Nonslender Delta Wings Research Articles

Experimental investigation of dynamic ground boundary condition for a non-slender delta wing

Experimental investigation of dynamic ground boundary condition for a non-slender delta wing

Method development for estimation of bleed momentum coefficient using surface pressure measurements and in situ hot wire calibration

Recently, as a passive flow control method, bleeding technique allowing flow of upstream towards suction side by interior passages, has been utilized on non-slender delta wing planforms, where effective geometries can enhance the flow field characteristics. In order to quantify the effect of the flow control, estimation of the momentum induced by the technique is of significant importance in addition to quantified aerodynamic favor. In the current work, passive bleeding momentum coefficient estimation method is developed and presented based on surface pressure measurements by utilizing Bernoulli equation and loss coefficient for a 45° swept non-slender delta wing. For the characterization of the presented technique in-situ hot wire measurements were performed allowing to extract across bleed slot pressure loss coefficient. Results represent that proposed approach can be used to compute the bleeding momentum where further advancements might lead to an important control parameter for potential active bleeding systems in different flow regimes.

Passive Poststall Flow Control on Nonslender Delta Wings Using Flags

The lift force on two nonslender delta wings with sweep angles of Λ=40 and 50°, with flags attached to the leading edge, was investigated. Substantial lift enhancement in the poststall regime of the clean wings and the stall delay were observed. With a flag, the flowfield measurements showed the reattachment of the shear layer onto the wing surface and the re-formation of the leading-edge vortex, whereas the clean wing had completely stalled flow at the same angle of attack. Over a wide range of angles of attack, mass ratios, and flag lengths, the maximum lift enhancement was attained when the dimensionless frequency of flag oscillations was in an optimal range and the flag-tip velocity had sufficient amplitude. The optimal dimensionless frequency range corresponded to the natural shear layer instabilities of the clean wing. The excitation in this range substantially increased the coherency of the separated flow due to the flag oscillations. The main mechanism of the lift increase is the excitation of the shear layer, which exhibits convective instability. This is very different from the lift enhancement mechanism on airfoils where the flags lock-in to the wake instability, which is known to have an absolute instability at the poststall angles of attack of the clean airfoil.

Effect of gurney flaps on a nonslender delta wing during large-amplitude and high-frequency dynamic pitching

The flow control effect of Gurney flaps on 50° swept delta wings during large-amplitude and high-frequency pitching were investigated both experimentally and numerically. Both force and velocity measurements were conducted and compared with the numerical simulation results. The lift hysteresis was found to be significantly affected by the Gurney flaps. It was found that the leading-edge Gurney flap (LG) acts to redistribute the vorticity over the suction surface of the wing during the upstroke process. The formation and development of the leading-edge vortex (LEV) on the suction surface was delayed by the LG, which was accompanied by enhancement of the lift at high angles of attack. The lower surface pressure can be altered with the flow velocity being slowed down by the trailing-edge Gurney flap (TG), which improves overall lift performance of the pitching wing at small to moderate angles of attack.

Sparse Pressure-Based Machine Learning Approach for Aerodynamic Loads Estimation During Gust Encounters

Estimation of aerodynamic loads is a significant challenge in complex gusty environments due to the associated complexities of flow separation and strong nonlinearities. In this study, we explore the practical feasibility of multilayer perceptron (MLP) for estimating aerodynamic loads in gusts, when confounded by noisy and spatially distributed sparse surface pressure measurements. As a demonstration, a nonslender delta wing experiencing various gusts with different initial and final conditions is considered. Time-resolved lift and drag, and spatially distributed sparse surface pressure measurements are collected in a towing-tank facility. The nonlinear MLP model is used to estimate gust scenarios that are unseen in training progress. A filtering process allows us to examine the fluctuation of the dynamic response from the pressure measurements on the MLP. Estimation results show that the MLP model is able to capture the relationship between surface pressure and aerodynamic loads with a minimum quantity of learning samples, delivering accurate estimations, despite the slightly large errors for the cases at the boundary of the datasets. The results also indicate that the dynamic response of the pressure measurements has an influence on the learning of MLP. We further utilize gradient maps to perform a sensitivity analysis, so as to evaluate the contribution of the pressure data to the estimation of gust loads. This study reveals the significant contribution of the sensors located near the leading edge and at the nose of the delta wing. Our findings suggest the potential for an efficient sensor deployment strategy in data-driven aerodynamic load estimation.

