High Angles Of Attack Research Articles

Comparative Study of Machine Learning Modeling for Unsteady Aerodynamics

Modern fighters are designed to fly at high angle of attacks reaching 90 deg as part of their routine maneuvers. These maneuvers generate complex nonlinear and unsteady aerodynamic loading. In this study, different aerodynamic prediction tools are investigated to achieve a model which is highly accurate, less computational, and provides a stable prediction of associated unsteady aerodynamics that results from high angle of attack maneuvers. These prediction tools include Artificial Neural Networks (ANN) model, Adaptive Neuro Fuzzy Logic Inference System (ANFIS), Fourier model, and Polynomial Classifier Networks (PCN). The main aim of the prediction model is to estimate the pitch moment and the normal force data obtained from forced tests of unsteady delta-winged aircrafts performing high angles of attack maneuvers. The investigation includes three delta wing models with 1, 1.5, and 2 aspect ratios with four determined variables: change rate in angle of attack (0 to 90 deg), non-dimensional pitch rate (0 to .06), and angle of attack. Following a comprehensive analysis of the proposed identification methods, it was found that the newly proposed model of PCN showed the least error in modeling and prediction results. Based on prediction capabilities, it is seen that polynomial networks modeling outperformed ANFIS and ANN for the present nonlinear problem.

Initial development of the hybrid semielliptical-dolphin airfoil

Iosif Taposu has formulated a mathematical model and generated a family of airfoils whose geometry resembles the dolphin shape. These airfoils are characterized by a sharp leading edge and experiments have proven that they can achieve better aerodynamic characteristics at very high angles of attack than certain classical airfoils, with the nose geometry inclined downwards. On the other hand, they have not been applied to any commercial general aviation aircraft. The authors of this paper have been motivated to compare the aerodynamic characteristics of widely used NACA 2415 airfoil with Taposu?s Dolphin that would have the same principal geometric characteristics. A CFD calculation model has been established and applied on NACA 2415. The results were compared with NACA experiments and very good agreements have been achieved in the major domains of lift and polar curves. The same CFD model has been applied on the counterpart Dolphin 2415. Results have shown that the Dolphin has a slightly higher lift/drag ratio in the lift coefficient domain 0.1-0.35 than NACA. On the other hand, at higher and lower lift coefficients, its aerodynamic characteristics were drastically below those of the NACA section, due to the unfavorable influence of the Dolphin?s sharp nose. A series of the Dolphin?s leading edge modifications has been investigated, gradually improving its aerodynamics. Finally, version M4, consisting of about 70% of Dolphin?s original rear domain and 30% of the new nose shape, managed to exceed the NACA?s characteristics, thus paving the way to investigate the Dolphin hybrids that could be suitable for the general aviation industry.

Multiplanes of the New and Old Worlds

This article for the first time expounds on the history of development and construction of multiplanes (having at least four wings) worldwide before the beginning of World War I. Five ideas are identified that led the designers to what at the first glance appeared to be an irrational design of a multiplane. These were not only increasing the total wing area and a desire to increase the wing aspect ratio to enhance the lift / drag ratio but also using the lattice wings to prevent air flow along the wingspan, tandem wing boxes to increase longitudinal stability, and a design of tandem monoplane on a moving frame to increase stability at high angles of attack and ensure efficient pitch control. The article lists 39 multiplane models built during that period, and specifies the ideas that guided their designers. It is shown that the interest in multiplanes in the USA and the UK was disproportionally high compared to other countries. It is proposed what it could have been associated with. It is stated for the first time that there has been not one but two different J. Zerbe’s quintaplanes. We provide an example of how statistical analysis of technical designs can help in the evaluation of the extent of originality of a country’s construction school.

