Post-stall Angles Of Attack Research Articles

High lift devices using compliant surfaces

Stall delay and lift enhancement play a crucial role in modern aircraft performance. This is commonly achieved by devices such as slats or flaps located at the leading edge or trailing edge of an aircraft's wing. In this paper, we report a feasibility study of using light-weight compliant surfaces for novel high lift devices. The effects of compliant flags with one end fixed or both ends fixed near the leading edge and trailing edge of an airfoil were studied by force, flag deformation, and flow field measurements in a wind tunnel. When a flag is placed near the leading edge, the excitation of the separated shear layer from the leading edge is the main mechanism in increasing the lift at the post-stall angles of attack. In contrast, the trailing-edge flag with an excess length and both ends fixed could increase the effective camber and the circulation around the airfoil in a time-averaged sense. The mechanism is similar to that of the conventional Gurney flap effect, and equally effective at pre-stall and post-stall angles of attack. When used together, the compliant flags can delay stall angle by 8° and increase the maximum lift coefficient by 67% in the parameter range tested presently. Compliant surfaces require no external power as a passive method. If they are to be used as active methods, they are light weight, and can be stored and deployed easily.

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.

Slip influence on a blade performance under different pitch-oscillating motion

Slip velocity effects on the performance of an SD7037 wind turbine blade are processed. The slip boundary condition is associated with applying a superhydrophobic coating on the leading edge to avoid icing. The computational fluid dynamics (CFD) approach and the Transition-SST (Shear Stress Transport) viscous model are applied for a pitch-oscillating and a static airfoil with a range of slip lengths at Reynolds number of 4×104. The dynamic motion of the airfoil causes the dynamic stall (DS) phenomenon. This study focuses on investigating the effects of superhydrophobicity on deep DS and resultant aerodynamic loads under different reduced frequencies. For the static blade, the slip causes significant influences on the aerodynamic coefficients at a post-stall angle of attack: the lift coefficient is increased by 20%, and the drag is decreased by 30%. The airfoil shows weak loadings at low angles of attack. For the dynamic blade under DS, the reduced frequency varies from 0.05 to 0.20. The aerodynamic characteristics have the maximum dependency on the slip boundary conditions when the reduced frequency equals 0.12. Increasing the slip length postpones the DS vortices growth and DS occurrence angle of attack, regardless of reduced frequencies.

Post-stall flow control on aerofoils by leading-edge flags

Self-excited oscillations of flags attached at the leading edge of aerofoils have been investigated at post-stall angles of attack at a chord Reynolds number of 100 000. Significant increases in the time-averaged lift coefficient and stall angle have been observed for three aerofoils: one symmetric, one cambered and one with a sharp leading edge. The aerodynamic improvement is due to the periodic formation of vortices caused by the flag oscillations. When the flag is near the aerofoil surface, it is lifted upwards by the induced velocity of the growing vortex. As the flag moves up, the vortex grows in strength and reaches maximum circulation when the flag is furthest from the aerofoil surface and subsequently sheds. Flags with large stiffness exhibit better spatial and temporal coherence of flag oscillations than the compliant flags, resulting in a larger maximum lift coefficient and higher stall angle. For all aerofoils tested, the best lift enhancement with respect to the clean aerofoils is found when the angle of attack is 6° to 10° above the stall angle of the clean aerofoil. High lift is observed when the flags are locked in with the wake instability in a narrow frequency band, depending on the flag mass ratio and length.

Hysteresis effect on airfoil stall noise and flow field

The effect of hysteresis on the lift coefficient of an airfoil is well-known, i.e., the lift coefficient is higher during the upstroke than during the downstroke for the same angle of attack (AoA). In contrast, the effect of hysteresis on the noise emission has not been reported before. In order to investigate this effect, a NACA 0012 airfoil is investigated during upstroke and downstroke between zero and post-stall AoA at a constant Reynolds number of 5.1×105. A map of the noise emission vs AoA is presented revealing tonal noise emission at low AoA, and broadband noise beyond the linear range of lift coefficient. The trend of overall sound pressure level vs AoA shows two local peaks: one at 4.4° with high tonal noise and a second one at 12.6° where the maximum lift coefficient is observed. Flow measurements near the leading edge and trailing edge by particle image velocimetry show that the airfoil noise trend vs AoA is dominated by the trailing-edge flow features. The magnitude of noise emission during the upstroke is higher than that during the downstroke in the hysteresis loop, although the velocity fluctuation magnitude is smaller. This implies that the vortex coherence during upstroke is higher than downstroke, which is evidenced by a more coherent vortex during the upstroke.

