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The effects of the aspect ratio, the sweep angle, and the leading-edge geometry of wings placed in a turbulent wake were investigated in wind-tunnel experiments at a chord Reynolds number of [Formula: see text]. The poststall lift enhancement due to the leading-edge vortex formation was studied at optimal locations in the wake. The effects of the strength of the leading-edge vortices, the ratio of the spanwise length scale of the incident vortex to the wingspan, and the degree of two-dimensionality of the wake–wing interaction were studied. The competition between the effects of the spanwise length scale of the incident wake and the strength of the leading-edge vortices determined the optimal aspect ratio, which was found to be around four. Increasing the sweep angle decreased the mean lift due to the decreased two-dimensionality of the vortex formation. Airfoils with sharp leading edge produced the strongest leading-edge vortices but further away from the wing surface, resulting in lower maximum lift. Relative to the performance in the undisturbed freestream, the increases in the stall angle and maximum lift coefficient were not significantly affected by the leading-edge shape.

The post-stall flow control using compliant flags of varying thickness and length, placed upstream a NACA0012 airfoil, was shown to be possible at an airfoil chord Reynolds number of 100,000. The flag wakes produced substantial increase in the stall angle and the maximum lift coefficient of the airfoil placed at optimal cross-stream locations from the wake centerline. Oscillating flags could generate periodic wakes with better spanwise coherence than the stationary bluff body. This resulted in the excitation, formation and shedding of the leading-edge vortices periodically, providing mean lift enhancement. There is an optimal range of the flag mass ratio for which the flag frequency coincides with the natural frequency of the vortex shedding instability or its subharmonic of the baseline airfoil wake. The flag dimensionless frequency is a function of the mass ratio only, which can be predicted by a reduced order model in the limit of very large mass ratio and by using the modified free-streamline theory for the separated flow. There is also an optimal range of the flag dimensionless frequency. Graphical abstract

A nearly two-dimensional vortex of small core size has been produced by the transient plunging motion of an upstream airfoil and it interacted with the downstream wings. Depending on the offset distance of the vortex and wing angle of attack, the incident vortex filament deforms, diffuses, and loses coherence, while inducing leading-edge vortex formation and shedding from the wing. No significant spanwise flow develops in the incident vortices during the interaction. The interaction with the swept wing at each spanwise plane appears to be unaffected by the other spanwise planes. The counter-clockwise vortex induces a positive lift peak as it approaches the wing, which can be predicted by the potential flow assumption. The peak lift force is proportional to the circulation of the incident vortex and has its maximum near the zero-offset distance. The minimum lift coefficient is reached after the vortex has just passed and caused flow separation on the lower surface. The maximum lift coefficients for the finite unswept and swept wings can be estimated by making a correction for the aspect ratio and using the independence principle. The only exception is observed for the swept wing at a post-stall angle of attack for which the leading-edge vortex shedding becomes parallel to the leading-edge and increases the peak lift force. Graphical abstract

A stationary wing placed in the wake of a bluff body experiences lift enhancement. The quasi-periodic flow in the wake causes excitation of the separated flow in the post-stall conditions. The increase in the time-averaged lift force is associated with the flow separation, leading-edge vortex formation and subsequent reattachment in a process similar to the dynamic stall of oscillating wings. The lift enhancement is maximum for an optimal offset distance from the wake centerline. At the optimal location, potential flow oscillations, rather than the direct impingement of large vortices in the wake, provide the excitation. The smaller amplitude flow oscillations lead to a large separation bubble in the time-averaged sense in the post-stall regime. The delayed flow separation in the wake has a similar mechanism and frequency to those of the active flow control methods for separation. The degree of the lift enhancement is remarkable, given that the wake at a Reynolds number of 50,000 is expected to be highly three-dimensional.