Utilizing the Magnus effect to produce more downforce than a standard wing

Wings on cars help keep the vehicle grounded at high speeds and improve traction by producing downward force known as downforce. To generate more downforce, the wings generally need a high angle of attack which leads to increased drag. A car’s performance is degraded with drag by the increase in air resistance, so a new method of generating downforce with better drag performance could be beneficial. The Magnus effect is the tendency of a spinning cylinder or sphere to produce force perpendicular to the flow of the air. Cylinders undergoing the Magnus effect have been found to have less drag than comparable wings and can generate lift and downforce with sufficient airflow. Thus, we saw it fit to test if a spinning cylinder was capable of generating more downforce than a typical wing via the Magnus effect. Based on the success of the Magnus effect for lift, we hypothesized similar outcomes when testing for downforce. We tested for downforce by attaching a motor-driven cylinder to a scale and used a leaf blower to simulate high wind speeds. We also tested a standard wing as a control. Overall, we found that at all speeds, the cylinder was significantly more effective at producing downforce with nearly 50% more force produced at the highest velocity while outperforming the wing by larger margins at lower speeds. Our experiments demonstrated that cylinders could be a potential replacement for the wing when downforce is a priority.

Experimental investigation on effect of partial flexibility at low aspect ratio airfoil – Part I: Installation on suction surface

Effects of flexible membrane mounted over suction surface of NACA 4412 airfoil were experimentally investigated at Reynolds number of 5x104 and low aspect ratio (AR=1) in this paper. The smoke-wire visualization method has been performed for flow visualization to demonstrate flow phenomena as laminar separation bubble (LSB), leading edge separation at z/c=0.4 and tip vortices at z/c=0.1. Values of velocity, Reynolds stress and turbulence statistics were measured by means of a constant temperature anemometer (CTA) system. Results of smoke-wire experiment revealed that size and height of LSB formed along z/c=0.4 at lower angles of attack such as α=8° was mitigated. Moreover, stall phenomenon as a result of boundary layer separation was apparently postponed at higher angles of attack. Velocity value was increased, whereas values of Reynold stress and turbulent kinetic energy was decreased with reduction of amount of fluctuations in flow. Consequently, using flexible membrane over suction surface of airfoil allowed the LSB to be mitigated or extinguished, resulting in exhibition of more stable flow characteristics.

Effect of surface roughness on aerodynamic performance of the wing with NACA 4412 airfoil at Reynolds number 1.7 × 105

Drag reduction is of paramount importance in an effort to enhance the aerodynamic performance of any aerial vehicle. Wings of avian and fins of aquatic animals which utilize their skin roughness to move faster beating the drag are the main source of conviction. This study conducts an experimental investigation to examine the effect of skin friction on the aerodynamic characteristics of a wing by inducting surface roughness at different chord locations (30% and 45% of the chord) of the wing. As a matter of course, wind tunnel tests were performed by fixing emery paper of two different grits (E480 and E4220) in chord-wise direction over two different chord positions on the upper side of the airfoil. The experiments were carried out on a wing having airfoil NACA 4412 at Reynolds number 1.7 × 105. Results revealed that the different surface roughness at different chord locations provide significant variation in lift and drag characteristics and also stall attributes. Surface roughness on the wing surface is found to play a major role in increasing the drag and delaying the stall separation. During takeoff and landing, the aircraft needs to fly at higher angles of attack in comparison to the cruising condition. The surface roughness can be activated only during the takeoff and landing phases by using a servo mechanism for the future aircraft applications.

Structural and CFD analyses of a reciprocating-airfoil (RA) driven UAV wing under maximum lift and inertia forces

The reciprocating-airfoil (RA) driven vertical take-off and landing (VTOL) aircraft is a new aircraft concept that utilizes the reciprocating motion of the wings to provide lift for take-off and landing. The RA wings are shaped like the wings of a fixed-wing airplane and work as fixed wings while cruising. The wing undergoes substantial linear motion during take-off and may generate lift similar to a fixed-wing aircraft. The unique structural characteristics of reciprocating wings are their high inertia and lifting force loadings. This study aims to conduct an internal structural analysis of the wing under the maximum lift and inertia force to validate the wing’s performance. The reciprocating motion of the wing in a stroke was analyzed to determine its maximum speed in the stroke and its inertia force loading in conjunction with a reciprocating driver. A 3D computational fluid dynamic (CFD) analysis was conducted at the highest angle of attack (AoA) and the determined maximum speed to obtain maximum lift and drag. The results of a finite element analysis (FEA) revealed acceptable stresses, demonstrating a safe load-carrying capacity of the wing structure, which may ensure the suitability of the wing for integration with the RA UAV module.

