Slender Delta Wing Research Articles

Effect of Dorsal Fin on Symmetry of Vortices over Slender Delta Wing

Three different sets of experiments are designed to reveal the behavior of the vortex flow over a delta wing with different fin heights. A sharp-edged flat-plate delta wing with an 82.6-deg-sweep angle is tested in low-speed wind tunnels at angles of attack up to 35 deg. The same tests are performed on an identical delta-wing model but with a flat-plate dorsal fin. The ratio of the local fin height to the local wing semispan varies from 0.3 to 1.5. The results of smoke laser-sheet visualization, particle image velocimetry, and force measurement are obtained and presented. The test results show that the vortices over the delta-wing-alone model at high angles of attack and zero sideslip are stable, symmetric, and steady before vortex breakdown, but a low dorsal fin renders the originally symmetric vortices asymmetric at the same angle of attack; the flow recovers symmetry and conicity when a higher fin is added. These results are in agreement with predictions by the stability theoryCaiet al., (“Stability of Symmetric Vortices in Two Dimensions and over Three-Dimensional Slender Conical Bodies,” Journal of Fluid Mechanics, Vol. null, 2003, pp. 65–94). This agreement demonstrates that the absolute type of instability is responsible for loss of symmetry of the vortices over a delta wing with a fin within the range of tested conditions.

Effect of aspect ratio on wing rock at low Reynolds number

Wing rock phenomenon has been studied extensively on slender delta wings but a little is understood about wing rock occurring on rectangular wing planforms, which is the focus of this paper. Specifically, a systematic procedure is discussed for experimental studies of single degree-of-freedom wing rock on rectangular wings. The characterization of limit cycle oscillations in terms of amplitude and frequency is based on the free-to-roll tests in wind tunnel. Rectangular wings with aspect ratio from 2 to 4.5 (with increment of 0.25) are studied at low Reynolds numbers (less than 0.1 million) and low Strouhal numbers (10−2) that are typical for the flight of micro air vehicles in urban environment. The changing trend in oscillation amplitude and frequency of the resulting sustained limit cycle oscillations is analyzed. The results indicate that wing rock occurs in post-stall regime (14–16° of angle of attack) and its amplitude tends to increase from aspect ratio 2 to 3.25 and decrease thereafter. The wing-rock frequency shows a decrease from aspect ratio of 2 to 3.75 and increase thereafter.

Suppression of Wing Rock in Slender Delta Wing by Horizontal Strakes

The experimental investigations of this paper show that delta- and reverse-delta-shaped small horizontal strakes (fineness ratio of 3–5%) are capable of suppressing the wing rock of a slender delta wing above 35 deg angle of attack. The parametric study reveals that 45 deg sweepback and sweepforward angles are the most favorable combinations, while the most advantageous streamwise location of the strake is the beginning of the delta wing. The suppression capability is also evident during pitching up, but no significant deterioration of the lift-to-drag ratio is observed. The suppression is achieved by changes in the vortical flowfield brought in by the strakes. During a roll, the strake 1) alters the away normal and outward spatial location of the up-going semispan’s vortex to closer normal and inward spatial location, respectively, and 2) reduces the difference in the normal distances of vortices of two semispans, both of which are the characteristics of no-wing-rock regime, such as 15 deg angle of attack. Thus, the strake changes the asymmetric flow to nearly symmetric. Such change in the vortical flowfield is caused by the interaction of two vortices emanating from the strake, with the vortex pair originating from the delta wing.

Extended state observer based robust control of wing rock motion

In this paper, a new design based on the Extended State Observer (ESO) technique for the robust control of wing rock motion of slender delta wings is proposed. The wing rock motion dynamics with varying angle of attack is significantly uncertain. The ESO is employed to simultaneously estimate the state and the uncertainty. The estimated uncertainty is used to robustify an Input–Output Linearization based controller designed for the nominal system. Closed loop stability of the overall system is established. The notable feature of the proposed design is that it neither requires accurate plant model nor any information about the uncertainty. The effectiveness of the ESO in estimation of the uncertainties and states and in regulating the rolling motion in the presence of significant uncertainties and un-modeled servo and sideslip dynamics is illustrated by simulation. Lastly, the efficacy of the proposed design is demonstrated by comparing its performance with some well-known existing designs.

