Delta Wing Research Articles

The hydrodynamic performance of duck feet for submerged swimming resembles oars rather than delta-wings

Waterfowl use webbed feet to swim underwater. It has been suggested that the triangular shape of the webbed foot functions as a lift-generating delta wing rather than a drag-generating oar. To test this idea, we studied the hydrodynamic characteristics of a diving duck’s (Aythya nyroca) foot. The foot’s time varying angles-of-attack (AoAs) during paddling were extracted from movies of ducks diving vertically in a water tank. Lift and drag coefficients of 3D-printed duck-foot models were measured as a function of AoA in a wind-tunnel; and the near-wake flow dynamics behind the foot model was characterized using particle image velocimetry (PIV) in a flume. Drag provided forward thrust during the first 80% of the power phase, whereas lift dominated thrust production at the end of the power stroke. In steady flow, the transfer of momentum from foot to water peaked at 45° < AoA < 60°, due to an organized wake flow pattern (vortex street), whereas at AoAs > 60° the flow behind the foot was fully separated, generating high drag levels. The flow characteristics do not constitute the vortex lift typical of delta wings. Rather, duck feet seem to be an adaptation for propulsion at a wide range of AoAs, on and below the water surface.

Shock migration on an oscillating straked delta wing using an unsteady Euler solver

Many investigations, including flight tests, wind tunnel tests, and computational methods have attempted to determine the drivers of limit cycle oscillations (LCO) on the F-16 fighter aircraft. One area of interest has focused on shock-induced trailing-edge separation (SITES) as a driver of limit cycle oscillations. The investigation presented herein aims to expand the understanding of unsteady, nonlinear aerodynamics by studying the flow over a wing undergoing pitch oscillations using Euler computational fluid dynamics with a boundary layer coupling scheme capable of estimating viscous flow effects within the boundary layer. A straked NACA 64A204 delta wing model was used for this analysis. Parameters such as trim angle of attack, pitch amplitude, and Mach number were varied to create 71 unique test cases. Results demonstrate that the same single shock systems at low angles of attack and two shock systems at moderate to high angles of attack that were observed by Cunningham in his wind tunnel tests were also present in this investigation. Patterns of significant shock movement and disappearance were observed within the transonic flow region. Hysteresis of the pitching moment was also found to be present. These findings support the theories that shock movement and SITES may play significant roles in driving LCO.

Delta wing rocking motion suppression by deep reinforcement learning

The wing rock phenomenon is a self-excited roll motion occurring on the aircraft at high angles of attack (AoAs), which negatively impact safety and maneuverability. The limit cycle oscillation is a typical characteristic of this self-excited roll motion. A controller employing a model-free training approach is constructed on the basis of deep reinforcement learning (DRL) to address this severe nonlinear control problem. For an 80° swept delta wing model, the analytical model describing the nonlinear behavior of wing rock is presented and implemented in the simulation environment. The DRL-based controller is trained with the proximal policy optimization algorithm. A reward function is carefully designed for training to achieve excellent convergence stability. Various simulations were conducted at a series of unlearned initial conditions (roll angles, AoAs) to demonstrate the effectiveness and generalization capability of the proposed method and verify and evaluate the performance of the trained DRL controller. Finally, the scenario that undergoes disturbance is added to confirm the effectiveness and robustness of the proposed DRL-based controller. Results show that the DRL-based controller could effectively regulate the oscillation and retain the capability to suppress the un-wanted behavior with favorable generalization capability and robustness. Therefore, delta wing rocking motion suppression based on DRL has powerful intelligent control performance and provides a new idea for predicting the evolution of aerodynamics in rocking motion suppression.