Improved Stall Delay Model for HAWT Performance Predictions using 3D Navier-Stokes Solver and Actuator Disk Method

The Actuator Disk Method (ADM), in its analytical formulation or combined with Navier-Stokes equations, is widely used for design and/or for aerodynamic analysis of Horizontal Axis Wind Turbines (HAWT). This method has demonstrated its capabilities for performance predictions of HAWT rotors for limited range of wind speeds with lower angles of attack values, i.e. attached flow conditions. However, for typical wind speeds that rotor encounters, under higher angles of attack i.e. stall conditions, such a method cannot describe accurately the flow characteristics around rotor-blades due to severe flow separations coupled with the effects of blades rotation as well as the radial flow over the blades. In this paper, original correction approaches have been proposed for the existing stall delay models to take into account both the blade rotation and the radial flow effects over the rotor blades. For this purpose, the ADM combined with 3D- NavierStokes equations formulation using Large Eddy Simulations (LES) model has been considered to describe the incompressible turbulent flow field around HAWT rotor blades. The resulting mathematical model has been solved using a 3D in-house subroutine developed with OpenFOAM code. The proposed numerical method has been validated against the well recognized reference measurements obtained using the New MEXICO and the NREL Phase VI experimental wind turbines. In comparison with existing stall delay models, the proposed correctionapproaches, especially the radial flow approach, have shown noticeable enhancements on performance predictions of HAWT rotors compared to the experimental measurements. It has been found very low discrepancies to the experimental torque and thrust values, up to 1% and 10% have been recorded respectively.

Experiment on longitudinal aerodynamic characteristics of flying wing model with plasma flow control

Horizontal tail is eliminated from the flying wing layout for improving the low observable and aerodynamic efficiency, resulting in degrading longitudinal maneuverability and fight stability. The low speed wind tunnel test study of improving the longitudinal aerodynamic characteristics of large aspect ratio flying wing model is carried out by using plasma flow control technology. The flying wing model has a leading-edge sweep angle of 34.5° and an aspect ratio of 5.79. The reasons for deteriorating the static maneuverability and stability of the flying wing model and the mechanism of plasma control of the flow field and longitudinal aerodynamic characteristics are studied by particle image velocimetry (PIV) flow visualization and static force measurement test. The control law of plasma control of the flight maneuverability and stability of the flying wing model is studied through flight test. The fact that the flow separation of the outer wing of the flying wing model occurs earlier than the inner wing and the wing is swept back can result in the forward movement of the aerodynamic center and the deterioration of the longitudinal static stability. The shock disturbance induced by plasma can suppress the flow separation of the suction surface, thereby extending the linear section of the lift curve of the model, preventing the aerodynamic center from moving forward, and improving the longitudinal static stability. When the wind speed is 50 m/s, the plasma control improves the horizontal rudder efficiency at a high angle of attack of the flying wing model, increases the maximum lift coefficient of the model by about 0.1, and postpones the stall angle of attack by more than 4° at different rudder angles. The plasma control allows the flying model to follow the command movement better while flying, increases the flying pitch limit angle from 11.5° to 15.1°, reduces the amplitude of longitudinal disturbance motion by 2°, and reduces the oscillation attenuation time from 15 to 8 s, thereby improving the longitudinal flight maneuverability and stability of the flying wing model. It can be seen that plasma flow control technology has great potential applications in improving the flight quality of flying wing layout.

５つの山頂部をもつ平板コルゲート翼の翼表面圧力特性

A two-dimensional unsteady numerical analysis of a corrugated airfoil was performed to clarify the variation of flow around the airfoil as a function of leading-edge angle. A corrugated airfoil with five crests was made to follow the outline of NACA4412, and three types of airfoils with leading edge angles of 25° and 35° were selected as the target airfoils. The Reynolds number was set to 1.0 × 104. Comparison of NACA4412 and the three types of corrugated blades showed that the lift coefficient was highest for NACA4412 at angles of attack of 4° or less and was significant for the corrugated blades at angles of attack higher than that. On the other hand, the drag coefficient was the lowest for NACA4412 at all angles of attack. Comparison of corrugated blades with different leading-edge angles revealed that at low angles of attack of 6° or less, the circulating flow in the first valley has a greater effect on aerodynamic characteristics than that in the second valley. On the other hand, at high angles of attack of 10° or more, flow cancellation was observed from the leading edge to the first crest of the upper surface of the wing, which is one of the reasons for the improved wing characteristics.