Aerodynamic studies over a manoeuvring UCAV 1303 configuration

AbstractThe flow past an unmanned combat air vehicle (UCAV) 1303 was investigated in a water tunnel in an effort to understand the complex flow field. The UCAV 1303 is characterised by a non slender, 47° swept-leading-edge delta wing, with a cranked trailing edge, a cropped wing tip and a fuselage. It also has no vertical tail. Dye visualisation pictures and five-axis strain gage load data were obtained on a 1:72 scale model in both steady flow and while the model executed preprogrammed manoeuvres. From the strain gage data, body axis force and moment data were derived. In addition to some similarities, comparison of the flow features observed with those of more well known slender delta wing flows showed some remarkable differences such as tip-stall, the presence of two like-sense vortices on the same side of the wing, their simultaneous spiral and bubble bursting. This paper predominantly discusses the flow visualisation results and attempts to correlate the load data to some of the observed flow features.

Effect of Nose Tip on Wing Rock of Slender Delta Wing

The root cause of wing rock is investigated by examining two slender delta wings (700 and 850 sweep back angle) in wind tunnel using force measurement, pressure measurement and PIV techniques. The results show presence of asymmetric flow at 200 angle of attack and initiation of wing rock at the same point for 850 model while there is neither asymmetric flow nor wing rock for 700 model suggesting close relation of flow asymmetry with wing rock. Investigation with three apparently identical nose sections reveals that the asymmetry comes from the area very close to the wing tip. This asymmetric flow causes the vortices to interact in a complex way resulting in wing rock when the vortices are in close proximity (such as for 850 model), which is not the case when the vortices are ‘comparatively away’ (such as 700 model) from each other.

A comprehensive analysis of wing rock dynamics for slender delta wing configurations

The paper deals with the study of an analytical model of wing rock, based on parameter identification of experimental data. The experiments were performed in the Aeronautical Laboratory of Politecnico di Torino, in the D3M Low Speed Wind Tunnel, on a 80∘ delta wing with a modular fuselage, designed with a cylindrical forebody and a conical nose tip. Free-to-roll tests have been used to determine build up and limit cycle characteristics of wing rock. An analytical nonlinear model was derived. Parameters were identified by means of least squares approximation of experimental data with coherent initial conditions. The consistency of time histories, reproduced by numerical integration, was also analyzed. This formulation correctly predicts stable limit cycles for a wide range of airspeeds, angles of attack, and release roll angles. Finally, the impact of aircraft configuration on wing rock parameters is here outlined.

Simplified type-2 fuzzy sliding controller for wing rock system

Wing rock is a highly nonlinear phenomenon in which aircrafts with slender delta wings undergo limit cycle roll oscillations at high angles of attack. A simplified type-2 fuzzy sliding controller is designed for suppressing wing rock phenomena and tracking the desired trajectories. To reduce the computational complexity of a type-reducer, the end points of a type-reduced set are approximated by the outputs of two standard fuzzy sliding mechanisms in the proposed simplified type-2 fuzzy sliding controller. Furthermore, the sliding modes of the fuzzy sliding control system are guaranteed. Simulation results are included to show the effectiveness of the proposed simplified type-2 fuzzy sliding controller.