Effect of wing planform on plunging wing aerodynamics

The unsteady aerodynamics of three plunging wings with different planform shapes was investigated in water tunnel experiments by means of force/moment and three-dimensional volumetric velocity measurements at a post-stall angle of attack. In contrast with a rectangular wing, a tapered wing does not exhibit a local minimum of the time-averaged lift coefficient at a reduced frequency of k≈ 1.7. This is due to the oblique shedding of trailing-edge vortices, which prevents the coupling with leading-edge vortices. The wing with a round-tip exhibits the same double maxima that correspond to the subharmonic and the fundamental frequency of vortex shedding in steady freestream. However, combined leading-edge and tip vortex produce larger time-averaged lift force with increasing reduced frequency. The variation of the bending moment is similar to that of the lift force in general. However, the mean bending moment arm shifts outboard with increasing reduced frequency due to the increased three-dimensionality of the vortex filaments and the tip-vortex effect. Maximum lift and bending moment become more similar for all three wings at higher reduced frequencies as the added mass effect becomes more significant.

Experimental Investigation of the Active Flow Control over a Cranked-Delta Wing

A series of experiments was carried out to investigate the effects of active flow control on the aerodynamic characteristics of a cranked-delta-wing model at a subsonic speed. A dielectric barrier discharge (DBD) plasma actuator was used to improve the aerodynamic performance of the half-span cranked-delta-wing model at various angles of attack. The model had inboard and outboard sweepback angles of 57 and 38 deg, respectively. The surface pressure distribution was obtained at five streamwise stations and at two ray stations. All data were obtained at a constant Reynolds number of 0.27×106 and at various angles of attack: 0 deg≤α≤36 deg. The results showed that the pulsed DBD plasma actuator improved the aerodynamic performance of the cranked-delta-wing model by preventing flow separation at high angles of attack. The deep stall angle of attack was delayed up to 6 deg, and the poststall suction coefficient was increased approximately 30% at an angle of attack of about α=34 deg when the pulsed DBD plasma was applied at its optimum condition (F+≈1 and duty cycle=10%). It was found that the pulsed plasma actuator suppressed the wake flow in the forward portion of the wing at poststall angles of attack, thus delaying the flow separation. However, the effects of the DBD actuator at very high angles of attack on the flow reattachment process at the rear portion of the wing for the present DBD setting were seen to be minimal. The induced flow by the pulsed DBD plasma actuator influenced the vortex merging phenomena and delayed the onset of the outer vortex burst at prestall angles of attack.

Experimental mitigation of large-amplitude transverse gusts via closed-loop pitch control

In this work, we experimentally demonstrate the utility of unsteady potential flow models in developing closed-loop control strategies for mitigating the lift transients on a wing experiencing large-amplitude transverse gusts. The developed closed-loop controller mitigates lift for gusts of various strengths and directions and for wings with pre- and post-stall angles of attack. Time-resolved force and flow field measurements are used to discover the salient flow physics during these encounters and illustrate how closed-loop actuation mitigates their lift transients.

Bio-inspired variable-stiffness flaps for hybrid flow control, tuned via reinforcement learning

A bio-inspired, passively deployable flap attached to an airfoil by a torsional spring of fixed stiffness can provide significant lift improvements at post-stall angles of attack. In this work, we describe a hybrid active–passive variant to this purely passive flow control paradigm, where the stiffness of the hinge is actively varied in time to yield passive fluid–structure interaction of greater aerodynamic benefit than the fixed-stiffness case. This hybrid active–passive flow control strategy could potentially be implemented using variable-stiffness actuators with less expense compared with actively prescribing the flap motion. The hinge stiffness is varied via a reinforcement-learning-trained closed-loop feedback controller. A physics-based penalty and a long–short-term training strategy for enabling fast training of the hybrid controller are introduced. The hybrid controller is shown to provide lift improvements as high as 136 % and 85 % with respect to the flapless airfoil and the best fixed-stiffness case, respectively. These lift improvements are achieved due to large-amplitude flap oscillations as the stiffness varies over four orders of magnitude, whose interplay with the flow is analysed in detail.

