. Identification of critical points, in conjunction with surface-normal vorticity and velocity, provides a basis for determining the effectiveness of a given leading edge. At high angle of attack, where large-scale three-dimensional separation occurs from the wing with a straight leading edge, an amplitude of the leading-edge protuberance as small as one-half of one percent of the chord of the wing can substantially alter the near-surface topology. When the amplitude reaches a value of four percent of the chord, it is possible to completely eradicate the negative focus of large-scale, three-dimensional separation, in favor of a positive focus of attachment. Moreover, alteration of the near-surface topology is most effective when the ratio of the wavelength to amplitude of the sinusoidal leading edge is maintained at a small value.     

Combined method of calculating supersonic ideal-gas flow over a wing with a supersonic blunt leading edge

A combined numerical method, based on the successive calculation of the flow regions near the blunt leading edge and center of a wing, is proposed on the assumption that the angle of attack and the relative thickness and bluntness radius of the leading edge are small. The flow in the neighborhood of the leading edge of the wing is assumed to be identical to that on the windward surface of a slender body coinciding in shape with the surface of the blunt nose of the wing and is numerically determined in accordance with [1]. The flow parameters near the center of the wing are calculated within the framework of the law of plane sections [2]. In both regions the equations of motion of the gas are integrated by the Godunov method. The flow fields around elliptic cones are obtained within the framework of the combined method and the method of [3], A comparative analysis of the results of the calculations makes it possible to estimate the region of applicability of the method proposed.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 1, pp. 159–164, January–February, 1989.The authors wish to express their gratitude to A. A. Gladkov for discussing their work, and to G. P. Voskresenskii, O. V. Ivanov, and V. A. Stebunov for making available a program for calculating supersonic flow over a wing with a detached shock.     

Supersonic three-dimensional flow around a delta wing with blunted leading edges

We consider the problem of steady flow of an inviscid, non-heat-conducting gas about a delta wing which is spherically blunted at the nose and cylindrlcally blunded on the leading edges, at an angle of attack.Several experimental and theoretical studies have been devoted to the investigation of this problem, of which we note [1–4], In the following the three-dimensional method of characteristics using the scheme proposed in [5] is used to calculate the flow fields about such bodies for freestream Mach numbers M=6, 7, 8, and , sweep angle =70°, and angles of attack from 0 to 15°.     

Computational modeling of vortex breakdown control on a delta wing

We present an effort to model the development and the control of the vortex breakdown phenomenon on a delta wing. The pair of the vortices formed on the suction side of a delta wing is the major contributor to the lift generation. As the angle of attack increases, these vortices become more robust, having high vorticity values. The critical point of a delta wing operation is the moment when these vortices, after a certain angle of attack, are detached from the wing surface and wing stall occurs. In order to delay or control the vortex breakdown mechanism, various techniques have been developed. In the present work, the technique based on the use of jet-flaps is numerically investigated with computational fluid dynamics by adopting two eddy-viscosity turbulence models. The computational results are compared with the experimental data of Shih and Ding (1996). It is shown that between the two turbulence models, the more advanced one, which adopts a non-linear constitutive expression for the Reynolds-stresses, is capable to capture the vortex breakdown location for a variety of jet exit angles. The performance assessment of the models is followed by the investigation of the effect of the jet-flap on the lift and drag coefficients.     

Vortex breakdown and its control on delta wings

In order to investigate the breakdown of vortices generated by the leading edge of delta wings, LDA-measurements have been performed in the flow on the suction side of a delta wing of aspect ratio A = 2. The measurements describe the growth of the vortex along the leading edge and reveal a certain radial structure upstream of the breakdown point. Moreover they shed light on the mechanism responsible for the onset of vortex breakdown on the suction side of a wing.

The occurrence of the breakdown phenomenon on a delta wing may be prevented or at least retarded by the use of spanwise blowing jets. The interaction of vortex and jets giving rise to these effects will also be discussed with the help of measured velocity profiles.     

Numerical analysis of the flow pattern and vortex breakdown over a pitching delta wing at supersonic speeds

A supersonic compressible flow over a 60° swept delta wing with a sharp leading edge undergoing pitching oscillations is computationally studied. Numerical simulations are performed by the finite volume method with the use of the k?ω turbulence model for various Mach numbers and angles of attack. Variations of flow patterns in a crossflow plane, hysteresis loops associated with the vortex core location, and vortex breakdown positions during a pitching cycle are investigated. Trends for various Mach numbers, mean angles of attack, pitching amplitudes, and pitching frequencies are illustrated.     

Non-uniform recovery of vortex breakdown over delta wing in response to blowing along vortex core

The transient characteristics of vortical structure over delta wing are studied experimentally when subject to single along-core blowing perturbation. Two half delta wing models with different sweep angle = 60° and = 75° are investigated in this study. For = 75°, the transient location of the onset of vortex breakdown moves upstream monotonously toward the unperturbed location. However, for =60°, there exists a chordwise region where the upstream propagation of the onset location of vortex breakdown is temporarily delayed. This delay causes the recovery process (upstream propagation) of the onset location of vortex breakdown to be non-uniform. In fact, this non-uniform recovery may take on different appearance such as a plateau or overshoots and will last for several times of the convective time scaleC/U . Furthermore, the location of chordwise region corresponding to this delay depends strongly upon the angle of attack (AOA). In additions, the non-uniform recovering characteristic of the onset of vortex breakdown may not be observed at high AOA if the blowing rate is too low. The mechanism governing the non-uniform recovering characteristic is clearly verified through the LDA measurement and the phase-locked flow visualization. Evidently the mutual interaction between the primary vortex structure and the secondary vortex is the key mechanism that leads to non-uniform recovering character over delta wing with sweep angle of 60°The authors are grateful for the support of the this investigation from National Science Foundation of the Republic of China under the grant no. NSC-83-0424-E-005-006.     