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.

Effect of ground on aerodynamics and longitudinal static stability of a non-slender delta wing

In this paper, the effect of ground on aerodynamics and longitudinal static stability of a non-slender delta wing with sweep angle of 45 degree and thickness-to-chord ratio (t/c) of 5.9% was characterized at Reynolds number of 9×104 in a low-speed wind tunnel using force and pressure measurements. The measurements were conducted for angles of attack varying between 0≤α≤30 degrees and non-dimensional heights between 3%≤h/c≤107% and height stability of the wing based on aerodynamic center in pitch (Xa) and aerodynamic center in height (Xh) was constructed. The results indicate that the effect of ground has substantial impact on both aerodynamic performance and stability of the delta wing. As the height of the wing decreases, both drag and lift forces increases where these effects were observed to be more pronounced for higher angles of attack. The maximum aerodynamic performance with increasing ground effect intensity was measured around 7 degree. Center of pressure (Xp) as well as Xa travel in a limited range over the wing chord, whereas Xh exhibits back and forth movement on the longitudinal axis resulting in interchanging stability characteristics varying with both height and angle of attack. The pressure measurements showed that the effect of ground has significant and complex impact on both vortex reattachment and vortex strength with varying angles of attack and heights such that it might promote reattachment and increases vortex strength but might cause earlier stall.

Study of the transient flow structures generated by a pulsed nanosecond plasma actuator on a delta wing

The transient flow structures produced by a pulsed nanosecond plasma actuator and the mechanism by which they are generated are investigated experimentally and through simulations for the case of flow control on a non-slender delta wing. Phase-averaged particle image velocimetry reveals a phenomenon in which, after each discharge pulse, two sub-vortices are generated in sequence and move along the shear layer regardless of the angle of attack, and this is confirmed by hot-wire anemometry. However, at high actuation frequencies (F+ = fc/U∞ ≥ 6.435), this phenomenon of double sub-vortices is not observed, and only one sub-vortex is generated per period. The results of pressure measurements indicate that each sub-vortex gives rise to a distinct pressure fluctuation on the wing surface. Numerical simulations reveal a number of residual heats resulting from plasma thermal effects in the shear layer, each of which turns out to induce a corresponding sub-vortex. At low actuation frequencies (F+ ≤ 4.29), there is a division of the initial residual heat into two independent residual heats and, hence, double sub-vortices per period, whereas at high actuation frequencies (F+ ≥ 6.435), residual heats from two consecutive periods merge into one, resulting in just one sub-vortex per period, which provides an explanation for the experimentally observed flow behavior.

Aerodynamic Efficiency Enhancement of Delta Wings for Micro Air Vehicles Using Shape-Optimization

The performance of fixed-wing micro air vehicles (MAVs) is impaired due to their small size, operating speeds, and altitude leading to limited lift, aerodynamic efficiency (lift-to-drag ratio), and low gust resilience. Nonslender delta wings have high stall angle and maneuverability, which are desirable features for MAVs. However, the aerodynamic efficiency of delta-winged MAVs is low, thus impacting the overall performance. We optimize the aerodynamic efficiency of a flat-plate delta wing using the adjoint-based aerodynamic shape optimization method coupled with incompressible Reynolds-averaged Navier–Stokes equations. The optimization improves the aerodynamic efficiency from 6.2 to 10.2 at 5° angle of attack while satisfying lift and moment constraints. The optimized wing is significantly cambered, and the axial cross section reveals that a circular arc can approximate the optimized wing airfoil. A parametric study of the circular camber delta wings with a 0–7.5% camber-to-chord ratio reveals that an increase in camber leads to linear growth in the lift, whereas the drag shows a quadratic growth. Consequently, an optimum camber exists at 2.5% for maximizing the aerodynamic efficiency (9.3); the wing has aerodynamic characteristics closely resembling the optimized wing, thus revealing that the essence of the efficiency optimization process is to identify the optimal camber distribution.

