High Angles Of Attack Research Articles

Research on transition corridor of distributed tilting power configuration

For a distributed tilt-power configuration aircraft, the transition corridor was studied based on limited conditions such as wing lift characteristics and propeller power requirement. By establishing the dynamic model of each component and balancing according to the balance equation, the boundary of the transition corridor of distributed tilting dynamic configuration is determined. The results show that the minimum forward flight speed required for a low Angle of attack is less than that for a high Angle of attack; the wing lift characteristic high-speed boundary and the propeller power limit boundary together form the high-speed boundary of the transition corridor.

Numerical Study on the Flow Characteristics of High Attack Angle around the Submarine’s Vertical Plane

When a submarine encounters an emergency situation, it should take emergency-surfacing actions by moving upward with a large angle of attack in the vertical plane. Previous research has often neglected the effect of vertical plane motion on the lateral force (Fy), rolling moment (Mx), and yawing moment (Mz). To examine the flow characteristics of submarines at high angles of attack on the vertical plane, the SST-DDES method is adopted in conjunction with adaptive mesh refinement (AMR) technology, and the new Omega vortex detection method is employed as the AMR criterion for numerical calculations. The obtained results are then appropriately verified by conducting water tank experiments, and the effects of different angles of attack and heel angles on Fy, Mx, and Mz are methodically examined. The results reveal that, in the flow around the vertical plane of a submarine, the influence of Fy and Mz cannot be ignored. In addition, when the vertical velocity of the hull is greater than 0.6 m/s, the influence of Mx cannot be overlooked either. When the angle of attack on the vertical plane of the submarine is greater than 25°, the effects of Fy, Mx, and Mz cannot be neglected, and the effect of Mz is particularly prominent, with its amplitude close to or greater than the average value of the pitch moment (My). The obtained results reveal that the presence of the heel angle (θ) intensifies the forces on the hull for Fy, Mx, and Mz, and the forces caused by the vertical velocity at Fy, Mx, and Mz cannot be neglected. These findings can provide a mechanical analysis basis for the analysis of nonlinear motion phenomena during submarine surfacing.

Analysis of Vortex Dominated Flow over Double Delta Wing

The maneuverability of fighter aircraft designed using double delta wing configuration at low subsonic speed in addition to under high angle of attacks are highly appealing and fascinating to comprehend. Flow pattern formed by double delta wing help to facilitate lifting force under high angle of attack and at such low subsonic speed. In comparison to delta wing, vortex formations by strake part of wing play an important role to create a persistent vortex core capable of delaying stalling. In this investigation, force coefficients of 75°/45° double delta wing configuration is extracted both from wing tunnel experiment and numerical calculation between 0° to 30° angles of attack at Reynolds number of 1.4 × 105 based on root chord length. It has also been attempted to investigate the flow pattern formation over wing leeward surface, interaction between vortices formed by strake/wing based on Q-criterion and pressure distribution along propagation path of vortices obtained from CFD (Computational Fluid Dynamics) calculations. Implications of interplay between vortices and pressure distribution in lift generation have also been analyzed. It has been found that around 15° angle of attack strake and wing vortices merge and gets more intensified. With higher angle of attack the lift generation significantly diminishes because of vortex bursting

Influence of Surface Roughness Modeling on the Aerodynamics of an Iced Wind Turbine S809 Airfoil

Ice formation on structures like wind turbine blade airfoils significantly reduces their aerodynamic efficiency. The presence of ice on airfoils causes deformation in their geometry and an increase in their surface roughness, enhancing turbulence, particularly on the suction side of the airfoil at high angles of attack. An approach for understanding this phenomenon and assessing its impact on wind turbine operation is modeling and simulation. In this contribution, a computational fluid dynamics (CFD) study is conducted using FENSAP-ICE 2022 R1 software available in the ANSYS package. The objective was to evaluate the influence of surface roughness modeling (Shin et al. and beading models) in combination with different turbulence models (Spalart–Allmaras and k-ω shear stress transport) on the estimation of the aerodynamic performance losses of wind turbine airfoils not only under rime ice conditions but also considering the less studied case of glaze ice. Moreover, the behavior of the commonly less explored pressure and skin friction coefficients is examined in the clean and iced airfoil scenarios. As a result, the iced profile experiences higher drag and lower lift than in the no-ice conditions, which is explained by modifying skin friction and pressure coefficients by ice. Overall, the outcomes of both turbulence models are similar, showing maximum differences not higher than 10% in the simulations for both ice regimes. However, it is demonstrated that the influence of blade roughness was critical and cannot be disregarded in ice accretion simulations on wind turbine blades. In this context, the beading model has demonstrated an excellent ability to manage changes in roughness throughout the ice accretion process. On the other hand, the widely used roughness model of Shin et al. could underestimate the lift and overestimate the drag coefficients of the wind turbine airfoil in icy conditions.