A novel approach to reduce the peak lift and pitching moment on a plunging airfoil is investigated through force, moment, and velocity measurements. This approach, unlike previous investigations of delayed flow separation and leading-edge vortex suppression, uses forced separation through deployment of a minitab near the leading edge. The device can be activated for short time intervals during a gust encounter or unsteady maneuver at the expense of short-duration drag increase. Depending on the frequency and the amplitude of the wing motion and the mean angle of attack, roll-up of vorticity and the formation of a vortex can be delayed or even prevented. This change in the vortex dynamics provides effective lift and moment alleviation for post-stall angles of attack and for low reduced frequencies. In contrast, at low angles of attack, the separated shear layer may roll up for the manipulated flow, resulting in vortex shedding, and lift and nosedown pitching moment increase. These two distinct flow regimes cause decreased or increased lift force, with the most effective frequencies scaling with the reduced frequency. In contrast, the borderline between the two regions scales with the Strouhal number based on amplitude and, in particular, with the minimum effective angle of attack during the cycle. The transient response was studied by investigating impulsively started plunging oscillations. During the first cycle, lift reduction is achieved for all frequencies within the range tested.

We present an experimental investigation of the flow structure in the near wake of a NACA0012 airfoil plunging sinusoidally at a chord Reynolds number of Re = 20 000 and for a wide range of reduced frequency k and Strouhal number based on peak-to-peak amplitude St. Estimated mean thrust coefficients using the mean and fluctuating velocity fields confirm the St2 dependence as well as a significant effect of the reduced frequency for k ≤ 1. Generally, time-averaged flow quantities are better correlated with St than k in the range tested (k ≤ 3.14 and St ≤ 0.24). Analysis of the streamwise flow and cross-flow in the near wake using two-point cross-correlations and proper orthogonal decomposition reveals that the unsteady characteristics are even better correlated with St than the mean flow quantities. The percentage energy of the fundamental wake modes of the streamwise flow and the flapping mode of the cross-flow increases with increasing St, but at different rates in the drag-producing and thrust-producing wakes. There are similarities to the wake synchronisation behind oscillating bodies. The spanwise-averaged cross-correlation coefficient in the measurement domain grows linearly for small St (in drag-producing wakes), and is nearly constant at a high value for larger St (in thrust-producing wakes). Results show that the Strouhal number is the most important parameter that determines the degree of two-dimensionality of the wake, and suggest that spanwise vortices are quasi-two-dimensional for St ≥ 0.05 and x/c ≤ 4. The implications for experimental gust generators using oscillating airfoils are discussed.

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.

The vortex dynamics of leading-edge vortices on plunging high-aspect-ratio (AR = 10) wings and airfoils were investigated by means of volumetric velocity measurements, numerical simulations and stability analysis to understand the deformation of the leading-edge vortex filament and spanwise instabilities. The vortex filaments on both the wing and airfoil exhibit spanwise waves, but with different origins. The presence of a wing-tip causes the leg of the vortex to remain attached to the wing upper surface, while the initial deformation of the filament near the wing tip resembles a helical vortex. The essential features can be modelled as the deformation of an initially L-shaped semi-infinite vortex column. In contrast, the instability of the vortices is well captured by the instability of counter-rotating vortex pairs, which are formed either by the trailing-edge vortices or the secondary vortices rolled-up from the wing surface. The wavelengths observed in the experiments and simulations are in agreement with the stability analysis of counter-rotating vortex pairs of unequal strength.

Unsteady aerodynamics of two flexible cantilevered wings in transverse gusts has been studied by means of force, velocity, and deformation measurements. The first bending frequencies of the wings were in the range of gust frequencies tested. The deformation measurements revealed that plunge angle due to the bending deflections is generally small. However, it becomes large at the resonant conditions and significantly attenuates the amplitude of the effective gust angle perceived by the wing, although the unsteady lift is difficult to measure accurately for the resonant cases. As the gust frequency increases, the more flexible wing experiences more separated flow during the cycle. Nevertheless, for both flexible wings, the lift amplitude normalized by the gust amplitude is smaller than those of the rigid wings with separated flows. When the maximum effective angle of attack is smaller than a critical value, the lift amplitude normalized by the gust amplitude is close to the theoretical predictions for attached flows and the effect of the gust reduced frequency is small. In contrast, above the critical value of the maximum effective angle of attack, the normalized lift can be amplified or attenuated, depending on the gust reduced frequency.

