Slender Delta Wing Research Articles

Generalized, Three-Dimensional Wind Tunnel Wall Correction Using Nonlinear Least-Squares Optimization

A novel three-dimensional technique for correcting wind tunnel measurements for wall interference has been developed. The flow around the test body and its wake is approximated as a series of inviscid singularities, while the wind tunnel walls and model support are approximated using a panel method. A nonlinear least-squares approach with trust-region reflective optimization is then used to obtain the strength and locations of singularities together with the domain boundary panel strengths by fitting to measured wall pressures. The technique is demonstrated using computational fluid dynamics for the test case of a well-documented slender delta wing body with high blockage and validated against both experiments and legacy data. The nonlinear least-squares wall correction method is shown to demonstrate a significant improvement in wall corrections compared to the classical two-variable technique.

Introduction of a biomimetic device designed to improve the flow over a slender delta wing: visualization study

Introduction of a biomimetic device designed to improve the flow over a slender delta wing: visualization study

Vortex Interaction Characteristics of Multiswept Wings at Subsonic Speeds

Despite the sufficient knowledge on the physics of vortices formed over slender delta wings, there is a lack of understanding about how corotating and counter-rotating vortices near the wing leading edge interact with each other. The vortex interaction will depend on the wing planform, leading-edge sweeps, leading-edge shapes, angle of attack, angle of sideslip, and the Reynolds number. In this study, experiments were performed on a range of planform configurations to determine the influence of multisweep wing configurations on the vortex interactions and breakdown thereof. Force/moment measurements were conducted to understand globally the effect of the planform configuration, whereas particle tracking surface oil flow and stereoscopic particle image velocimetry were employed to detect interactions locally in the flowfield. Results were validated against corresponding numerical simulations and past experiments. Comprehensive results indicated a strong correlation between the chord ratios of the initial sweep on the suppression of vortex breakdown. Specifically, two cases with varying chord ratios were identified that exhibited unique vortex breakdown progression (gradual and rapid). Flowfield measurements elucidated key flow features responsible for the differing breakdown progression.

Wing rock mode and its mechanism of a flying-wing aircraft

The flying wing is an aerodynamic configuration with high efficiency, but the lack of lateral-directional stability has always been an obstacle that limits its application. In this study, the wing rock motion of a 65° swept flying-wing aircraft is studied via wind tunnel experiments and numerical simulations at a low speed, and various unsteady motion phenomena are focused on. Both the experimental and numerical results show that the flying wing has a bicyclic ${C_l}$ – $\phi $ hysteresis loop during its wing rock, different from the slender delta wing, rectangular wing, generic aircraft configuration, etc., which have a tricyclic hysteresis loop. This form of hysteresis loop implies a different energy exchange manner of the flying wing in the wing rock oscillation. Further analysis shows that the flying wing forms a unilateral leading-edge vortex (LEV) under a high roll angle, with its wing rock oscillation driven by the ‘vortex–shear-layer’ structure, which is different from that of slender and non-slender delta wings. Moreover, the quantitative dynamic hysteresis characteristics of the LEV's strength and location for the flying wing and the slender delta wing are also different. These results have proven the existence of a wing rock mode which is different from previous investigations, which enriches the understanding of self-induced oscillation. Present discoveries are also conducive to the aerodynamic shape design and flight manipulation of a flying-wing aircraft, which is significant for its wider application.

Numerical Investigation of a Slender Delta Wing with Antisymmetric Leading-Edge Flap Deflection

Numerical Investigation of a Slender Delta Wing with Antisymmetric Leading-Edge Flap Deflection

The leading-edge vortex over a swift-like high-aspect-ratio wing with nonlinear swept-back geometry

The leading-edge vortex (LEV) is a common flow structure that forms over wings at high angles of attack. Over the years, LEVs were exploited for augmenting the lift of man-made slender delta wings aircraft. However, recent observations suggested that natural flyers with high-aspect-ratio (high-AR) wings, such as the common swift (Apus apus), can also generate LEVs while gliding. We hypothesize that the planform shape and nonlinear sweep (increasing towards the wingtip) enable the formation and control of such LEVs. In this paper, we investigate whether a stationary LEV can form over a nonlinear swept-back high-AR wing inspired by the swift’s wing shape and evaluate its characteristics and potential aerodynamic benefit. Particle image velocimetry (PIV) measurements were performed in a water flume on a high-AR swept-back wing inspired by the swift wing. Experiments were performed at four spanwise sections and a range of angles of attack for a chord-based Reynolds number of . Stationary LEV structures were identified across the wingspan by utilizing the proper orthogonal decomposition (POD) method for angles of attack of 5∘–15∘. The size and circulation of the stationary LEV were found to grow towards the wingtip in a nonlinear manner due to shear layer feeding and spanwise transport of mass and vorticity within the LEV, thus confirming that nonlinear high-AR swept-back wings can generate stationary LEVs. Our results suggest that the common swift can generate stationary LEVs over its swept-back wings to glide slower and at a higher rate of descent, with the LEVs potentially supporting up to 60% of its weight.