Effect of Slenderness Ratio on Free-to-Roll Wing Aerodynamics

M ANNED fighter aircraft and unmanned combat air vehicles may require high angle-of-attack maneuvers where the roll dynamics become crucial. Small unmanned air vehicles (UAVs) and fixed-wingmicro air vehicles often have separated flows [1,2]. These small vehicles fly at relatively high angles of attack and close to the stall conditions due to the poor lift. Vertical gusts might induce roll instabilities at high angles of attack. This possibility is suggested based on the recent experiments in which rectangular, elliptical, and Zimmerman planforms exhibited self-induced roll oscillations, even before the stall angle [3,4]. Most of the knowledge on the aerodynamics of free-to-roll wings is on slender delta wings [5–7], which exhibit well understood freeto-roll oscillations about a zero mean roll angle. Free-to-roll aerodynamics of delta wings with lower sweep angles (or higher aspect ratios) have been investigated in recent experiments [8–14]. The existence of equilibrium at nonzero roll angles was reported for sweep angles 65 . The roll asymmetries are demonstrated in Fig. 1a for a simple delta wing with sweep angle of 55 (taken from [14]). The variation of themean roll angle as a function of angle of attack shows that nonzero roll angles are observed until the wing stalls. Here, we use the term “stall” to describe the onset of the flow regime for which there is no flow reattachment, and hence the roll angle becomes zero. Previouswork [12] confirms that the “stall angle of the free-to-roll wing” is near the stall angle of the fixed wing (at zero roll angle) at which the lift force is maximum. The stall of the free-to-roll wing is very sudden as the angle of attack is increased. In contrast, for a slender delta wing with sweep angle 70 , roll asymmetries mostly disappear, as shown in Fig. 1b. Contour plots in Fig. 2 are adapted from the data in [14], which show the mean roll angle and the standard deviation of roll angle in the free-to-roll experiments. The variation of the angle of attack at which the vortex breakdown appears at the trailing edge of the wing, as well as the stall angle of attack at which the lift is maximum (for zero roll angle), is added from another source [15] (shown as dashed lines). It is seen in Fig. 2a that, for wings with low sweep angles, roll asymmetries appear when the vortex breakdown is over thewing and becomes maximum near the stall angle. The magnitude of the mean roll angle decreases with increasing sweep angle and almost disappears for the most slender wing. The stall of the free-to-roll wing also becomes more gradual with increasing sweep angle. Figure 2b shows that, in a certain range of sweep angles and near the stall, self-induced roll oscillations are possible. In this Note, we extend our experiments to cropped delta wings with an initial leading-edge sweep angle of 55 , as shown in Fig. 3. We anticipate that the slenderness ratio, defined as the chord length divided by the span, may be important. It was shown [7] that if the distance from the vortex origin divided by the span is larger than two, the wing rock of aircraft configurations is likely. Similarly, this ratio may be important for the onset of the roll asymmetries for the nonslender free-to-roll wings. Table 1 lists the aspect ratio and slenderness ratio for the wings shown in Fig. 3. For comparison, we also added the simple delta wing with 70 into our study.

Yaw Angle Effect on Flow Structure over the Nonslender Diamond Wing

Yaw Angle Effect on Flow Structure over the Nonslender Diamond Wing

Effects of Unsteady Trailing-Edge Blowing on Delta Wing Aerodynamics

The effects of unsteady trailing-edge blowing on delta wing aerodynamics were investigated experimentally to understand the aerodynamics-propulsion interaction for dynamic thrust vectoring. Two models with sweep angles of A = 50 and 65 deg, representing nonslender and slender delta wings, respectively, were tested in a water tunnel. Flow visualization and velocity and force measurements were conducted at stall and poststall incidences. For the periodic blowing, it was found that the dynamic response of leading-edge vortex breakdown and wing normal force coefficient exhibit phase lags for both nonslender and slender delta wings. The estimated time constants are larger than those reported in the literature for unsteady wings undergoing pitching or plunging. In the case of transient blowing, the time delay for the decelerating jet is significantly larger than that for the accelerating jet. The variation of the circulation and the reattachment process near the wing surface were studied by means of velocity measurements. The range of the estimated time constants is similar at the stall and poststall incidences for both the slender and nonslender wings.

Gurney Flaps on Slender and Nonslender Delta Wings

*† ‡ A parametric low-speed wind tunnel study has been undertaken of the effect of Gurney flaps on the aerodynamic characteristics of slender and non-slender delta wings with sweep angles of 40°, 60° and 70°. In contrast to published data, no fundamental difference in behaviour with sweep angle was found. Gurney flaps were found to give similar lift and pitching moment changes to conventional hinged control surfaces, but with a much greater drag penalty. Gurney flap effectiveness was found to increase as sweep angle is reduced; a characteristic explained by a simple analysis of lift and drag on a sharp-edged delta wing. For a 40° sweep angle (typical of a UCAV configuration), Gurney flap deployment can significantly improve lift/drag ratio at higher incidences. Further improvements can be obtained by tapering or serrating the flap.