View Video Presentation: https://doi.org/10.2514/6.2022-1362.vidExperiments were carried out in a wind tunnel to study the effects of a small flag attached to the surface of an airfoil. It was found that the flag exhibited quasi-periodic oscillations between the airfoil surface and the freestream at post-stall angles of attack of the clean airfoil when it was attached near the leading-edge. Digital image correlation (DIC) measurements of deformation field and particle image velocimetry (PIV) measurements of velocity field indicated that the limit cycle oscillations of compliant latex flags produced leading-edge vortices when the flag tip reached its maximum displacement, thus forming a separation bubble, and promoting flow reattachment to the airfoil surface. As a result, the stall angle could be delayed significantly, and the lift coefficient could be increased remarkably, e.g., up to 73% increase in the mean lift coefficient of the clean airfoil in the post-stall regime was observed in the present investigation. Proper orthogonal decomposition analyses of the flag deformation field showed that the dominant mode of flag oscillations occurred mostly in the first beam mode, whereas the higher modes in spanwise direction revealed the three-dimensionality of the flag deformation field. Flags reinforced with plastic shims, which have larger bending rigidity, however, can improve the two-dimensionality of the flag displacement fields, thus improving the lift enhancement and stall-delay even further. The oscillation frequencies of the flags suggest coupling with wake instabilities and wake resonance at the post-stall angles of attack. This is a passive flow control method that relies on unsteady effects and requires no external power.

We present experiments demonstrating that a small flag attached to the surface of an airfoil near the leading edge exhibits self-excited oscillations when the airfoil is set at a post-stall angle of attack. The limit cycle oscillations of the flag between the airfoil surface and the freestream have a dominant frequency and promote reattachment of the flow, producing a remarkable increase in the maximum lift coefficient and stall angle. Leading-edge vortex shedding and formation of a separation bubble are observed when the flag tip reaches around the maximum displacement. The instantaneous flow and the range of oscillation frequency have strong similarities to those of the active flow control methods for separation, yet this is a passive flow control method that relies on unsteady effects. The flag oscillations occur mostly in the first structural mode, while the higher modes reveal the three-dimensionality of the flag displacement field. The oscillation frequencies of the flag are in the same range of vortex shedding frequency of the baseline airfoil, suggesting that the wake resonance mechanism may be behind the observations of the enhanced lift at the post-stall angles of attack. There is virtually no effect of the compliant flag on the mean lift at small angles of attack.

Load control is the reduction of extreme aerodynamic forces produced by gusts, maneuvers, and turbulence to enable lighter, more efficient aircraft. To design an effective control system, the actuator's response in terms of amplitude and phase lagmust be known. Current load control technologies are limited to low-frequency disturbances due to their large inertia. This paper evaluates a potential high-frequency alternative: The minitab using periodic and transient deployments on a NACA0012 airfoil in wind-tunnel experiments. Periodic deployment for reduced frequencies, k ≤ 0.79 exhibits a normalized lift response amplitude, which decays with increasing k comparable to Theodorsen's circulation function but with substantially higher lag. Transient deployment, at rates as low as τdeploy = U∞;tdeploy/c = 1, illustrates a delay in aerodynamic response. The delay is larger for outward minitab motion than inward; τ; ≈ 6 and4, respectively, forα = 0 degandincreases withα. Theflowfields showthat the delay in response and the reduction in effectiveness for dynamic minitab deployment are due to delayed growth of the separated region behind the minitab. The aerodynamic response due to minitab deployment is approximated as the response of a first-order system, which is pertinent to control system design. This simple characterization for amplitude reduction and delay in response makes it well suited to load control.