An Aerial–Aquatic Robot with Tunable Tilting Motors Capable of Multimode Motion

Aerial–aquatic robots have promising applications due to their ability to operate in air and water with different motion modes. However, challenges such as water/air transition and multimedia movement still limit their development. Inspired by the aerial–aquatic behavior of wild ducks, an aerial–aquatic robot with tunable tilting motors and a delta wing is introduced. The robot can perform multimode motion, including flying in air, swimming on the water surface, transitioning between air and water, and diving underwater. Besides, the buoyancy design enables the robot to float at the air–water interface without consuming energy. Herein, the performance of the robot, including vertical and horizontal cross‐media performance, power consumption, and endurance in various motion modes, is evaluated. The field experiments show that the proposed robot possesses multimode motion and high maneuverability. This study provides a new idea for the structural design of next‐generation practical aerial–aquatic robots.

Experimental and numerical analysis of the aerodynamics and vortex interactions on multi-swept delta wings

The flow field around a generic multi-swept delta wing configuration is investigated under transonic flow conditions, both experimentally and numerically. A special focus is on the analysis of vortex/vortex and vortex/shock interactions at moderate angles of attack. In the present study, the Mach number is varied between textrm{Ma} = {0.50} and textrm{Ma} = {1.41} and the angle of attack is varied between alpha = 8^circ and alpha = 28^circ. Numerical results are validated using experimental surface pressure data from pressure taps, as well as forces and moments based on strain gauge measurements. For selected cases, velocity field data from particle image velocimetry (PIV) measurements are available as well. Over a broad range of angle of attack and Mach number, strong vortex/vortex interactions, including vortex braiding and vortex merging, occur. The location of vortex merging is moving downstream with increasing angle of attack and increasing Mach number. Additionally, at textrm{Ma} = {0.85}, vortex/shock interaction occurs above the wing. For moderate angles of attack, shock-induced vortex breakdown is observed.

Flow Field Investigation on a Double Delta Wing

In this paper flow field characteristics of double delta wing have been studied both experimentally and numerically at different angles of attack ranging from 0° to 30° at a Reynolds number of 2 x 10&lt;sup&gt;5&lt;/sup&gt; based on the root chord length of the wing model. Surface oil flow visualization has been studied on the leeward surface of a sharp edged single beveled 76°/40° cropped double delta wing at subsonic wind tunnel facility. The flow physics is further validated with the help of surface pressure measurement in the context of surface flow pattern. Subsequently CFD (Computational Fluid Dynamics) analysis is carried out using ANSYS Fluent. Comparisons of both qualitative as well as quantitative measurement with experimental techniques and computational work have been carried out. For computational work, the computational domain is meshed with unstructured grids and SA-turbulence model is used.

An adaptive law based online system identification of cropped delta unmanned aerial vehicles from flight tests

Online system identification of an aircraft has multifaceted advantages, such as real-time health monitoring, fault detection and reconfiguration of control, to name a few. Researchers have utilized variants of the equation error method, which involves non-gradient-based computation to carry out online system identification, due to its feasibility for real-time implementation. However, it was found to be underperforming while handling high angles of attack as well as time-varying bias parameters. In the current manuscript, an attempt has been made to alleviate the aforementioned limitation by proposing a novel online system identification technique based on adaptive law and its applicability is demonstrated by using compatible flight data sets, acquired from two cropped delta wing unmanned aerial vehicles, pertaining to linear, moderately nonlinear, and near stall flight regimes. It is observed that the proposed method is able to predict unknown aerodynamic derivatives, irrespective of flight envelopes, if dynamic modes are adequately excited. Confidence in the estimated aerodynamic parameters is assured with the two-sigma error bound. It is noticed that the maximum relative standard deviation in the estimates obtained with the proposed method is less than the recursive least squares method and higher than the sequential least squares-recursive Fourier transform method in the linear flight regime. Simulated responses using the estimates obtained from aforementioned methods are compared with measured flight data, and it is witnessed that the simulated nondimensional lift coefficient with the proposed method has a relative offset of less than 34.2% w.r.t measured value in near-stall flight regimes, which is least among all three methods.