Vortical flow characteristics of a slender delta wing in ground effect

This study examined the ground effect on the vortical flow structures over a 70-deg-swept slender delta wing qualitatively and quantitatively using dye flow visualization and Particle Image Velocimetry techniques. The observations with dye visualization and PIV methods were conducted in the plan-view, side-view, and end-view planes with the angles of attack of α = 25° and 35°. The distance, h between the delta wing and the ground surface nondimensionalized with the chord, c of the delta wing, was considered as h/c = 0.1, 0.25 0.55 as well as the ground-free effect flow condition. Also, the experiments were performed at a normalized chordwise location of x/c = 0.8 over the end-view plane. The results demonstrate that magnitudes of leading-edge vortices progressively diminish when the delta wing moves from the ground-free effect region towards the wing location, h/c = 0.1. Also, the rate of inboard flow from the side leading-edge towards the central axis of the wing gets higher by the influence of the ground over the suction surface of the wing. When the wing altitude is lowered towards the ground surface, the size of the leading-edge vortex core region enlarges for both angles of attack, α = 25°, and 35°. The efficiency of the wing aerodynamics is positively affected by its closeness to the ground due to the air cushion effect that occurs as a result of the pressure increase on the suction surface of the wing due to the stagnation of air in the space between the delta wing and ground.

A new extended state observer for uncertain nonlinear systems

Extended state observer (ESO) is a class of high-gain observers and a powerful tool for output feedback control of uncertain nonlinear systems. In this paper, we propose a new ESO design technique which is based on cascading a series of first-order ESOs. Benefited from the cascade design, saturations can be inserted into the observer internal variables to limit the maximal implemented gain. The convergence of the new ESO with a class of nonlinear gain functions is established, and it is shown that adopting appropriate nonlinear gain mechanisms leads to improvement in measurement noise tolerance. What is more, the new ESO based output feedback control has stronger uncertainty estimation and compensation capability than the standard ESO based output feedback control. Specifically, the proposed approach only relies on the knowledge of the control direction of the uncertain nonlinear system. Finally, the obtained results are applied to suppress the wing rock motion of a slender delta wing.

Design and computational fluid dynamics analysis of bio-inspired non slender cropped delta wingsuit

This study examines the effect of bio-inspired serrations on aerodynamic performance of a Non Slender cropped delta wingsuit model. In this study, Gottingen 228 aerofoil is used for designing the wingsuit model having AR of 1.00 using CAD software Solidworks. Ansys Fluent solver has been utilised to solve the Reynolds Averaged Navier–Stokes (RANS) equations with a k− $$\omega$$ turbulence model. The flow velocity was kept at 45 m/s with Re ~ 2.6 × 106 and angle of attack was varied from 0° to 45°. Computations revealed that the wingsuit with serrations in comparison with clean basemodel had a notable increase in maximum lift coefficient. The results were compared with the experimental and CFD results of existing literature in the open domain. The serrated non slender delta wingsuit performs extremely well giving a lift coefficient of 2.74 and $$C_{{\text{L}}}$$ / $$C_{{\text{D}}}$$ of 9.72. The results were validated by comparing them with flat plate and non slender cropped delta wing results available in the existing literature. A good agreement in terms of trends was obtained for $$C_{{\text{L}}}$$ and $$C_{{\text{D}}}$$ which indicates that proposed wingsuit should perform well aerodynamically under typical wingsuit flying conditions.