Anexperimental study was performed in a water tunnel to evaluate the effects of continuous and pulsed blowing jets on the counter-rotating vortices generated by the afterbody of a slanted base cylinder. Drag reductions from continuous blowing through circular jets were found to vary significantly with direction and location, and approached 7% when blowing outboard from upstream locations on the upswept face. However, for all circular jets tested, the external power required was larger than the power saved due to the drag reduction. Jet vortices restricted shear layer development, leading to smaller afterbody vortex cores further from the surface. A high-aspect-ratio jet flap, ejecting nearly parallel to the freestream, achieved drag reductions close to 9%, equating to the net energy savings of almost 3% for the best case. Jet vortices shortened the shear layer, resulting in vortices with reduced circulation, which were displaced away from the upswept face. Pulsing the jet flap resulted in improved drag reductions and energy savings (up to around 6%) compared with the equivalent continuous blowing case at the same time-averaged jet momentum coefficient. Pulsed blowing caused an increase in vortex separation and meandering, whereas the circulation was reduced by up to 10% of that for continuous blowing.

The vortex dynamics of thrust and lift generation for periodically plunging airfoils and finite wings at low Reynolds numbers is reviewed. Effects of vortex lock-in phenomenon, leading-edge vortex,...

Abstract A novel small-amplitude high-frequency gust generator has been developed that works by oscillating a small fence on the wind tunnel wall. The gust generator produces approximately constant local angle of attack in the chordwise direction. Due to the challenges of measuring small and slightly non-uniform gust angles the gust generator was calibrated using direct lift measurements on a symmetric wing set at zero geometric angle of attack. Unsteady lift force measurements and the Theodorsen's theory were used for the dynamic calibration of the gust angle. At nonzero geometric angles of attack, if the wing’s effective angle of attack remains below the stall angle, unsteady lift closely follows the static lift curve with very small hysteresis. Beyond the stall angle, dynamic stall and larger lift hysteresis are observed. Interestingly, in this regime, if flow is separated and a separation bubble is maintained on the wing throughout the cycle then increasing frequency reduces lift hysteresis. The slope of the lift curve, averaged over the cycle, may be greater than that of attached flow. The gust response is more sensitive to maximum effective angle of attack than the reduced frequency or the reduced pitch rate. The normalized lift change is much larger for separated flows than for attached flows. Graphic abstract

To investigate the effect of sweep on the leading-edge vortex of high aspect ratio plunging wings, this article presents force, bending moment, and volumetric velocity measurements for high aspect ratio AR = 10 wings with sweep angles of 0° and 40°. The effect of the sweep angle on the bending moment is the largest at the minimum effective angle of attack. This is because as the leading-edge vortex sheds it moves inboard on the unswept wing while moving outboard on the swept wing. Where the leg of the leading-edge vortex connects with the wing there is significant three-dimensional flow. The axial velocity along the vortex filament, which may be towards to the wing tip or the wing root (reversed flow), exhibits increasing magnitude as the effective angle of attack decreases and the vortex filament deforms. Reversed axial flow along the vortex filament has the largest magnitudes for the unswept wing. In the vortex core, jetlike, wakelike, and uniform axial velocity profiles were observed. Unlike the classical vortex breakdown, the transition from the jetlike to the wakelike axial flow does not appear to be abrupt. The measurements also revealed evidence of spanwise instabilities in the leading-edge vortex filament.

Increased aerodynamic loads during gusts, turbulence, and maneuvers define the outer envelope of aircraft structural design. Minitabs, which are small spanwise strips that protrude normal to the ai...