Experimental and Numerical Investigation of the Control of the Flow Structure on Surface Modified Airfoils

In this study, experimental and numerical flow analysis was performed on three different blade profiles with a chord length of 165 mm using passive flow control method. The first of the airfoil is the standard NACA 0018 profile. The second airfoil type has a NACA 0018 profile with a gap in the suction surface. The last airfoil is the NACA 0018 profile which is 66% of the trailing edge cut from the chord length. All airfoil profiles were analyzed at the Reynolds number, Re=2x104, and angles of attack α=0o, 5o, 10o, 12o and 14o in both experiment and numerical studies. The experiments were carried out using the Particle Image Velocimetry (PIV) method in a closed-loop open water channel, and the time-averaged velocity vectors, streamlines, and vorticity contours of the flow field were obtained. Subsequently, numerical analyses were performed using the ANSYS Fluent package program, one of the Computational Fluid Dynamics (CFD) programs used frequently in the literature. The streamlines and pressure contours of the airfoil profiles have been compared visually at the same Re and different angles of attack. In addition, according to the angle of attack of the airfoil profiles, lift coefficient CL, drag coefficient CD, and the ratio of lift coefficient to drag coefficient CL/CD graphs were presented. It has been shown that the gap on the airfoil at high attack angles caused changes in lift (up to 0.7) and drag (up to 0.15). These features can allow these models to be used for different purposes in the aerodynamics field.

Laminar post-stall wakes of tapered swept wings

While tapered swept wings are widely used, the influence of taper on their post-stall wake characteristics remains largely unexplored. To address this issue, we conduct an extensive study using direct numerical simulations to characterize the wing taper and sweep effects on laminar separated wakes. We analyse flows behind NACA 0015 cross-sectional profile wings at post-stall angles of attack $\alpha =14^\circ$ – $22^\circ$ with taper ratios $\lambda =0.27$ – $1$ , leading-edge sweep angles $0^\circ$ – $50^\circ$ and semi aspect ratios $sAR =1$ and $2$ at a mean-chord-based Reynolds number of $600$ . Tapered wings have smaller tip chord length, which generates a weaker tip vortex, and attenuates inboard downwash. This results in the development of unsteadiness over a large portion of the wingspan at high angles of attack. For tapered wings with backward-swept leading edges, unsteadiness emerges near the wing tip. On the other hand, wings with forward-swept trailing edges are shown to concentrate wake-shedding structures near the wing root. For highly swept untapered wings, the wake is steady, while unsteady shedding vortices appear near the tip for tapered wings with high leading-edge sweep angles. For such wings, larger wake oscillations emerge near the root as the taper ratio decreases. While the combination of taper and sweep increases flow unsteadiness, we find that tapered swept wings have more enhanced aerodynamic performance than untapered and unswept wings, exhibiting higher time-averaged lift and lift-to-drag ratio. The current findings shed light on the fundamental aspects of flow separation over tapered wings in the absence of turbulent flow effects.

Sensitivity Analysis of an Unsteady Flow Around a Pitching Airfoil

Various ways to control the loads and thus the output of the wind turbine by pitching the blades and by controlling the rotational speed already exist. However, adverse pressure gradient leading to flow separation at the trailing edge, with relatively high angles of attack (AOA) due to blade-pitching motion, largely affects the airfoil aerodynamic performance. This work presents the results of a numerical study of the unsteady flow around a pitching FFA-W3-301 airfoil at a Reynolds 1.6 × 106 using OpenFOAM®, as obtained by first performing 2D Unsteady Reynolds-Averaged Navier- Stokes (URANS) simulations whereby the flow characteristics are simulated by the shear stress transport (SST) k − ω model and Spalart-Allmaras (SA). The influence of various parameters on the numerical results is investigated, namely y+, computational grid resolution, dynamic mesh technique, time step and turbulence model. Integral aerodynamic forces and detailed flow patterns are compared with experimental measurements presented in the literature. A range of angles of attack, including early stalls, are examined. For best-performing parameters, an adequate refinement close to the wall was imperative in order to match experiments, especially during the upstroke motion of the airfoil, while for the downstroke phase, some differences still appeared. The agreement was greatly improved by using a 2.5D hybrid RANS-LES approach with enhanced delayed-Detached Eddy Simulation (DES) capabilities.