Near-surface particle image velocimetry measurements over a yawed slender delta wing

In this study, extensive instantaneous velocity measurements were conducted within a flow area by stereo particle image velocimetry (SPIV) to investigate the influence of the yaw angle, β, on the vortical flow structure formed on a slender delta wing. This sideslip angle, β, in the yaw plane was varied from 4° up to 20° with an interval of 4° at two critical angles of attack, α = 25° and 35°, respectively. In order to reveal the influence of the yaw angle, β over the flow structure of the delta wing, time-averaged flow statistics, and instantaneous flow data obtained by the SPIV technique in the plan-view plane close to the suction surface of the delta wing were presented. It was observed that even a low yaw angle, for instance β = 8°, becomes to be effective on the flow characteristics of the delta wing, and this effect was augmented with increasing β. The influence of β is quite high on the vortical flow structure at α= 35° compared to the angle of attack of α = 25°. The flow structure that is symmetrical with respect to the centerline of the wing in the case of no yaw has disrupted with the existence β. Furthermore, the extent of the asymmetry enlarges with increasing β. The leading-edge vortex (LEV) on the windward side broken earlier and dominated the flow on the wing surface. It is concluded that this asymmetric flow structure can deteriorate the aerodynamic performance and cause other adverse effects such as unsteady loading.

The impact of the pitching motion on the structure of the vortical flow over a slender delta wing under sideslip angle

The primary purpose of this investigation is to observe the effect of the pitching motion on the vortical flow structure and bursting of leading-edge vortices over a delta wing under the sideslip angle, β using a dye visualization technique. In the current work, a delta wing with a sweep angle of Λ = 70° was oscillated in upstroke and downstroke direction to be able to discover the influence of pitching motion on the flow characteristics of the delta wing. The values of mean angles of attack were selected as αm = 25° and αm = 35°, and the sideslip angle was altered from β = 0 to 16°. The delta wing oscillated with the various periods of Te = 5 s, 20 s, and 60 s, respectively. Amplitude of motion was adjusted as αo = ± 5°. It is found that the pitching motion of the delta wing under the sideslip angle β varies the location of the vortex bursting and vortical flow structure substantially.

Linear Adaptive Fault Tolerant Control against Aircraft Actuator Failures for Wing Rock Suppression

In this paper we consider the problem of the suppression of wing rock phenomenon in presence of actuator faults. The main aim of this work is the application of a linear adaptive actuator failure compensation control scheme for wing rock motion of a slender delta wing. First, the analytical nonlinear model that characterizes the phenomenon of wing rock is given, and its behavior for different angles of attack is presented. Then, an adaptive actuator failure compensation redundant control design is proposed, and applied for a rolling motion regulation in the presence of actuator faults. Finally, numerical simulations are conducted to show the effectiveness of the control scheme for wing rock suppression for different actuator failures scenarios.

Salınım hareketinin sapma açısı altındaki bir delta kanada etkileri

In this study, the main objective is to reveal the change of the vortical flow characteristics on a slender delta wing with a sweep angle of Λ=70º under the alteration of three different sideslip angles of β=0º, 4º, and 8º at three different main angles of attack of αm=25º, 30º and 35º experimentally. Dye visualization experiments were conducted at a Reynolds number of Re=2x104 to analyze qualitatively. The perturbation motion which has amplitudes of α0=±0.5º and ±1º was applied continuously under the periods of time of Te=0.5s, 1s, and 2s. However, the effects of the low-frequency perturbations to which the aircraft is exposed during the maneuver were examined. It is observed that Kelvin-Helmholtz in the case of the perturbed motion vortices is more prominent than static case results of the slender delta wing. The vortex breakdown location on the perturbed delta wing has been changed in a wider range than the static case of the delta wing at zero sideslip angle, β.

Investigation of crossflow features of a slender delta wing

In the present work, the main features of primary vortices and the vorticity concentrations downstream of vortex bursting in crossflow plane of a delta wing with a sweep angle of Λ=70° were investigated under the variation of the sideslip angles, Β. For the pre-review of flow structures, dye visualization was conducted. In connection with a qualitative observation, a quantitative flow analysis was performed by employing Particle Image Velocimetry (PIV). The sideslip angles, Β were varied with four different angles, such as 0°, 4°, 12°, and 20° while angles of attack, α were altered between 25° and 35°. This study mainly focused on the instantaneous flow features sequentially located at different crossflow planes such as x/C=0.6, 0.8 and 1.0. As a summary, time-averaged and instantaneous non-uniformity of turbulent flow structures are altered considerably resulting in non-homogeneous delta wing surface loading as a function of the sideslip angle. The vortex bursting location on the windward side of the delta wing advances towards the leading-edge point of the delta wing. The trajectory of the primary vortex on the leeward side slides towards sideways along the span of the delta wing. Besides, the uniformity of the lift coefficient, CL over the delta wing plane was severely affected due to unbalanced distribution of buffet loading over the same plane caused by the variation of the sideslip angle, Β. Consequently, dissimilarities of the leading-edge vortices result in deterioration of the mean value of the lift coefficient, CL.