Counter flowing wall jets are proposed to reduce extreme loads, enabling lighter and more efficient aircraft. Upstream blowing on the upper surface of a NACA0012 airfoil at a Reynolds number of 660,000 was investigated by means of force, pressure and particle image velocimetry measurements, and was found to be more effective than blowing normal to the surface. Blowing locations near the trailing edge are more effective for low angles of attack; locations near the leading edge are more effective for higher angles of attack. At the stall angle of attack, a maximum reduction in the lift coefficient was observed as Δ C L = − 0.33 when the blowing location was at x J / c = 0.08 for the largest flow rate coefficient ( C Q = 0.44 % ) tested. Even for completely separated flows, upstream blowing near the trailing-edge can modify the recirculation region and cause the shear layer to deflect upwards, resulting in a reduction of lift. This is a significant advantage over mechanical flaps placed near the trailing-edge, which remain in the separated flow and lose their effectiveness.

Location, strength, and structure of tip and edge vortices shed from wings and bodies can be manipulated by using flow control techniques. Flow physics of these approaches involve flow separation from the edge, roll-up into the vortex, wing flow regime, vortex instabilities, vortex–vortex interactions, and vortex–turbulence interactions. Actuators include continuous and unsteady blowing as well as suction, bleed, and control surfaces, which add momentum, vorticity, and turbulence into the vortices. It is noted that actuation may have effects on more than one aspect of the flow phenomena. A comparative review of the control of delta-wing vortices, tip vortices, and afterbody vortices is presented. The delay of vortex breakdown and the promotion of flow reattachment require different considerations for slender and nonslender delta wings and may not be possible at all. Tip vortices can be controlled to increase the effective span, to generate rolling moment, to attenuate wing rock, and to attenuate vortex–wing interactions. Although there are different approaches for each application, opportunities for future research on turbulence ingestion, bleed, and excitation of vortex instabilities exist. Recent research also indicates that active and passive flow control can be used to manipulate the afterbody vortices to reduce the drag.

An investigation of symmetry breaking and naturally occurring instabilities over thin slender delta wings with sharp leading edges was carried out in a water tunnel using particle image velocimetry (PIV) measurements. Time-averaged location, strength and core radius of conical vortices vary almost linearly with chordwise distance for three delta wings with $75^{\circ }$, $80^{\circ }$ and $85^{\circ }$ sweep angles over a wide range of angles of attack. Properties of the time-averaged vortex pairs depend only on the similarity parameter, which is a function of the angle of attack and the sweep angle. It is shown that time-averaged vortex pairs develop asymmetry gradually with increasing values of the similarity parameter. Vortex asymmetry can develop in the absence of vortex breakdown on the wing. Instantaneous PIV snapshots were analysed using proper orthogonal decomposition and dynamic mode decomposition, revealing the shear layer and vortex instabilities. The shear layer mode is the most periodic and more dominant for lower values of the similarity parameter. The Strouhal number based on the free stream velocity component in the cross-flow plane is a function of only the similarity parameter. The dominant frequency of the shear layer mode decreases with the increasing similarity parameter. The vortex modes reveal the fluctuations of the vorticity magnitude and helical displacement of the cores, but with little periodicity. There is little correlation between the fluctuations in the cores of the vortices.

The effect of suction on an airfoil surface at various locations downstream of the leading edge of a thin flat-plate airfoil was studied in a wind tunnel at a low Reynolds number. At poststall angles of attack, substantial lift enhancement and delay of stall can be achieved if a large separation bubble is generated by reattaching the massively separated flow near the trailing edge. The effects of location and volumetric flow rate of suction were investigated by means of force and velocity field measurements. There is an optimal location of suction around xs/c=0.4, which generates the maximum lift coefficient for suction coefficients less than 3%. When suction is applied closer to the leading edge, it may be possible to reattach the flow for smaller suction coefficients, but the resulting small separation bubble causes smaller lift increase. Large separation bubbles are needed for the maximum time-averaged lift enhancement, however, they exhibit shear layer flapping, intermittent reattachment, and larger l...