ANALYSIS OF VORTEX ON SHARP-EDGED DELTA WING WITH BLOWING EFFECT

The formation of the primary vortex above the wing is dependent on the Reynolds number, the angle of attack, the Mach number, and the bluntness of the leading edge. This experiment is done in order to assess the impacts of blowing at three sites on the delta wing's flow topology. The blower was placed on the leading edge at the 15%, 50%, and 70% wing sections. The experiment was conducted at a Reynolds number of 8 x 105, which corresponds to a wind speed of 25 m/s and a mean aerodynamic chord. The UTM generic delta wing model with a 55° sweep angle was employed, and the experiment was conducted at the Low-Speed Wind Tunnel at UniversitiTeknologi Malaysia (UTM-LST). During each experiment, the model was subjected to surface pressure monitoring techniques. This project's findings show the influence of blowing on the flow topology above the wing.In specific test settings, the installation of the blower at positions 50 percent significantly increased the size of the primary vortex while delaying its breakdown at greater angles of attack.

Dynamic Structural Scaling Concept for a Delta Wing Wind Tunnel Configuration Using Additive Manufacturing

Considering aeroelastic effects plays a vital role in the aircraft design process. The construction of elastic wind tunnel models is a critical element in the investigation of occurring aeroelastic phenomena. However, the structural scaling between full-scale and reduced-scale configurations is a complex design and manufacturing task and is usually avoided in wind tunnel testing. This work proposes a numerical approach for a dynamic aeroelastic scaling technique, which is applied to a fictive delta wing configuration. This scaling methodology is designed to optimise the structural layout of wind tunnel models with an integrated rib and spar structure to meet the behaviour of a realistic full-scale equivalent. For the modelling approach of the wing structure, a beam and shell structure is utilised. The applied scaling laws for the relevant quantities and the applied procedures are described. Computational fluid dynamics (CFD) calculations are performed by solving the Reynolds-averaged Navier–Stokes (RANS) equations for the assumption of a rigid full-scale and down-scaled wing. These calculations are used to verify the aerodynamic scaling assumptions, which are applied to the scaling procedure of the wind tunnel model. Global aerodynamic coefficients are evaluated for a variety of angles of attack. The local flow phenomena of the full-scale and the scaled model are compared in more detail for a medium and a high angle of attack. The pressure coefficient distribution shows a proper accordance for the full-scale and the scaled model. To verify the results of the structural scaling optimisation, a high-fidelity structural full-scale model is compared with the scaled model using the ELFINI FEM solver. Therefore, all structural components are modelled by 2D elements. The results for the reduced eigenfrequencies and according modes of the full-scale and the scaled model show a high level of similarity. A static deformation of the structural grids is performed by applying the aerodynamic loads from the CFD simulations. The results show that the deviation of the nondimensional deformation between the scaled and the full-scale model is negligible. Consequently, the applied scaling methodology proves to be a valuable tool for the conceptual approach of designing aeroelastically scaled wind tunnel models considering 3D-printed material.

Mathematical Modelling and Fluidic Thrust Vectoring Control of a Delta Wing UAV

Pitch control of an unmanned aerial vehicle (UAV) using fluidic thrust vectoring (FTV) is a relatively novel technique requiring no moving control surfaces, such as elevators. In this paper, the authors’ previous work on the characterization of a static co-flow FTV rig is further validated using the free to pitch dynamic test bench. The deflection of a primary jet after injection of a high-velocity secondary jet was captured using the tuft flow visualization technique, along with the experimental recording of subsequent normal force impinged on the Coanda surface resulting in the pitching moment. The effect of primary and secondary flow velocities on exhaust jet deflection, as well as on the pitch angle of the aircraft, is examined. Aerodynamic gain as well as the inertia of a delta wing UAV test bench are computed through experiments and fed into the equation of motion (e.o.m). The e.o.m developed aided in the design of a model-based PID controller for pitch motion control using the multi-parameter root locus technique. The root locus tuned controller serves as a benchmark controller for performance evaluation of the genetic algorithm (GA) and particle swarm optimization (PSO) tuned controllers. Furthermore, the frequency domain metric of gain and phase margins were also employed to reach a robust control design. Experiments conducted in a simulation environment reveal that PSO-PID results in a better response of the UAV in comparison to the baseline pitch controller.