دراسة تجريبية لتأثير مولد الدوامة على الطبقة الحدودية للوحة مسطحة

This paper is dealing with an experimental study to show the influence of the geometric characteristics of the vortex generators VG son the thickness of the boundary layer (∂) and drag coefficients (CD) of the flat plate. Vortex generators work effectively on medium and high angles of attack, since they are "hidden" under the boundary layer and practically ineffective at low angles. The height of VGs relative to the thickness of the boundary layer enables us to study the efficacy of VGs in delaying boundary layer separation. The distance between two VGs also has an effect on the boundary layer if we take into account the interference between two pairs of VGs. The effect of the changing in (h- the height of vortex generator, d- the average distance between tow vortex generators) on the thickness of the flat plate boundary layer and the drag coefficients has been studied for triangular vortex generator. The measurements of the vortex generator have been changed to determine the optimum boundary layer thickness and the change in drag coefficients. An experiment was done at an average free stream velocity, (U∞,) of 28 m/s. The experiment was conducted in the wind tunnel UTAD-2 University (NAU) Kiev, Ukraine.

Stall flutter suppression of NACA 0012 airfoil based on steady blowing

Stall flutter is a classical aeroelastic phenomenon that seriously affects the performance of flexible wings at high angle of attack. In this article, steady blowing from the leading edge is designed to suppress stall flutter of a two-dimensional airfoil model at 15° static equilibrium angle of attack. The dynamic responses are measured to analyze the nonlinear characteristics of the aeroelastic system. It is found that the leading-edge blowing changes the local bifurcation behavior from a subcritical Hopf bifurcation followed by a saddle–node bifurcation to a supercritical Hopf bifurcation, with the flutter amplitude decreasing and the flutter critical velocity increasing. Time-resolved particle image velocimetry (TR-PIV) measurements are acquired to observe the flowfield evolution during stall flutter. The unsteady aerodynamic moment and flowfield results show that the leading-edge blowing effectively controls dynamic stall during oscillations. Low-momentum injection promotes shedding of dynamic stall vortices (DSVs) and reduces aerodynamic moment fluctuation, whereas high-momentum injection further suppresses the formation of DSVs and shrinks the aerodynamic moment hysteresis loop. During pitching-down of the airfoil, leading-edge blowing induces rapid recovery of the aerodynamic moment, which promotes the decay of stall flutter. When the blowing excitation is strong enough, stall flutter can be completely suppressed within several cycles.

A new many-objective aerodynamic optimization method for symmetrical elliptic airfoils by PSO and direct-manipulation-based parametric mesh deformation

A new many-objective aerodynamic optimization method is proposed to design the distinctive symmetrical elliptic airfoil applied on the canard rotor wing aircraft with special requirements on both low drag in different Mach numbers and large drag divergence Mach number. A parameterized direct manipulation of free form deformation (DFFD) method is proposed to realize the one-step process on geometry parameterization and mesh deformation. Then an efficient multi and many-objective particle swarm optimization algorithm (LMPSO) is built to optimize the symmetrical elliptic airfoil to comprehensively improve its aerodynamics. The biobjective optimization of drag reductions at hover stage and at cruise flight shows that the new method can significantly reduce the drag of the elliptic airfoil but at the cost of drag increases at high angles of attack. The four-objective optimization results indicate that the consideration of the high angle of attack flight conditions can improve the aerodynamic performance of the airfoil over a large lift range, which is more practical in engineering application. The five-objective optimization presents that the maximization of the drag divergence Mach number of the elliptic airfoil can be significantly promoted when adding three more design conditions at high Mach number. In general, the optimization results demonstrate that the developed new method can effectively improve the aerodynamic performance of symmetrical elliptic airfoils and provide support on the development of the canard rotor wing aircraft.