Counter-flowing wall jets actuated on the upper surface of an airfoil are investigated for the purpose of controlling gust encounters. For periodic and transient actuations, phase-averaged lift force and particle image velocimetry measurements are presented for a NACA 0012 airfoil, at a Reynolds number of 660,000, for a range of reduced frequencies and three jet locations, x J / c = 0.08 , 0.60 and 0.95. For periodic actuation, amplitude of lift oscillations decrease and phase delay increase with increasing reduced frequency. The effect of reduced frequency on the amplitude and phase is more significant for blowing locations near the leading-edge and with increasing angle of attack. Transient actuation reveals the slow response of the separated flow, and therefore lift, with the delay becoming more pronounced for blowing near the leading-edge. Estimated time constants are similar to previous observations for forced separation and reattachment.

The peak loads experienced by aircraft of all scales will typically be during gusts, turbulence or extreme manoeuvres. Understanding the aerodynamic response to these transient disturbances is therefore crucial, particularly when Leading-Edge Vortices (LEVs) occur. This fundamental study investigates the aerodynamic response to a wide range of transient plunging motions. The peak loads exhibited a strong dependence to motion amplitude yet remained relatively insensitive to motion duration. Within the parameter range tested (motion duration of T i 20 τ , or equivalent reduced frequency k i 1, and plunge amplitude of α p l , p e a k ≤ 30°), the peak lift did not exceed that of the quasi-static thin airfoil theory prediction, permitting its use as a safe limit for structural design. The normalized peak lift change displayed weak collapse with the timescale of the motion and instead showed better correlation with the non-dimensional plunge rate. The peak pitching moment scales well with plunge rate according to the theoretical prediction due to the added-mass component for plunge-up motions, but quickly diverges for plunge-down motions. At post-stall angles of attack, large-scale vortex shedding was observed and caused decaying oscillations in the loads long after the transient motion ends. For both a NACA 0012 and flat plate airfoil, the first vortex shedding cycle after the transient motion occurs around the subharmonic of the static shedding frequency. Subsequent shedding cycles then increase in frequency and asymptotically approach the static shedding frequency in around 15 to 20 convective times. This is the first study to experimentally quantify this behavior and is an aspect currently missing in existing reduced-order models, which could be significant for the prediction of successive transient disturbances. Finally, Reynolds number insensitivity was demonstrated for transient disturbances between 20,000 and 150,000, even for post-stall angles of attack where large-scale vortex shedding can occur.

Near-wakes of axisymmetric cylinders with slanted base were investigated in wind tunnel experiments for an upsweep angle of 28°. Effects of splitter-plate, cavity, and flaps on the afterbody vortices and separated flow were studied by means of surface pressure and Particle Image Velocimetry measurements. The splitter plate causes more diffused afterbody vortices due to the turbulence ingestion from the separation region. When the slanted base is replaced with a deep cavity, there is weaker roll-up of vorticity due to the lack of streamwise flow and a solid surface. Varying the splitter plate length in the range tested did not have significant influence on the flowfield, apart from affecting the strength of the splitter-plate vortex. With the splitter plate and cavity, unsteadiness is dominated by the flow separation region, as opposed to the afterbody vortices in the baseline case. A pair of vertical flaps attached to the side-edges of the splitter plate can reduce the unsteadiness at the measurement plane immediately downstream of the splitter plate.

Drag force, surface pressure and particle image velocimetry measurements havebeen carried out in wind tunnel experiments in order to study the effect of minispoilers on the drag of an axisymmetric cylinder with a slanted base that has an upsweep angle of Φ = 28°. Two different spoiler heights (h/D = 2.5% and 5%) were examined at various chordwise locations along the slanted surface. Placing the spoiler closer to the beginning of the slanted section caused an increased drag due to the separation induced by the spoiler itself, although a more diffused afterbody vortex was produced due to the turbulence ingestion. Drag reduction was observed when placing the spoiler closer to the trailing-edge, with the optimum location x′s /c = 87.5%, resulting in approximately 5% drag reduction for both spoiler heights. This reduction is due to the displacement of the afterbody vortex away from the surface as the mini-spoiler acts similarly to a Gurney flap. The same effect can also be produced by introducing a smaller co-rotating vortex to manipulate the afterbody vortextrajectory.