Experimental Investigation of Ground Effect on the Vortical Flow Structure of a 40° Swept Delta Wing

The ground effect influences the flow structure on a delta wing during landing and take-off processes. In this regard, comprehensive instantaneous velocity measurements and flow visualizations were carried out by particle image velocimetry and dye flow visualization techniques to reveal the ground effect on leading-edge vortex characteristics of a 40° swept delta wing. The flow behaviors on the delta wing under the impact of the ground were analyzed at two angles of attack, 8° and 11°, and the space between the ground and lower surface of the wing was nondimensionalized with the wing’s chord length. It was found that the presence of the ground caused premature leading-edge vortex breakdown due to the increasing adverse pressure gradient on the wing’s suction side along the chord direction. The ground effect caused an increase in peak value and distributions of turbulent kinetic energy on the wing surface that depended on the earlier leading-edge vortex bursting and complex and disorganized flow structures. The value of time-averaged vertical velocity was lower when the delta wing descended from the free-stream flow zone into the ground effect zone because of the blocking of fluid flow in the gap between the ground and pressure surface of the wing. Thus, it can be concluded that the ground effect is very influential on the change of vortical flow characteristics of nonslender delta wings.

Aerodynamics of non-slender delta and reverse delta wings: Wing thickness, anhedral angle and cropping ratio

The effects of thickness-to-chord (t/c) ratio, anhedral angle (δ), and cropping ratio from trailing-edge (Cr%) on the aerodynamics of non-slender reverse delta wings in comparison to non-slender delta wings with sweep angle of 45 ° were characterized in a low-speed wind tunnel using force and pressure measurements. The measurements were conducted for total of 8 different delta and reverse delta wings. Two different t/c ratios of 5.9% and 1.1%, and two different anhedral angles ofδ = 15° and 30° for non-cropped and cropped at Cr = 30% conditions were tested. The results indicate that the reverse delta wings generate higher lift-to-drag ratio and have better longitudinal static stability characteristics compared to the delta wings. The wing thickness has favorable effect on longitudinal static stability for the reverse delta wing whereas longitudinal static stability is not influenced by wing thickness for the delta wing. For reverse delta wings, the anhedraled wing without cropping has adverse effect on aerodynamic performance and decreases the lift-to-drag ratio. Cropping in anhedraled wing causes significant improvement in lift-to-drag ratio, shift in aerodynamic and pressure centers towards the trailing-edge, and enhancement in longitudinal static stability.

Low speed aerodynamic characteristics of non-slender delta wing at low angles of attack

Low-speed wind tunnel experiments are conducted to study the aerodynamic performance of a half-span delta wing with 45° leading-edge sweep at subsonic flow regime. The experiments are carried out at a Reynolds number of 8.37 × 105, a free-stream Mach number of 0.1 and angles of attack up to 25°, in steps of 5°. The test model was designed with thirty-two pressure taps fixed on its surfaces (sixteen on each side). Multi-tube manometers were connected to these taps using long tubes to enable recording the pressure readings. Surface pressure distributions and aerodynamic characteristics were calculated at different span-wise locations along the non-dimensional chord-wise distance. Results exhibited that most lift on the studied wing is generated in the region close to the leading edge for all the studied incidence angles. Additional lift is created in the region close to the root chord rather than the tip chord, whereas drag forces increases from tip to root. This can be attributed to the formation of trailing edge vortexes due to the flow separation at the wing leading edge that produces more drag, hence suppressing lift. The study showed also that angle of attack increases the drag coefficient from tip to root, especially at high angle of attack, indicating unfavourable behaviour for manoeuvring. Moreover, the angle of attack increased the pitching moment coefficient up to 10° before it drops sharply until it reaches the tip of the wing model.

Estimation of aerodynamic coefficients of a non-slender delta wing under ground effect using artificial intelligence techniques

Estimation of aerodynamic coefficients of a non-slender delta wing under ground effect using artificial intelligence techniques

FTLE and Surface-Pressure Signature of Dynamic Flow Reattachment During Delta-Wing Axial Acceleration