Vortex Dynamics of Free-to-Roll Slender and Nonslender Delta Wings

Experiments were conducted on various free-to-roll delta wings with sweep angles in the range of 40―70 deg in a wind tunnel. For low sweep angles, nonzero trim angles are observed until the wing stalls after which the trim angle becomes zero. The magnitude of the roll trim angle decreases with increasing sweep angle and almost disappears for the most slender wing. In a certain range of sweep angles, and just before the stall, large amplitude self-induced roll oscillations with nonzero mean were observed for thin wings with sharp leading edges for which separation point is fixed. Two different regions of angle of attack with free-to-roll oscillations were found for certain sweep angles. Velocity measurements suggest that the oscillations are initiated by the increasing proximity of the shear layers as the angle of attack is increased. Possible mechanisms of the self-induced roll oscillations based upon the phase-averaged velocity measurements have been proposed, incorporating varying vortex size and strength as well as the movement of the shear-layer reattachment locations. Hysteresis and time lag effects are significant even though typical dimensionless frequencies are very low.

Detached-Eddy Simulations of Vortex Breakdown over a 70-Degree Delta Wing

An understanding of the vortical structures that comprise the vortical flowfield around slender bodies is essential for the development of highly maneuverable and high angle-of-attack flight. This is primarily due to the physical limits this phenomenon imposes on aircraft and missiles at extreme flight conditions. Demands for more maneuverable air vehicles have pushed the limits of current computational fluid dynamics methods in the high Reynolds number regime. Simulation methods must be able to accurately describe the unsteady, vortical flowfields associated with fighter aircraft at Reynolds numbers more representative of full-scale vehicles. One of the goals of this paper is to demonstrate the ability of detached-eddy simulation, a hybrid Reynolds-averaged Navier―Stokes large-eddy simulation method, to accurately predict the vortical flowfield over a slender delta wing at Reynolds numbers above 1 × 10 6 . Although detached-eddy simulation successfully predicted the location of the vortex breakdown phenomenon in previous work, the goal of the current effort is to further validate the method with additional experimental data from the Office National d'Etudes et Recherches Aerospatiales, such as surface pressures and turbulent kinetic energy in the vortex core. The effect of grid density and an adaptive mesh refinement technique is also assessed through comparisons with the experiment. Detailed wind-tunnel geometry, such as tunnel walls and the sting mount system, are simulated and found to make a measurable difference. Finally, modeling the laminar-to-turbulent transition is demonstrated to have a significant effect on the vortical flowfield.

Active Flow Control on a Nonslender Delta Wing

The effects of active flow control by oscillatory blowing at the leading edge of a nonslender delta wing with a Λ=50° sweep angle have been investigated. Pressure measurements and Particle Image Velocimetry measurements were conducted on a half wing to investigate the formation of leading edge vortices for oscillatory blowing, compared to the stalled flow for the no blowing case. Stall has been delayed by up to 8, and significant increases in the upper surface suction force have been observed. Velocity measurements show that shear layer reattachment is promoted with forcing, and a vortex flow pattern develops. The time averaged location of the centre of the vortical region moves outboard with increased excitation. The near-surface flow pattern obtained from the PIV measurements shows reattachment in the forward part of the wing. There is no measurable jet-like axial flow in the vortex core, which seems to break down at or very near the apex. This highlights that unlike slender delta wings, vortex breakdown is not a limiting factor in the generation of lift for nonslender delta wings. Phase averaged measurements reveal the perturbation due to the pulsed blowing, its interaction with the shear layer and vortex, apparent displacement of the vortex core, and relaxation of the reattachment region. The flow in a phase averaged sense is highly three dimensional. Experiments indicate that unsteady blowing at Strouhal numbers in the region of St=0.5 to St=0.75, and in the region of St=1.25 to St=1.5 can be a highly effective. Reattached flow can develop from stalled flow after pulsing has been initiated with a time constant of tU/c=5 for unsteady blowing at St=0.75, and tU/c=7 for St=1.5. Experiments with excitation from finite span slots located in the forward half of the wing show that partial blowing can be more effective at low momentum coefficients. Force measurements of a full delta wing confirmed that the effectiveness of this method of flow control was not only confined to half delta wings.