Effect of leading-edge protrusion shapes for passive flow control measure on wind turbine blades

In the present study, protrusions in the leading edge of the NACA 4415 airfoil were incorporated as a passive flow control measure on the wind turbine blades. Three airfoil models: unmodified leading-edge (ULE), spherical leading-edge protrusion (SLEP), and triangular leading-edge protrusion (TLEP), were investigated experimentally. Thereafter, CFD investigations were carried out using ANSYS 14.0 simulation tool to observe the flow characteristics around the airfoils. Further, LEP-based passive design was applied on a horizontal axis wind turbine (HAWT) to investigate its performance. An amplitude (A) of 1% of the chord length (C) was considered as the height of protrusion, while a distance of 0.25C was maintained between the protrusions to fabricate the experimental models. The study was performed at a low Reynolds number (Re) of 1.5 × 105 for a wide angle of attack (α) of 0°–20°. The experimental results demonstrate that the SLEP model has a higher lift coefficient at α ≥ 18°, whereas the TLEP model performs poorly when compared with the ULE model. The instantaneous lift coefficient plot indicates that the SLEP model generates a more stable force at a higher angle of attack. The computational analysis reveals that a larger primary circulation (extended till 0.2C) is observed in the ULE model at a post-stall angle of attack (α = 18°), indicating early flow separation. Whereas, in the SLEP and TLEP models, it is extended to 0.70C and 0.54C, respectively, indicating the best flow controlling measures achieved by using the SLEP model. The investigations of HAWT rotors with LEP revealed that SLEP HAWT exhibit 8.2% more power coefficient than ULE HAWT.

Mechanisms of Morphing Wall Flow Control by Traveling Waves over an Airfoil

The main two mechanisms of morphing wall flow control are direct injection of momentum in the streamwise direction and indirect momentum transfer via triggering instabilities. Traveling waves have been shown to perform better than standing waves, probably because they can use both mechanisms. However, the relative importance of the two mechanisms is not known. To differentiate between the mechanisms, a range of parameters (frequency, amplitude, and starting location) at stall (15 deg angle of attack) and poststall (20 deg angle of attack) is tested using wall-resolved large-eddy simulations with a sharp-interface curvilinear immersed boundary method at a low Reynolds number of over a NACA0018 airfoil. The results of the simulations demonstrate that the flow is reattached within a range of nondimensional frequencies, actuation amplitudes, and starting locations of oscillation at the stall and poststall angles of attack. Significant lift enhancement and drag reduction are also observed within these ranges. The nondimensional frequency range at which the flow is reattached is found to be similar to the dominant nondimensional frequencies of leading-edge vortex shedding of the unactuated airfoil. These indicate that the indirect transfer of momentum is the dominant mechanism because direct injection of momentum increases with the increase of amplitude and frequency; that is, separation should reduce as they increase. Nevertheless, direct injection of momentum improves the performance relative to pure excitations of standing waves when instabilities are triggered.

Experimental and computational investigations of aerodynamic characteristics of a finite rectangular wing-in-ground effect

Wind tunnel experiments are carried out on a 3D wing-in-ground effect at pre- and post-stall angles of attack at a relatively low Reynolds number, Re ≈ 8.8 × 104. The rectangular wing, aspect ratio, AR = 6.4 used in the present work has a symmetric airfoil section, NACA0012 and is studied at different ranges of ground proximity and also compared to a wing of cambered airfoil section, NACA4415 not in ground proximity. Unsteady and time-averaged six-component aerodynamic characteristics, coefficients of Lift, Drag, Side forces, Pitch, Roll, Yaw moments at several angles of attack in the range – 8° < α < 18° are reported. Aerodynamic efficiency and drag polars are studied in lieu of ground proximity at several angles of attack. Effect of ground proximity on the occurrence of stall and [Formula: see text] is studied. Numerical calculations using steady-state CFD are also made for quantitative comparison. The effect of ground proximity on span-wise pressure distribution and induced velocity in the vicinity of the wing is also discussed in detail.