Computational study of a conceptual re-entry vehicle's aerodynamic characterization at Mach number-2

Re-entry vehicles are becoming an important topic of study in recent times due to the interest of many countries in space endeavor. They are incorporated with double-delta wings having many profiles, different strake and wing swept angles. A conceptual re-entry vehicle having a double delta wing with NACA0020 symmetrical aerofoil at the top surface and the reflex aerofoil at the bottom is studied. Here, the main objective is to examine the effect of viscosity on the aerodynamic characteristics of the vehicle by varying the angle of attack from 10°to 50°at a free stream Mach number of 2. The steady-state Reynolds Averaged Navier-Stokes equations with Spalart-Allmaras turbulence model and Euler equations are solved using ANSYS Fluent for examining the contribution of the viscous term on flow past the re-entry vehicle. The non-dimensional coefficients of lift, drag, pressure, Mach number, and density gradients obtained from both the models are compared. It is observed that the coefficients of lift and drag obtained from Euler simulations differed significantly from the Spalart-Allmaras turbulence model at all angle of attacks. The maximum value of lift coefficient is seen at 40°in Euler simulation and it appeared at 35°in Spalart-Allmaras model. Both the models have predicted the zones of attached flow, separated flow, leading-edge vortices and the shock pattern over the vehicle. But the flow transition and vortex breakdown are observed earlier in Spalart-Allmaras model compared to inviscid simulations. It suggested that the viscosity and fluctuation in the upstream flow has a significant role in the aerodynamic characteristics of re-entry vehicles.

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.

Structural design and dynamic analysis of an inflatable delta wing

Structural stiffness and strength, as well as the maintenance of wing shape are very important for the usage of inflatable wings. Firstly, inspired by the deployable dorsal fin of fish, structural design on an bionic inflatable delta wing is proposed. The outer flexible surface of delta wing can be stretched through the interval inflatable airbeams to unfold the spanwise-rotatable deployable ribs to achieve the maintenance of wing shape. Then, a dynamic theoretical model of equivalent multi-beam-springs array is established to describe the interval inflatable airbeams and the ribs by the Lagrange equations. The multi-point excitation modal test method is also used to identify the modal of the inflatable delta wing. Moreover, the transient response is analyzed by the wavelet spectrum for shock excitation. The equivalent multi-beam-springs array model is verified by experimental testing for inflatable structural framework of delta wing. Finally, the effects of airbeam spacing and internal pressure on the structural frequency of inflated wing are predicted by numerical simulation. These results can provide a new way for the design of the inflatable delta wing.

Analysis of the vortex-dominated flow field over a delta wing at transonic speed

AbstractThe present work provides an advancement in the prediction of delta wing flow and an improved understanding of various flow physical phenomena which occur over the wing in transonic conditions. Scale-resolving simulations of the vortex-dominated flow around a sharp leading-edge VFE-2 wing have been performed using the SA-based IDDES model. The complex leading-edge vortex pattern with embedded shocks and subsequent shock-vortex interaction is investigated. A promising accuracy has been achieved using the high-fidelity flow field data provided by the scale-resolving simulation results. Besides the assessment of sensitivity to spatial and temporal resolution, physical aspects are presented, which are not accessible in experimental data in such detail and require scale-resolving simulation approaches. This includes the observation of the vortex system and the shocks in the fully three-dimensional flow field data. Finally, turbulence-related quantities such as eddy viscosity and resolved Reynolds-stresses and their behaviour during the vortex formation and sustaining process are analysed.