An experimental investigation was conducted on the aerodynamics of an NACA0012 airfoil wing with triangular planform geometry undergoing steady and axial accelerations at as a model of an unmanned combat air vehicles encountering unsteady environments at pre- and poststall angles of attack (Marzanek, M., and Rival, D., “Separation Mechanics of Non-Slender Delta Wings During Streamwise Gusts,” Journal of Fluids and Structures, Vol. 90, Oct. 2019, pp. 286–296.). Ensemble-averaged flowfields, forces, moments, and surface-pressure distributions were measured in the Optical Towing Tank for Energetics Research facility at Queen’s University. To characterize the evolution of the flowfield coherent structures around a wing under axial gusts or accelerations, a Lagrangian flowfield analysis including the finite-time Lyapunov exponent (FTLE) was conducted. Results indicate that axial acceleration can induce a dynamic flow reattachment at the poststall angle of attack, and clear FTLE ridges extend from the wing surface and travel down the chord as reattachment progresses. A distinct signature in the surface-pressure distribution tracks closely with the location where the FTLE ridges meet the wing. The timing of how this surface-pressure distribution moves aft is shown to correlate with the relatively large-scale fluctuation in the pitching moment. The current work reveals the correlation between FTLE and surface pressure, which further hints at a potential approach of dynamic estimation by using Lagrangian coherent structures.

Identification of Sharp Edge Non-Slender Delta Wing Aerodynamic Coefficient Using Neural Network

Delta wing formed a vortical flow on its surface which produced higher lift compared to conventional wing. The vortical flow is complex and non-linear which requires more studies to understand its flow physics. However, conventional flow analysis (wind tunnel test and computational flow dynamic) comes with several significant drawbacks. In recent times, application of neural network as alternative to conventional flow analysis has increased. This study is about utilization of Multi-Layer Perceptron (MLP) neural network to predict the coefficient of pressure (Cp ) on a delta wing model. The physical model that was used is a sharp edge non-slender delta wing. The training data was taken from wind tunnel tests. 70% of data is used as training, 15% is used as validation and another 15% is used as test set. The wind tunnel test was done at angle of attack from 0°-18° with increment of 3°. The flow velocity was set at 25m/s which correspond to 800,000 Reynolds number. The inputs are angle of attack and location of pressure tube (y/cr) while the output is Cp . The MLP models were fitted with 3 different transfer functions (linear, sigmoid, and tanh) and trained with Lavenberg-Marquadt backpropagation algorithm. The results of the models were compared to determine the best performing model. Results show that large amount of data is required to produce accurate prediction model because the model suffer from condition called overfitting.

Yer Etkisi Altındaki Delta Kanat Üzerinde Oluşan Taşıma Katsayısının Yapay Sinir Ağı Kullanılarak Tahmin Edilmesi

The estimation of the lift coefficient, CL of a non-slender delta wing under the ground effect, is performed by employing an artificial neural network (ANN). The purpose of the study is to estimate the lift coefficient, CL acting on the delta wing for the ground distance h/c=0.4 by utilizing the actual lift coefficient, CL for the ground distances h/c=1, 0.7, 0.55, 0.25 and 0.1. In this ANN model, the angle of attack, α and ground distance, h/c were used as input parameters and lift coefficients, CL as the output parameter. While mean absolute percentage error (MAPE) and root mean squared error (RMSE) were found as 1.60% and 0.0114 in the testing stage, they were calculated as 1.77% and 0.01 in the training stage. Hence, this investigation shows that the lift coefficient, CL of the delta wing in ground effect can be correctly estimated by developing an ANN model.

Effect of ground on flow characteristics and aerodynamic performance of a non-slender delta wing

The aerodynamic performance and the structure of vortical flow on a delta wing are affected by the influence of the ground in take-off and landing stages. In this context, the effect of the ground on a delta wing having a sweep angle of 40° was investigated by employing Particle Image Velocimetry (PIV), aerodynamic force measuring system and the dye flow visualization technique. Flow characteristics of delta wing were examined under two different angles of attack, α=8° and 11° and variation of the distance between trailing edge of the wing and ground, h normalized with the root chord length, c of the wing. It was observed that the existence of the ground attenuates the magnitude of peak values of primary and secondary vortices due to incomplete development of vortices. The ground effect caused the outboard movement of the leading-edge vortex in a spanwise direction as well as an increase in the size of vortices. Furthermore, the presence of the ground induced a decrease in Strouhal number, St due to the slowing down of vortex formation. Lift and drag coefficient, CL, and CD of the delta wing were observed to increase with descending from unbounded flight region into ground effect region. Finally, it was found that CL/CD increases by reducing the distance between ground and wing, h/c, and a rise of CL/CD is much more effective under the lower angles of attack, α.