Numerical investigation for subsonic performance of the high-pressure capturing wing configuration with wing dihedral

High-pressure capturing wing (HCW) aerodynamic configuration demonstrates favorable aerodynamic performance under hypersonic conditions, and its novel additional lifting wing (also known as HCW) has the potential to enhance lift characteristics under subsonic conditions. Therefore, this configuration presents a promising option for wide-speed-range vehicles. However, the stability characteristics of this novel configuration under subsonic conditions have not yet been investigated. In this paper, the effects of wing dihedral angles on the subsonic aerodynamic characteristics of a parametric conceptual HCW configuration with two lifting wings were investigated. Specifically, the design variables for this study were the dihedral angles of the upper HCW and the lower delta wing. To obtain the distributions of various aerodynamic parameters over the design space, a combination of the uniform experimental design method, computational fluid dynamics numerical simulation techniques, and kriging surrogate model algorithm was employed. The findings suggest that wing dihedral angles have a greater impact on the lift-drag ratio (L/D) at low angles of attack compared to high angles of attack. L/D can be enhanced by incorporating a positive dihedral angle in HCW, and as the delta wing's negative dihedral angle rises, L/D tends to increase earlier and decrease later at low angles of attack. Furthermore, for the longitudinal, lateral, and directional stability characteristics of this configuration, the positive dihedral angles of the delta wing offer greater overall advantages than negative ones in improving them, and the positive dihedral angles of HCW yield more significant enhancements in stability compared to negative ones.

High-Speed Virtual Flight Testing Platform for Performance Evaluation of Pitch Maneuvers

To research serious nonlinear coupling problems among aerodynamics, flight mechanics, and flight control during high maneuvers, a virtual flight testing platform has been developed for a large-scale, high-speed wind tunnel, based on the real physical environment, and it can significantly mitigate risks and reduce the costs of subsequent flight tests. The platform of virtual flight testing is composed of three-degrees-of-freedom model support, measuring devices for aerodynamic and motion parameters, a virtual flight control system, and a test model. It provides the ability to realistically simulate real maneuvers, investigate the coupling characteristics of unsteady aerodynamics and nonlinear flight dynamics, evaluate flight performance, and verify the flight control law. The typical test results of a pitch maneuver with open-loop and closed-loop control are presented, including a one-degree-of-freedom pitch motion and a two-degrees-of-freedom pitch and roll motion. The serious pitch and roll-coupled motion during a pitch maneuver at a high angle of attack is revealed, and the flight control law for decoupled control is successfully verified. The comparison of the test results and the flight data of a real pitch maneuver proves the reliability and capability of virtual flight testing.

Leading-Edge Vortex Controller (LEVCON) Influence on the Aerodynamic Characteristics of a Modern Fighter Jet

The purpose of this paper is to assess the influence of a novel type of vortex creation device called the leading-edge vortex controller (LEVCON) on the aerodynamic characteristics of a fighter jet. LEVCON has become a trending term in modern military aircraft in recent years and is a continuation of an existing and widely used aerodynamic solution called the leading-edge root extension (LERX). LEVCON is designed to operate on the same principles as LERX, but its aim is to generate lift-augmenting vortices, i.e., vortex lift, at higher angles of attack than LERX. To demonstrate the methodology, a custom delta wing fighter aircraft is introduced, and details about its aerodynamic configuration are provided. The LEVCON geometry is designed and then incorporated into an existing three-dimensional (3D) model of the aircraft in question. The research is conducted using OpenFOAM 8, a high-fidelity computational fluid dynamics (CFD) open-source software. The computational cases are designed to simulate the aircraft’s flight at stall velocities within a high range of angles of attack. The results are assessed and discussed in terms of aerodynamic characteristics. A conclusion is drawn from the analysis regarding the perceived improvements in fighter jet aerodynamics. The analysis reveals that both lift and critical angle of attack can be manipulated positively. With the addition of LEVCON, the average lift gain in the high angle of attack (α) range is between 8.5% and 10%, while the peak gain reaches 19.4%. The critical angle of attack has also increased by 2°, and a flatter stall characteristic has been achieved.