Vortex force maps for three-dimensional unsteady flows with application to a delta wing

Abstract

Leading-Edge Vortex as a High-Lift Mechanism for Large-Aspect-Ratio Wings

Enabling a leading-edge vortex (LEV) is a possible mechanism to significantly increase the lift on wings. This is well known for slender delta wings, for which various analytical models were developed and show accurate predictions of the lift enhancement. This paper shows that, when suitably designing the wing planform, LEVs can also exist on high-aspect-ratio (high-AR) wings. In this study, a quasi-three-dimensional flow model is presented for solving the stationary LEV phenomenon over a high-AR swept back wing, with sweep increasing toward the wingtip. Our model captures the three-dimensional phenomenon by satisfying conservation of mass and vorticity within the LEV, and using a combination of strip theory and the lifting-line theory. Our model predictions are compared with flow visualization data on a parabolic swept back wing. Results confirm that suitable sweep wing geometry can fix an LEV steadily over the upper wing surface, resulting in significant lift enhancement of up to 70%.

Experimental Research on Wingtip Vortices using a Half Deltawing at the Tips

The counter rotating wing tip vortices produced by the aircraft continues to be a big concern for the aviation industry and the aircraft manufacturers due to its hazardous effects on the flight safety and aircraft efficiency. The strength of the vortices poses severe problems to the aircraft operations. Manufacturers developed various wingtip devices to alleviate this problem, but still it is not fully understood and solved. In this thesis, the effectiveness of using a half delta wing at the tips is investigated. The flow field over a low aspect ratio NACA 0015 wing fitted with a slender sharp half delta wing with a leading edge sweep angle 700 at a Reynolds number 1.87 ×105 is investigated. Particle image velocimetry is used to quantify the vortex structure and force balance measurements are used to calculate the aerodynamic data of the wing. The peak vorticity, peak tangential velocity are decreased due to the addition of half delta wing. The over-all radius of the wingtip vortex increased showing a diffused vortex due to the addition of the half delta wing. The core circulation is decreased leading to a lower strength vortex. Though the tip device increased the drag, it increases the aerodynamic efficiency through the improvement in L/D

Experiment on Wingtip Vortices using a Half Deltawing at the Tips

The counter rotating wing tip vortices produced by the aircraft continues to be a big concern for the aviation industry and the aircraft manufacturers due to its hazardous effects on the flight safety and aircraft efficiency. The strength of the vortices poses severe problems to the aircraft operations. Manufacturers developed various wingtip devices to alleviate this problem, but still it is not fully understood and solved. In this thesis, the effectiveness of using a half delta wing at the tips is investigated. The flow field over a low aspect ratio NACA 0015 wing fitted with a slender sharp half delta wing with a leading edge sweep angle 700 at a Reynolds number 1.87 ×105 is investigated. Particle image velocimetry is used to quantify the vortex structure and force balance measurements are used to calculate the aerodynamic data of the wing. The peak vorticity, peak tangential velocity are decreased due to the addition of half delta wing. The over-all radius of the wingtip vortex increased showing a diffused vortex due to the addition of the half delta wing. The core circulation is decreased leading to a lower strength vortex. Though the tip device increased the drag, it increases the aerodynamic efficiency through the improvement in L/D.

Effect of Yaw Angles on Aerodynamics of a Slender Delta Wing

AbstractIn the present investigation, flow structures over a slender delta wing with a 70° sweep angle, Λ at an angle of attack, α of 30°, and under the variation of yaw angles within the range of ...

Effects of coarse riblets on air flow structures over a slender delta wing using particle image velocimetry

This research investigates the aerodynamic performance and flow characteristics of a delta wing with 65° sweep angle and with coarse axial riblets, and then compares with that of a smooth-surface delta wing. Particle Image Velocimetry (PIV) were utilized to visualize the flow over the wing at 6 cross-sections upright to the wing surface and parallel to the wing span, as well as 3 longitudinal sections on the leading edge, symmetry plane, and a plane between them at Angles of Attack (AOA) = 20° and 30° and Re = 1.2 × 105, 2.4 × 105, and 3.6 × 105. The effects of the riblets were studied on the vortices diameter, vortex breakdown location, vortices distance from the wing surface, flow lines pattern nearby the wing, circulation distribution, and separation. The results show that the textured model has a positive effect on some of the parameters related to drag reduction and lift increase. The riblets increase the flow momentum near the wing’s upper surface except near the apex. They also increase the flow momentum behind the wing.