Synthesis of fluorinated copolymers and their applications as pressure‐responsive materials

AbstractPressure‐sensitive paints (PSPs) have been demonstrated to be very important for aerodynamics research. In order to improve the pressure sensitivity and adhesion ability of PSP, herein, a series of fluorine‐containing copolymers compositing of an oxygen and pressure‐responsive probe derived from platinum(II)‐5,10,15,20‐tetrakis‐(2,3,4,5,6‐pentafluorophenyl)‐porphyrin (PtTFPP) named as poly[(tBs)a‐co‐(TFEM)b‐co‐(PtTFPP‐1MA)c] were synthesized and investigated as PSPs. These PSPs were constructed with different fluorine ratios for the investigation of the effect of fluorine contents on PSPs. Especially, PSP (F0.33) with the trifluoroethyl methacrylate (TFEM) molar ratio of 33% in the copolymers could adhere well to the surface of an aircraft model. F0.33 showed high oxygen sensitivity (I0/I100 = 32.87), reasonable response time to oxygen (44 s), high pressure sensitivity (SP = 0.75%/kPa at 20°C) and low temperature response (ST = −0.77%/°C at 100 kPa). F0.33 was further sprayed on the surface of a delta wing of an aircraft model and its pressure response was verified in a wind tunnel. The data from F0.33 film at the same wind speed and position were close to the pressure data measured by using traditional method using open pressure holes, showing the potential application of these PSPs.

Flow Field Investigation of a Blended Wing Body in Low Speeds

With the advancement in the modern transport system, it has become imperative to achieve larger aircraft volume at minimum cost wherein aerodynamics plays a major role. The present work aims to investigate the flow over a typical blended wing body (BWB) configuration at different angles of attack adopting experimental and computational techniques. Experiments consisted of the force measurements and oil flow visualizations at a free stream velocity of around 19 m/s. Computations were made using the commercially available software ANSYS 18.1. Results indicated that the use of a Blended Wing Body Configuration resulted in improved aerodynamic characteristics in comparison to other wing configurations such as Delta/ Double delta wings. Surface flow visualization carried out over the BWB configuration indicated the reason behind increased lift coefficients. Reasonable agreement of experiments and computations was observed.

Flow Visualization of Wing Vortices of Delta Wings

Photo Gallery Flow Visualization of Wing Vortices of Delta Wings Philipp Epple, Philipp Epple Coburg University of Applied Sciences, Department of Mechanical Engineering, Coburg, Germany e-mail: philipp.epple@hs-coburg.de Search for other works by this author on: This Site PubMed Google Scholar Dustyn Kramer, Dustyn Kramer Coburg University of Applied Sciences, Department of Mechanical Engineering, Coburg, Germany e-mail: Dustyn.Kramer@stud.hs-coburg.de Search for other works by this author on: This Site PubMed Google Scholar Michael Steppert Michael Steppert Coburg University of Applied Sciences, Department of Mechanical Engineering, Coburg, Germany e-mail: michael.steppert@hs-coburg.de Search for other works by this author on: This Site PubMed Google Scholar Author and Article Information Philipp Epple Coburg University of Applied Sciences, Department of Mechanical Engineering, Coburg, Germany Dustyn Kramer Coburg University of Applied Sciences, Department of Mechanical Engineering, Coburg, Germany Michael Steppert Coburg University of Applied Sciences, Department of Mechanical Engineering, Coburg, Germany e-mail: philipp.epple@hs-coburg.de e-mail: Dustyn.Kramer@stud.hs-coburg.de e-mail: michael.steppert@hs-coburg.de J. Fluids Eng. Jun 2023, 145(6): 060904 (1 pages) Paper No: FE-23-1073 https://doi.org/10.1115/1.4057070 Published Online: March 10, 2023 Article history Received: February 14, 2023 Revised: February 22, 2023 Accepted: February 22, 2023

Aerodynamic characterisation of delta wing unmanned aerial vehicle using non-gradient-based estimator