Flow development over inclined flat plates in ground effect and relation to aerodynamic loads

The aerodynamics of finite-span inclined flat plates in ground effect is experimentally investigated at a chord-based Reynolds number of 50 000 for aspect ratios of 1 and 2. The minimum ground height is varied between 0.1 and 1.0 chord lengths, and lift and drag forces are measured using a force balance for angles of attack between −90° and 90°. Planar, two- and three-component particle image velocimetry is used to perform streamwise and cross-plane measurements at the midspan and one chord length downstream of the trailing edge, respectively. Ground effect is significant at ground clearances below 0.5 chord lengths, most notably near the stall angle, where it leads to significant changes to flow development. At sufficiently low free flight pre-stall angles, the increase in edge velocity at low gap ratios caused greater suction, generating higher lift with a minimal increase in drag for both orientations. Closer to the free flight stall angle, a decrease in aerodynamic loading is observed for negative orientations due to earlier onset of stall with a decreasing gap ratio. The exception was the higher aspect ratio plate at negative orientations, where the loading was largely invariant to changes in gap ratio for all angles tested. At positive orientations, the increase in average static pressure along the pressure surface in ground proximity led to an overall increase in loading prior to deep-stall conditions for both aspect ratios. The ground effect was minimal at post-stall angles of attack. The results may be used to guide the design of photovoltaic supports at relevant latitudes.

Hydrofoil geometry and leading-edge modifications: Influence of section profile, aspect ratio, and sweep

Modifications mimicking the leading-edge protuberances on the pectoral fin of humpback whales have been widely adopted to the designs of foils and control surfaces to delay stall and provide post-stall superior lifting characteristics. However, the efficacy of these modifications in flow control and lift augmentation depends on the wing geometry. The present work investigates the lift, drag, and flow characteristics of finite-span hydrofoils with different section profiles-both symmetric and cambered (NACA 0012, NACA 0018, NACA 634-021, NACA 2412, and NACA 4415) having twin-protuberances over their leading-edge at a Reynolds number of 2 × 105. The influence of aspect ratio (1, 2, and 3) and leading-edge sweep angle (0 deg, 15 deg, and 26.1 deg) on the hydrodynamic performance of modified foils is also investigated. The characteristic feature of twin-protuberance foil designs is the restriction of flow separation between the chordwise vortices shed from the two protuberances in certain post-stall regimes. The use of leading-edge modifications is observed to be more conducive for lift enhancement in thin foil sections, especially NACA 0012, having a smaller stall angle. Considering the other parameters, leading-edge protuberances are observed to be advantageous for hydrofoils with aspect ratios 2 and 3 at post-stall angles of attack, and for 0 deg and 15 deg sweep angles for post-stall lift enhancement. The influence of Reynolds number on the performance modifications is also investigated.

Compounding Transient Airfoil Motions and the Effectiveness of Linear Superposition

This paper investigates, for the first time, the effects of compounding transient airfoil motions and the predictive capability of the linear superposition principle in vortex dominated flows. Significant increases in the peak lift and nose-down pitching moment were observed during the second of two transient plunging motions at a poststall angle of attack. The load response of the second motion was estimated through linear superposition of the first motion response with a surprising level of accuracy. Flowfield measurements revealed this performance to coincide with a constructive merging of leading-edge vortices (LEVs). LEV merging showed sensitivity to motion timing. Breakdown of the linear superposition prediction coincided with LEV detachment and trailing-edge vortex formation, which disrupted constructive LEV merging. The amplitude of the second motion showed no discernible effect on LEV merging, and subsequently the accuracy of the linear superposition prediction. An extension to periodic motion was investigated, where linear superposition of a single sinusoidal cycle was compared with the true periodic response. This was found to capture the mean lift increase for low to moderate reduced frequencies. Lift amplitude, however, was captured with reasonable accuracy across the range of reduced frequencies and amplitudes tested.

Fluid-structure interaction of a bio-inspired passively deployable flap for lift enhancement

Covert-feathers-inspired passive flow control is beneficial for aerodynamic performance at post-stall angles of attack. However, most studies model covert feathers as a rigidly attached or a freely moving flap on a wing. The performance of a flap mounted via a torsional spring, emblematic to the finite stiffness of bird feathers, has remained unexplored. We simulate strongly coupled fluid-structure interactions of flow past a stationary airfoil with a passively deployable, torsionally mounted flap on the suction surface. We then perform a parametric study and an in-depth analysis of the dominant flow features that affect performance.