Discovering optimal flapping wing kinematics using active deep learning

This paper focuses on the discovery of optimal flapping wing kinematics using a deep learning surrogate model for unsteady aerodynamics and multi-objective optimisation. First, a surrogate model of the unsteady forces experienced by a 3-D flapping wing is built, based on deep neural networks. The model is trained on a dataset of randomly generated kinematics simulated using direct numerical simulation (DNS). Once trained, the neural networks can quickly predict the unsteady lift and torques experienced by the wing, using sparse information on the kinematics. This fast surrogate model allows multi-objective optimisation to be performed. The resulting Pareto front consists of new kinematics that may be very different from the kinematics of the initial dataset. A few arbitrarily chosen kinematics on the Pareto front are thus simulated using DNS and used to enhance the database. The new dataset is used to train again the networks, and this active deep learning/optimisation framework is performed until convergence, obtained after only two iterations. Overall, this method reduced the cost of optimisation by 83 %. Results reveal two distinct families of motions. Kinematics promoting high efficiency are characterised by large stroke amplitudes and relatively low angles of attack, as observed for fruit flies, honeybees or hawkmoths. For those, lift production is driven by quasi-steady effects and the formation of a stable leading edge vortex. Kinematics promoting high lift are characterised by small stroke amplitudes and high angles of attack, reminiscent of mosquitoes. Lift production is driven by the rapid generation of vorticity at the trailing edge.

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.

Design and flight test of the fixed-flapping hybrid morphing wing aerial vehicle

Bionic aircrafts have become a hot topic in the research on unmanned aerial vehicles (UAVs). By mimicking the flight mechanics and wing morphology of living organisms, bionic aircraft not only improve aerodynamic performance, stealth, and convenience but also enable single aircraft to complete multiple specific tasks. To further improve the practical performance of bionic aircraft, a fixed-flapping hybrid morphing wing aerial vehicle (F-FWAV) is designed, and the aerodynamic and flight mechanics are studied. First, kinematic simulations are conducted on the designed F-FWAV, and the kinematic characteristics of the flapping wing are analyzed. Then, flight tests are conducted on the aircraft model, and time-domain and frequency-domain analyses were performed on the data of the pitch angle and roll angle during level flight, high angle of attack (AOA) maneuvering and landing, indicating that the pitch and roll angles of the F-FWAV are coupled. CFD numerical simulations are used to obtain the cruise points of the F-FWAV under different AOAs, frequencies, and freestream velocities. The results of the mutual interference between the fixed wing and flapping wing sections show that the flapping wing section provides more than 50 % of the lift of the entire flight process, and the downward flapping movement of the flapping wing section increases the edge vortex of the fixed wing section, thereby increasing the lift of the fixed wing section.

Experimental Validation of NACA 6321 Airfoil Characteristics Obtained Using Different Turbulence Models

Abstract: Numerical analysis of NACA 6321 aerofoil is conducted at different angles of attack with constant velocity using three turbulence models and the results are validated with experimental findings. The simulation study is conducted by solving the steady-state governing equation of continuity and momentum using the Spalart-Allmaras, k-omega, and k-epsilon models; the obtained results are compared with experimental data. Aerodynamic parameters are calculated and then juxtaposed with experimental data acquired from experiments performed in a subsonic tunnel. The study reveals that the results generated by the k-omega model exhibit a strong correlation with the experimental findings at low and high angles of attack when compared to other turbulence models. In contrast to the k-epsilon and the Spalart-Allmaras models, the prediction of the stalling angle of attack has an error of  20% in comparison to the experimental evaluation. The results of the k-omega turbulence model predict the turbulence properties pretty well in the NACA 6321 aerofoil with an error of less than 4%. Keywords: NACA 6321, Spalart-Allmaras, k-omega and k-epsilon, Turbulence models.

Experimental studies of airliner aerodynamic characteristics at overcritical angles of attack

PurposeThe main purpose of this study was to determine the basic aerodynamic characteristics of the airliner Tu-154M at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°.Design/methodology/approachWind tunnel tests of the Tu-154M aircraft model at the scale 1:20 were performed in a low-speed wind tunnel T-3 by using a six-component internal aerodynamic balance. Several model configurations were also investigated.FindingsThe results of the presented studies showed that at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°, the Tu-154M aircraft flap deflection affected the values of the drag and lift coefficients and generally had no major effect on the values of the side force and pitching moment coefficients.Research limitations/implicationsThe model vibration which was the result of flow separation at high angles of attack was the wind tunnel test limitation.Practical implicationsStudies of the airliner aerodynamic characteristics at the wide range of the overcritical angles of attack and sideslip angles allow assessment of the aircraft aerodynamic properties during possible unexpected situations when the passenger aircraft is found to have gone beyond the conventional flight envelope.Social implicationsThere are no social implications of this study to report.Originality/valueThe presented wind tunnel test results of the airliner aerodynamic characteristics at overcritical angles of attack and sideslip angles is an original contribution to the existing not-too-extensive database available in the literature.