AbstractAerodynamic characterisation from flight testing is an integral subroutine for evaluating a new flight vehicle’s aerodynamic performance, stability and controllability. The estimation of aerodynamic parameters from flight test data has extensively been explored, in the past, using estimation methods such as the equation error method, output error method and filter error method. However, in the current era, non-gradient-based estimation techniques are gaining attention from researchers due to their inherent data-driven optimisation capability to find the global best solution. In this paper, a novel non-gradient-based estimation method is proposed for the aerodynamic characterisation of unmanned aerial vehicles from flight data, which relies on the maximum likelihood method augmented with particle swarm optimisation. Flight data sets of a wing-alone unmanned aerial vehicle are used to demonstrate the capabilities of the proposed method in estimating aerodynamic derivatives. Estimates from the proposed method are corroborated with the wind tunnel test and output error method results. It has been observed that simulated flight vehicle responses using estimated parameters are in good agreement with measured data in most of the manoeuvers considered. Confidence in the estimates of linear and nonlinear aerodynamic parameters is well established with the lower limit of Cramer-Rao bounds, which are minimal. The proposed method also demonstrates good predictability of the quasi-steady stall aerodynamic model by estimating stall characteristic parameters such as aerofoil static stall characteristics parameter, hysteresis time constant and breakpoint. The overall performance of the proposed estimation method is on par with the output error method and is validated with the proof-of-match exercise.

Numerical Simulation of Surface Pressure of a Wedge at Supersonic Mach Numbers and Application of Design of Experiments

This paper aims to numerically simulate the flow field for wedge, cone, and ogives. Usually, wedge shapes are used as a stabilizing surface for supersonic projectiles like rockets, missiles, and planes in defense applications. Wedge-shaped or delta wings are often the principal wing used for the stability of supersonic aircraft/missiles. The main goal of the current study is to estimate the pressure at the wedge-shaped plane's/missiles nose at different Mach numbers and incidence angles. Analytical pressure distribution is determined using the strip and piston theory. Later the outcomes from the numerical simulations are compared with the results obtained analytically. The analytical and CFD findings matching is very good. The findings demonstrate that the Mach number and wedge angle are the variables that influence the deviation of static pressure. The static pressure on the surface of the wedge grows with the rise in the semi-vertex angle of the wedge angle and the Mach number. This increase in the surface pressure ratio is linear for the increase in Mach number as well as the semi-vertex angle of the wedge. However, the magnitude of increase with the Mach numbers is not the same as what it was for low Mach numbers. The enhancement in the pressure decreases with the Mach numbers.

Parametric model identification of delta wing UAVs using filter error method augmented with particle swarm optimisation

AbstractFrom arsenal delivery to rescue missions, unmanned aerial vehicles (UAVs) are playing a crucial role in various fields, which brings the need for continuous evolution of system identification techniques to develop sophisticated mathematical models for effective flight control. In this paper, a novel parameter estimation technique based on filter error method (FEM) augmented with particle swarm optimisation (PSO) is developed and implemented to estimate the longitudinal and lateral-directional aerodynamic, stability and control derivatives of fixed-wing UAVs. The FEM used in the estimation technique is based on the steady-state extended Kalman filter, where the maximum likelihood cost function is minimised separately using a randomised solution search algorithm, PSO and the proposed method is termed FEM-PSO. A sufficient number of compatible flight data sets were generated using two cropped delta wing UAVs, namely CDFP and CDRW, which are used to analyse the applicability of the proposed estimation method. A comparison has been made between the parameter estimates obtained using the proposed method and the computationally intensive conventional FEM. It is observed that most of the FEM-PSO estimates are consistent with wind tunnel and conventional FEM estimates. It is also noticed that estimates of crucial aerodynamic derivatives ${C_{{L_\alpha }}},\;{C_{{m_\alpha }}},\;{C_{{Y_\beta }}},\;{C_{{l_\beta }}}$ and ${C_{{n_\beta }}}$ obtained using FEM-PSO are having relative offsets of 2.5%, 1.5%, 6.5%, 3.4% and 7.6% w.r.t. wind tunnel values for CDFP, and 1.4%, 1.9%, 0.1%, 9.6% and 7.5% w.r.t. wind tunnel values for CDRW. Despite having slightly higher Cramer-Rao Lower Bounds of estimated aerodynamic derivatives using the FEM-PSO method, the simulated responses have a relative error of less than 0.10% w.r.t. measured flight data. A proof-of-match exercise is also conducted to ascertain the efficacy of the estimates obtained using the proposed method. The degree of effectiveness of the FEM-PSO method is comparable with conventional FEM.