Large Eddy Simulation of optimal Synthetic Jet Actuation on a SD7003 airfoil in post-stall conditions

Aerodynamic performances may be optimised by the appropriate tuning of Active Flow Control (AFC) parameters. For the first time, we couple Genetic Algorithms (GA) with an unsteady Reynolds-Averaged Navier-Stokes (RANS) solver using the Spalart-Allmaras (SA) turbulence model to maximise lift and aerodynamic efficiency of an airfoil in stall conditions [1], and then validate the resulting set of optimal Synthetic Jet Actuator (SJA) parameters against well-resolved three-dimensional Large Eddy Simulation (LES). The airfoil considered is the SD7003, at the Reynolds number Re=6×104 and the post-stall angle of attack α=14∘. We find that, although SA-RANS is not quite as accurate as LES, it can still predict macroscopic aggregates such as lift and drag coefficients, provided the free-stream turbulence is prescribed to reasonable values. The sensitivity to free-stream turbulence is found to be particularly critical for SJA cases. Baseline LES simulation agrees well with literature results, while RANS-SA would seem to remain a valid model to a certain degree. For optimally actuated cases, our LES simulation predicts far better performances than obtained by suboptimal SJA LES computations as reported by other authors [2] for the same airfoil, Re and α, which illustrates the applicability and effectiveness of the SJA optimisation technique applied, despite using the less accurate yet computationally faster SA-RANS. The flow topology and wake dynamics of baseline and SJA cases are thoroughly compared to elucidate the mechanism whereby aerodynamic performances are enhanced.

Aerodynamics of a wing in turbulent bluff body wakes

The aerodynamics of a stationary wing in a turbulent wake are investigated. Force and velocity measurements are used to describe the unsteady flow. Various wakes are studied with different dominant frequencies and length scales. In contrast to the pre-stall angles of attack, the time-averaged lift increases substantially at post-stall angles of attack as the wing interacts with the von Kármán vortex street and experiences temporal variations of the effective angle of attack. At an optimal offset distance from the wake centreline, the time-averaged lift becomes maximum despite of small amplitude oscillations in the effective angle of attack. The stall angle of attack can reach 20° and the maximum lift coefficient can reach 64 % higher than that in the freestream. Whereas large velocity fluctuations at the wake centreline cause excursions into the fully attached and separated flows during the cycle, small-amplitude oscillations at the optimal location result in periodic shedding of leading edge vortices. These vortices may produce large separation bubbles with reattachment near the trailing-edge. Vorticity roll-up, strength and size of the vortices increase with increasing wavelength and period of the von Kármán vortex street, which also coincides with an increase in the spanwise length scale of the incident wake, and all contribute to the remarkable increase in lift.

Numerical Study of Multiple Bio-Inspired Torsionally Hinged Flaps for Passive Flow Control

Covert feathers are a set of self-actuating, passively deployable feathers located on the upper surfaces of wings that augment lift at post-stall angles of attack. Due to these benefits, the study of covert-inspired passive flow control devices is becoming an increasingly active area of research. In this work, we numerically investigate the aerodynamic benefits of torsionally mounting five covert-inspired flaps on the upper surface of a NACA0012 airfoil. Two-dimensional high-fidelity simulations of the flow past the airfoil–flap system at low Re=1000 and a high angle of attack of 20∘ were performed. A parametric study was conducted by varying the flap moment of inertia and torsional hinge stiffness to characterize the aerodynamic performance of this system. Lift improvements as high as 25% were attained. Two regimes of flap dynamics were identified that provided considerable aerodynamic benefits. A detailed investigation of the flow physics of both these regimes was conducted to understand the physical mechanisms by which the passively deployed flaps augmented the lift of the airfoil. In both regimes, the flap was found to act as a barrier in preventing the upstream propagation of reverse flow due to flow separation and trailing edge vortex. The torsional spring and flap inertia yielded additional flap dynamics that further modulated the surrounding flow and associated performance metrics. We discuss some of these fluid–structure interaction effects in this article.

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.