Exploring the signature of distributed pressure measurements on non-slender delta wings during axial and vertical gusts

For a broad range of aerodynamic bodies, vortex structures arising from perturbations such as gusts cause characteristic surface pressure signatures that are coupled to the observed aerodynamic loads. The present study evaluates the extent to which sparsely measured pressure signatures can be used to identify the spatio-temporal evolution of vortex structures and, specifically, their relationship to the bulk aerodynamic loads. A non-slender delta wing experiencing axial and vertical gusts under various initial stall conditions is selected as a test case. Time-resolved loads, distributed surface pressures, and time-resolved flow fields (particle image velocimetry) are collected for a wide range of parameters in a towing-tank facility. By linearly mapping the sparse pressure data to the aerodynamic loads, the spatio-temporal relation of loads and pressure can be extracted. The static mapping coefficients are determined through linear regression at each incidence angle as well as for an angle-independent (aggregate) case. Despite slightly larger errors when compared to the angle-specific fits, the aggregate method maintains a good fit quality over all angles of attack and thereby provides a robust pressure-load mapping. Thus, the existence of a common mechanism across gusts and angles of attack is identified despite the stark differences in flow conditions, i.e., light vs deep dynamic stall. In addition, the lasso regularization used in the study provides valuable insight into sensor reduction. The distribution of fewer regression predictors indicates specific pressure ports that capture the footprint of dominant flow features and thereby suggest sensitive locations for future clusters of sensors.

Corrugation Assisted Enhancement of Aerodynamic Characteristics of Delta Wing for Micro Aerial Vehicle

Some of the existing challenges of fixed-wing micro aerial vehicle designs at low Reynolds number are low lift and reduced aerodynamic efficiency. Delta wings, despite their characteristic maneuverability, are rarely used for micro aerial vehicles, perhaps due to the limited lifting capacity because of the reduced aspect ratio. Corrugations on insect wings are ubiquitous and augment the aerodynamic performance in a similar Reynolds number regime. Herein, corrugations are applied to enhance the lift and aerodynamic efficiency of a delta wing. Three delta-wing geometries are examined using numerical simulations at a Reynolds number of : a nonslender delta wing with 53 deg leading-edge sweep and two variants with corrugations in the rear planform. The simulations reveal that the variants potentially offer approximately 11 and 27% increments in the aerodynamic efficiency and the lift, respectively, over the conventional delta wing at an angle of attack of 5 deg. Moreover, the corrugations strengthen the leading-edge vortices and generate an additional nosedown moment about the leading edge. The introduction of corrugations on micro aerial vehicle wings could be a passive technique to potentially improve the aerodynamic characteristics, without any significant weight or drag penalty.

Control of flow structure over a non-slender delta wing using passive bleeding

Recently, it has been demonstrated that the bleeding, which utilizes passages inside the wing to allow the fluid flowing from the pressure side to the suction side by using inherent pressure difference, could be used as an effective flow control method for non-slender delta wings. In the present study, this technique is applied to a non-slender delta wing with sweep angle of Λ=45 deg with particular interests in delaying stall and eradication of three-dimensional surface separation. The experiments are conducted in a low speed wind tunnel using surface pressure measurements, Particle Image Velocimetry (PIV) measurements on near surface and cross flow planes, and aerodynamic force measurements. The bleeding passage orientation defined as back angle θ and bleeding passage size defined as bleed opening ratio bor are varied and the corresponding wings along with the Base planform are tested for the Reynolds numbers 35000≤Re≤100000 and the angles of attack 0≤α≤36 deg. The results indicate that at sufficiently high angle of attack where the pronounced surface separation appears on the Base planform, which is indicated by focus point along with large-scale swirl in the near surface streamline pattern, the elimination of surface separation is achieved with passive bleeding. This is incorporated with significant increases in the magnitudes of surface normal vorticity and suction pressure coefficient −Cp, which indicate recovery of leading-edge vortex. Considering the back angle and bleed opening ratio, the results indicate that increases in both parameters provide better performances in terms of elimination of surface separation. In addition, aerodynamic coefficient measurements show that proper bleeding configuration provides significant improvement in stall angle and lift coefficients.