Reynolds number effects on leading edge vortex development on a waving wing

The waving wing experiment is a fully three-dimensional simplification of the flapping wing motion observed in nature. The spanwise velocity gradient and wing starting and stopping acceleration that exist on an insect-like flapping wing are generated by rotational motion of a finite span wing. The flow development around a waving wing at Reynolds number between 10,000 and 60,000 has been studied using flow visualization and high-speed PIV to capture the unsteady velocity field. Lift and drag forces have been measured over a range of angles of attack, and the lift curve shape was similar in all cases. A transient high-lift peak approximately 1.5 times the quasi-steady value occurred in the first chord length of travel, caused by the formation of a strong attached leading edge vortex. This vortex appears to develop and shed more quickly at lower Reynolds numbers. The circulation of the leading edge vortex has been measured and agrees well with force data.     

Effects of Gurney flaps on the lift enhancement of a cropped nonslender delta wing

An experimental wind-tunnel investigation was undertaken to determine the effects of Gurney flaps on a 40-deg cropped nonslender delta wing at a chord Reynolds number of 250,000. In the experiment, the height of the Gurney flaps was varied from 0.01C to 0.05C, and the sideslip angle of the model was selected as 0, 5, 10 and 20 deg. In addition, the 0.05C Gurney flap was serrated with different heights of 0.01C to 0.05C separately. In comparison with the baseline clean configuration results, it was found that the model with plate Gurney flaps can indeed increase the lift-to-drag ratio at moderate-to-high lift coefficients for the wing, and the greatest increment was obtained for the 0.01C Gurney flap. The effect of Gurney flap on the increment of lift-to-drag ratio tends to be not significant with the increase of sideslip angle. Moreover, the 0.05C serrated Gurney flap provides the best performance among the serrated Gurney flaps. Received: 6 July 2000 / Accepted: 21 June 2001 Published online: 29 November 2001     

Detailed measurements in the transonic vortical flow over a delta wing

Some of the results of an international experiment are presented which documents surface and field information on compressible flow over a delta wing, with particular emphasis on the vortical flow on the lee side. The main aim of the experiment, to provide data for testing the validity of using Euler computations for such flows, is also discussed.     

Vortex diffraction on a swept wing

The forces acting on a swept wing in the presence of a vortex induced by a delta wing, as well as the velocity field in the vicinity of the swept wing, have been measured. By means of the “frozen,” vortex model and a specially-developed numerical panel method, the forces and moments acting on the wing are calculated from the known velocity field. Comparison of the calculated and measured force characteristics makes it possible to determine the extent to which the model fits the physical flow pattern. It is shown that for the intense vortex considered in this study the model gives results which disagree sharply with the experimental data. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 6, pp. 98–105, November–December, 1998. The study was supported by the International Scientific and Technical Center under grant No. 201.     

Vortex control by the spanwise suction flow on the upper surface of delta wing

The numerical investigation has been performed to explore the feasibility of vortex control by leading edge sucking excitation on a delta wing. The results reveal that the flow on the upper surface of the delta wing changes significantly in a wide range of the angle of attack. For the vortical flow at moderate angle of attack, the secondary and tertiary vortices are weakened or suppressed, and the total lift is almost unchanged. For the stalled flow at high angle of attack, the leading edge concentrated vortex is recovered, and the lift is enhanced with increasing suction rate. For the bluff-body flow at even high angles of attack, the lift can still be improved. The concentrated vortex disappears on the upper surface, and the load increment is nearly unchanged along the chordwise direction. The project supported by the National Natural Science Foundation of China (19802018).     

Experiments on the unsteady nature of vortex breakdown over delta wings

 Vortex breakdown location over delta wings is not steady and exhibits fluctuations along the axis of the vortices. Experiments on the nature and source of these fluctuations were carried out. Spectral analysis and other statistical concepts were used to quantify the unsteady behaviour of vortex breakdown location obtained from flow visualization. The fluctuations consist of quasi-periodic oscillations and high-frequency low amplitude displacements. The quasi-periodic oscillations are due to an interaction between the vortices, which cause the antisymmetric motion of breakdown locations for left and right vortices. The oscillations are larger and more coherent as the time-averaged breakdown locations get closer to each other as angle of attack or sweep angle is varied. The frequency of this organized motion is much smaller than the frequency of any other known instabilities. On the other hand, the most probable frequency for the high-frequency small-amplitude fluctuations of breakdown location is in the same range as the frequency of Kelvin–Helmholtz instability of the separated shear layer. A mechanism for the interaction between the vortices causing the oscillations of breakdown location was proposed. When a splitter plate was placed in the symmetry plane of the wing, the large amplitude quasi-periodic oscillations of breakdown location were suppressed. Received: 10 March 1998 / Accepted: 27 October 1998