Numerical investigation of virtual control surfaces for lift control on NACA0015 airfoil

The plasma actuators as virtual control surfaces are assessed numerically as a means to control the lift of NACA0015 airfoil at the full angle of attack (without stall). The virtual control surface for increasing lift is realized by the pressure side (PS) plasma actuators that induce an upstream jet and the suction side (SS) plasma actuators that induce a downstream jet (SSD plasma actuator), while the one for reducing lift is realized by the SS plasma actuators that induce an upstream jet (SSU plasma actuator). Numerical simulation is achieved by solving the two-dimensional Reynolds averaged Navier–Stokes using the finite volume method. The plasma actuator adopts the empirical model proposed by the author before. The simulation of the air flow was performed for the freestream velocity of 20 m/s (Re=1.03 × 106) and the induced jet momentum coefficient between 0.0846% and 0.9027%. The calculation results show that the optimal number of DBD actuators for increasing the lift is related to the angle of attack. The SS flow separation of the high angle of attack greatly reduces the control effect of the PS actuator, which can be eliminated by arranging the actuators in front of the separation point. Finally, a virtual control surfaces configuration containing three groups of seven plasma actuators is obtained.

Experimental and Numerical Flutter Analysis Using Local Piston Theory with Viscous Correction

Due to the maneuver and overload requirements of aircraft, it is inevitable that supersonic fins experience high angles of attack (AOAs) and viscous effects at high altitudes. The local piston theory with viscous correction (VLPT) is introduced and modified to account for the 3-dimensional effect. With the contribution of the explicit aerodynamic force expression and enhanced surface spline interpolation, a tightly coupled state-space equation of the aeroelastic system is derived, and a flutter analysis scheme of relatively small computational complexity and high precision is established with a mode tracking algorithm. A wind tunnel test conducted on a supersonic fin confirms the validity of our approach. Notably, the VLPT predicts a more accurate flutter boundary than the local piston theory (LPT), particularly regarding the decreasing trend in flutter speed as AOA increases. This is attributed to the VLPT’s ability to provide a richer and more detailed steady flow field. Specifically, as the AOA increases, the spanwise flow evolves into a gradually pronounced spanwise vortex, yielding an additional downwash and energizing the boundary layer, which is not captured by LPT. This indicates that the precision of LPT/VLPT significantly depends on the accuracy of steady flow results.

Virtual deflection with synthetic jet actuators at high angles of attack

Abstract A hybrid drone called a biplane quadrotor operates in both low (during the horizontal flight) and high (transition maneuver) Angle of Attack (AoA). So, this paper focuses on enhancing aerodynamic force during the transition maneuver. Synthetic Jet actuators (SJAs) can modify airfoil shapes virtually, so either the flow reattaches or flow separation will be delayed. This delay can enhance the aerodynamic force. In this paper, CFD analysis is performed using Ansys Fluent to study the impact of SJA on NACA 0015 airfoil at high (30°) AoA. This study aims to find the best location of SJA for high AoA to get maximum lift enhancement. The outcome of this study reveals that at 85 %, we can get maximum enhancement in the lift.

Application of fast-response pressure sensitive paint to low-speed vortical flow at high angles of attack

Application of fast-response pressure sensitive paint to low-speed vortical flow at high angles of attack

Aerodynamic Performance Assessment of an Airfoil at Reynolds Number 10000: A Computational Approach

Effects of a low Reynolds number flow on an airfoil with cavities of a specified shape are analyzed computationally to assess the aerodynamic performance by observing the coefficients of lift and drag with increasing angles of attacks. The Reynolds number considered for the simulation is 1×104 operating at atmospheric conditions at sea level. Indentations with a definite shape (right-angled triangle) are utilized to find the vortex effects of air in the cavity on the shear layer separation on airfoil surface to retain the flow even at higher angles of attack. The simulation is run in ANSYS Fluent to study the increase in the aerodynamic performance, research findings include an increase in coefficient of lift values by approximately, 111.86% of cavity with depth 0.05C indented at 70% of the chord length at 6° angle of attack, 68.13% of cavity with depth 0.05C indented at 50% of chord length at 6° angle of attack, 23.93% of cavity with depth 0.025C indented at 70% of chord length at 6° angle of attack, 3.96% of cavity of depth 0.025C at 6° angle of attack, 9.59% of cavity of depth 0.05C at 18° angle of attack.