A new dynamic stall prediction framework based on symbiosis of experimental and simulation data

Dynamic stall requires both accurate and efficient predictions. To model the unsteady aerodynamics of dynamic stall, a symbiosis method for dynamic stall prediction is proposed through fusing experimental data and numerical simulations based on computational fluid dynamics. With only a fraction of wind tunnel test data of the National Advisory Committee for Aeronautics 0012 airfoil, the proposed framework is able to predict the lift and moment coefficients of dynamic stall under different balanced angles of attacks, amplitudes, and reduced frequencies. Results indicate that compared with the Unsteady Reynolds-Averaged Navier–Stokes simulation, the proposed model reduces the prediction error about two to five times. In addition, a posteriori analysis shows that with efficient hyperparameter optimization, the framework can separate the dynamics for attached and separated flows adaptively. The proposed data fusion model provides a way to combine the physics of the dynamic stall phenomenon to prediction models for the aerodynamic loading at high angles of attack.

The Effect of Gurney Flap and Trailing-edge Wedge on the Aerodynamic Behavior of an Axial Turbine Blade

In this research, the effect of Gurney flap and trailing-edge wedge on the aerodynamic behavior of blunt trailing-edge airfoil Du97-W-300 which is equipped with vortex generator is studied. To do this, the role of Gurney flap and trailing-edge wedge on the lift and drag coefficient and also aerodynamic performance of the airfoil is studied. Validation of the numerical model is performed by comparison of the obtained results with those of experiment. Results show that before stall, Gurney flap leads to the increase in the aerodynamic performance in a wider range of angle of attack. Numerical findings reveal that the maximum increment for the aerodynamic performance is obtained at low angle of attack when trailing-edge wedge is employed. It is found that for the highest considered value of Gurney flap and trailing-edge wedge heights, where the highest values for the lift occur, the higher aerodynamic performance at low angle of attack is obtained when trailing-edge wedge is used and at high angle of attack, the Gurney flap results in a higher aerodynamic performance. It is also shown that when high aerodynamic performance is concerned, addition of Gurney flap to the airfoil leads to the higher value for the lift. Doi: 10.28991/HIJ-2021-02-04-03 Full Text: PDF

Development of a passive‐adaptive slat for a wind turbine airfoil

AbstractA passive‐adaptive slat concept was designed to avoid separation in the root region of a horizontal‐axis wind turbine blade. This concept incorporates an autonomously moveable slat device only driven by the aerodynamic forces acting on it without the need for mechanical or electrical actuation. It opens at high local angles of attack to delay the stall angle and closes for small angles of attack to increase the lift to drag ratio of the blade segment. This article describes the development of a passive‐adaptive slat for a DU‐91‐W2‐250 airfoil, which is a segment of the reference rotor blade in the project SmartBlades 2.0. In the course of the passive‐adaptive slat design, the optimization of the slat and its extended position is presented. This is followed by the development of two passive‐adaptive slat kinematics, which are opening and closing the slat passively at different angles of attack. With the designed passive‐adaptive slat the stall of the airfoil is delayed by 20° in incidence and the maximum lift of the airfoil is increased by about 130% at the same time in comparison to the original airfoil. Furthermore, the DU‐91‐W2‐250 airfoil with moveable passive‐adaptive slat has in most conditions a higher climb index and therefore a better aerodynamic performance than the same airfoil with a fixed integrated slat.

Wind tunnel measurements of the aerodynamic characteristics of a 3:2 rectangular cylinder including non-Gaussian and non-stationary features

This paper presents a wind tunnel investigation of the aerodynamic characteristics (force and pressure coefficients) of a static 3:2 rectangular cylinder in smooth flow for angles of attack between −4° and 90° at various values of Reynolds number. In contrast to much of the existing literature, this study shows clear dependence of the mean drag coefficients on Reynolds number. In addition, the variation of aerodynamic parameters with angle of attack is fully mapped for a symmetric section revealing that the peak values of the mean values of drag, lift, moment and Strouhal number do not occur at the same critical angle of attack.The present study also presents the first map for identifying the locations of the reattachment and stagnation points as well as the zones of angle of attack where the flow is separated and attached on a face of the section. The phenomenon of switching flow is observed at the angle of attack 25°, leading to strong non-stationarity and non-Gaussian distributions of the aerodynamic forces and pressure. Another phenomenon, so-called unsteady low-frequency vortex shedding, is also observed at higher angles of attack. These phenomena need to be accounted for when estimating wind loading and aeroelastic instability for these sections.