Experimental Studies on Unsteady Lateral Blowing on NACA 0012

T HE initial study of steady lateral blowing was conducted by Dixon in 1969 [1]. The mechanism of this concept is analogous to the flow over slender delta wing platforms, which produces a leading-edge vortex and nonlinear lift curves. Steady lateral blowing might be thought of as providing the sweep effects such as those of the delta wings, for which the effective sweep is a function of the jet momentum. The lateral blowing jet can be described in terms of the chordwise position along the airfoil section, the nozzle height, and the nozzle size. Dixon et al. [2] andClarke [3] suggested that the best vertical nozzle positionmay be a function of the nozzle diameter and that the jet from a nozzle placed too near the airfoil surface may have a deleterious effect on the flow over the airfoil. Wong and Kontis [4] performed a comprehensive study on the steady spanwise blowing on a NACA 0012 airfoil section. The force measurements showed that the lift coefficient increasedwhen the steady lateral blowingwas applied, and the most effective blowing location was found to be at x=c 0:25 (x is the chordwise location and c is the airfoil chord), because the lift augmentation ratio Cl=C was always above one for all positive angles of attack [ Cl is the increment in the lift coefficient due to blowing, andC is the blowing momentum coefficient; the definition ofC is shown inEq. (1)].However, largemomentumwas required to generate the lift enhancement on the airfoil and this became a major issue in the effectiveness of blowing. As a consequence, Meyer and Seginer [5] performed some initial tests for generating the same lift increase using lower momentum through pulsed blowing. One of the main disadvantages of lateral blowing is due to the fact that air for blowing is bled from the compressor stage of the engine. This reduces the amount of air for combustion, hence reducing thrust. Unsteady blowing provides the beneficial effect of reducing the mass-flow requirement from the engine. Another adverse effect of the technology arises from the need to install plumbing, control valves, regulators, oscillators, etc., to supply and control the compressed air from the engine. This adds weight to the aircraft, and a higher lift is required. C thrust

Steady vortex force theory and slender-wing flow diagnosis

The concept vortex force in aerodynamics is systematically examined based on a new steady vortex-force theory (Wu et al., Vorticity and vortex dynamics, Springer, 2006) which expresses the aerodynamic force (and moment) by the volume and boundary integrals of the Lamb vector. In this paper, the underlying physics of this theory is explored, including the general role of the Lamb vector in nonlinear aerodynamics, its initial formation, and its relevance to the total-pressure non-uniformity. As a typical example, the theory is applied to the flow over a slender delta wing at a large angle of attack. The highly localized flow structures with high Lamb-vector peaks are identified in terms of their net contribution to various constituents of the total aerodynamic force. This vortex-force diagnosis sheds new light on the flow control and configuration optimization.

Analysis of Delta-Wing Vortical Substructures Using Detached-Eddy Simulation

An understanding of the vortical structures that comprise the vortical flowfield around slender bodies is essential for the development of highly maneuverable and high-angle-of-attack flight. This is primarily because of the physical limits these phenomena impose on aircraft and missiles at extreme flight conditions. Demands for more maneuverable air vehicles have pushed the limits of current computational fluid dynamics methods in the high-Reynolds-number regime. Simulation methods must be able to accurately describe the unsteady, vortical flowfields associated with fighter aircraft at Reynolds numbers more representative of full-scale vehicles. It is the goal here to demonstrate the ability of detached-eddy simulation (DES), a hybrid Reynolds-averaged Navier-Stokes/large-eddy-simulation method, to accurately model the vortical flowfield over a slender delta wing at Reynolds numbers above one million. DES has successfully predicted the location of the vortex breakdown phenomenon, and the goal of the current effort is to analyze and assess the influence of vortical substructures in the separating shear layers that roll up to form the leading-edge vortices. Very detailed experiments performed at ONERA using three-dimensional laser-Doppler-velocimetry measurement will be used to compare simulations utilizing DES turbulence models. The computational results provide novel insight into the formation and impact of the vortical substructures in the separating shear layers on the entire vertical flowfield.

Response of a Free-to-Roll Slender Delta Wings toPitching and Plunging

Response of a Free-to-Roll Slender Delta Wings toPitching and Plunging

Reinforcement adaptive fuzzy control of wing rock phenomena

Wing rock phenomena is a limit cycle roll oscillation experienced by aircraft with slender delta wings at a high angle of attack. A new reinforcement adaptive fuzzy control scheme is proposed for suppressing or tracking wing rock phenomena. A fuzzy logic system is used to approximate a unknown nonlinear dynamics in which a variable universe technique is applied to modify the premise parameters of FLS for improving its interpolation precision, and the derived adaptive law combining with a reinforcement-learning strategy is applied to tune its consequent parameters. The novelty of the proposed control scheme is mainly in its simplicity, robustness and stability so that it can be applied for online learning and real-time control. Three cases of time-varying wing rock control under disturbances are simulated to confirm the effectiveness and robustness of the proposed scheme.