Aerodynamic Performance Of Delta Wing Research Articles

Experimental study of flow control on delta wings with different sweep angles using pulsed nanosecond DBD plasma actuators

A low-speed wind tunnel study of active flow control for improving aerodynamic performance of delta wings at high angles of attack is presented. A dielectric barrier discharge actuator driven by pulsed nanosecond voltage is used to create periodic perturbations along the leading edge of delta wing. Four wings with sweep back angles of Λ = 30°, 47°, and 60° are tested. Data obtained by means of force measurements show that the effectiveness of the actuation strongly depends on the wing’s sweep angles and the actuation frequency. For Λ = 30°, 47° delta wings, the pulsed actuation could increase the lift at post stall regions. For Λ = 60° delta wing, the pulsed actuation at the leading edge almost has no effect on its aerodynamic performance. At last, the control mechanism is discussed.

Control of vortex on a non-slender delta wing by a nanosecond pulse surface dielectric barrier discharge

Wind tunnel experiments are conducted for improving the aerodynamic performance of delta wing using a leading-edge pulsed nanosecond dielectric barrier discharge (NS-DBD). The whole effects of pulsed NS-DBD on the aerodynamic performance of the delta wing are studied by balanced force measurements. Pressure measurements and particle image velocimetry (PIV) measurements are conducted to investigate the formation of leading-edge vortices affected by the pulsed NS-DBD, compared to completely stalled flow without actuation. Various pulsed actuation frequencies of the plasma actuator are examined with the freestream velocity up to 50 m/s. Stall has been delayed substantially and significant shifts in the aerodynamic forces can be achieved at the post-stall regions when the actuator works at the optimum reduced frequency of F + = 2. The upper surface pressure measurements show that the largest change of static pressure occurs at the forward part of the wing at the stall region. The time-averaged flow pattern obtained from the PIV measurement shows that flow reattachment is promoted with excitation, and a vortex flow pattern develops. The time-averaged locations of the secondary separation line and the center of the vortical region both move outboard with excitation.

Effects of flexibility on aerodynamic performance of delta wings with different sweep angles

Force measurement and surface oil flow visualization experiments were conducted in a wind tunnel to investigate the effects of flexibility on aerodynamic performance of delta wings with different sweep angles. The experimental results indicate that the maximum lift coefficient is increased and the stall angle is delayed as the sweep angle increases for both rigid and flexible wings. It is also found that the maximum lift coefficients of the flexible wings with a sweep angle from 35° to 50° are higher than those of the rigid ones. The increment of the maximum lift coefficient in particular achieves 32.9% compared with the case without lift enhancement for the 40° flexible delta wing. Moreover, the surface oil flow visualization experiments show that the stall of the flexible wing of the moderate low sweep angle is accompanied by helical flow structure, while the vortex bursting appears on the corresponding rigid wing.

Effect of leading-edge vortex flaps on aerodynamic performance of delta wings

The effect of leading-edge vortex flaps on the aerodynamic characteristics of highly swept-back wings is analytically investigated, using the free vortex sheet method. The method, based on a three-dimensional inviscid flow model, is an advanced panel type employing quadratic doublet distributions to represent the wing surface, rolled-up vortex sheet and wake and is capable of computing forces, moments and surface pressures.

Aerodynamic Performance of Delta Wings at Supersonic Speeds

In recently published papers the authors have discussed the problem of drag of delta wings with straight trailing edges and also the lift of this type of wing. The present paper consists of an extension of these results to delta wings with swept-back or swept-forward trailing edges. Drag at zero lift for two slope surfaces and lift is computed for these airfoils, and a discussion is given of the drag due to lift. Centers of pressure are also computed. In addition, some results for drag of wings with three slope surfaces are given. Discussion is made of optimum design from the standpoint of leading and trailing edge angles, and location of maximum thickness. I t is shown that considerable improvements over two-dimensional airfoil performance can be obtained. Results are presented in the form of conventional engineering coefficients.