Aeroelastic mode decomposition framework and mode selection mechanism in fluid–membrane interaction

In this study, we present a global Fourier mode decomposition framework for unsteady fluid–structure interaction. We apply the framework to isolate and extract the aeroelastic modes arising from a coupled three-dimensional fluid–membrane system. We investigate the frequency synchronization between the vortex shedding and the structural vibration via mode decomposition analysis. We explore the role of flexibility in the aeroelastic mode selection and perform a systematic comparison of flow features among a rigid flat wing, a rigid cambered wing and a flexible membrane. The camber effect can enlarge the pressure suction area on the membrane surface and suppress the turbulent intensity, compared to the rigid flat wing counterpart. With the aid of our mode decomposition technique, we find that the dominant structural mode exhibits a chordwise second and spanwise first mode at different angles of attack. The structural natural frequency corresponding to this mode is estimated using an approximate analytical formula. By examining the dominant frequency of the coupled system, we show that the dominant membrane vibration mode is selected via the frequency lock-in between the dominant vortex shedding frequency and the structural natural frequency. From the fluid modes and the mode energy spectra at α=20∘ and 25°, the aeroelastic modes corresponding to the non-integer frequency components lower than the dominant frequency are observed, which are associated with the bluff body vortex shedding instability. The non-periodic aeroelastic behaviors observed at higher angles of attack are related to the interaction between aeroelastic modes caused by the frequency lock-in and the bluff-body-like vortex shedding. Using the mode decomposition analysis, we suggest a feedback cycle for flexible membrane wings undergoing synchronized self-sustained vibration. This feedback cycle reveals that the dominant aeroelastic modes are selected through the mode and frequency synchronization during fluid–membrane interaction to exhibit similar modal shapes in the membrane vibration and the pressure pulsation.

Hydrofoils with Differing Leading-Edge Protuberances: CFD and Experimental Investigations

Leading-edge protuberances on the pectoral fin of humpback whales have been widely adopted to the designs of foils to provide superior lifting characteristics in the post-stall regimes. The present work investigates the lift, drag and flow characteristics of finite-span rectangular hydrofoils having different configurations of two protuberances over the leading edge with NACA 634-021 as the base design section. The results obtained from CFD analyses are validated using lift and drag measurements from experiments. The influence of using a transition-sensitive turbulence model on the results is investigated. It is observed that, in general, a foil with smaller separation between protuberances has better post-stall lift characteristics whereas that with protuberances at larger separation have better pre-stall characteristics. Depending on the separation between them, streamwise vortices are generated from the leading-edge protuberances. The two protuberances can restrict the zone of separation between them at high angles of attack. The influence of Reynolds number on the lifting performance is also investigated.

Aerodynamic Performance and Static Stability Characteristics of Aircraft with Tail-Mounted Propellers

In this combined experimental and numerical study, the propeller–airframe aerodynamic interaction is characterized for an aircraft configuration with propellers mounted to the horizontal tailplane. The contributions of the propeller and airframe to the overall loading are distinguished in the experimental analyses by using a combination of external balance and internal load cell data. Validated computational fluid dynamics simulations are then employed to quantify the interaction at a component level. The results show that the propeller installation shifts the neutral point aft with increasing propeller thrust. For the configuration considered herein, the yawing moment due to sideslip is increased by approximately 10%, independent of the propeller thrust coefficient. The changes in propeller loading due to the airframe-induced flowfield are the dominant factor to change the airframe stability and performance. The prominent installation effects occur at high angle of attack, because in that condition the propeller experiences a significant nonuniform inflow that affects the propeller and tailplane. The relatively large propeller diameter compared with tailplane span leads to a change of the tailplane root vortex that causes the tailplane effectiveness to reduce with an inboard-up rotating propeller.