The interaction of meandering tip vortices shed from a leading wing with a downstream wing was investigated experimentally in a water tunnel using flow visualization, particle image velocimetry measurements, and volumetric velocity measurements. Counter-rotating upstream vortices may exhibit sudden variations of the vortex core location when the wing-tip separation is within approximately twice the vortex core radius. This is caused by the formation of vortex dipoles near the wing tip. In contrast, co-rotating upstream vortices do not exhibit such sensitivity. Large spanwise displacement of the trajectory due to the image vortex is possible when the incident vortex is further inboard. For both co-rotating and counter-rotating vortices, as long as there is no direct impingement upon the wing, there is a little change in the structure of the time-averaged vortex past the wing, even though the tip vortex shed from the downstream wing may be substantially weakened or strengthened. In the absence of the downstream wing, as well as for weak interactions, the most energetic unsteady modes represent the first helical mode |m| = 1, which is estimated from the three-dimensional Proper Orthogonal Decomposition modes and has a very large wavelength, on the order of 102 times the vortex core radius, λ/a = O(102). Instantaneous vorticity measurements as well as flow visualization suggest the existence of a smaller wavelength, λ/a = 5–6, which is not among the most energetic modes. These two-orders of magnitude different wavelengths are in agreement with the previous measurements of tip vortices and also exhibit qualitative agreement with the transient energy growth analysis. The very long wavelength mode in the upstream vortex may persist during the interaction, and reveal coupling with the trailing vortex as well as increased meandering.

The interaction of trailing vortices of an upstream wing with rigid and flexible downstream wings has been investigated experimentally in a wind tunnel, using particle image velocimetry, hot-wire, force and deformation measurements. Counter-rotating upstream vortices exhibit increased meandering when they are close to the tip of the downstream wing. The upstream vortex forms a pair with the vortex shed from the downstream wing, and then exhibits large displacements around the wing-tip. This coupled motion of the pair has been found to cause large lift fluctuations on the downstream wing. The meandering of the vortex pair occurs at the natural meandering frequency of the isolated vortex, with a low Strouhal number, and is not affected by the frequency of the large-amplitude wing oscillations if the downstream wing is flexible. The displacement of the leading vortex is larger than that of the trailing vortex, however it causes highly correlated variations of the core radius, core vorticity and circulation of the trailing vortex with the coupled meandering motion. In contrast, co-rotating vortices do not exhibit any increased meandering.

In this study lift, pitching moment and particle image velocimetry (PIV) measurements are presented for a NACA 0012 airfoil with a Reynolds number of 20,000 at a geometric angle of attack in the range 0° to 20° undergoing a transient plunging motion. The transient motion is defined by a Strouhal number based on the chord length up to 0.60 and effective plunge angle amplitude from -30° to 30°. The largest peak in both lift and pitching moment was observed during the motion for all cases. This was attributed primarily to circulatory effects for the lift and added-mass effects for pitching moment. PIV showed the presence of a coherent leading-edge vortex (LEV) at positive plunge angle amplitudes which had the biggest effect on the pitching moment during motion. Substantial peaks in lift and pitching moment were observed for post-stall geometric angles of attack after the end of motion. This is postulated to be due to strong interactions between the leading and trailing-edge vortices. An increase in plunge angle amplitude caused an increase in the magnitudes of lift/moment change. This shedding phenomena was found for both positive and negative plunge angles and showed remarkably similar peak relative timings, highlighting the insensitivity of the shedding process to the initial transient forcing.

The interaction of trailing vortices from a leading wing with rigid and flexible downstream wings has been investigated experimentally in an open jet wind tunnel using force measurements, particle image velocimetry measurements, hotwire measurements and deformation measurements. Counter-rotating leading vortex has been found to have substantial meandering when it is close to the tip of the downstream wing. The leading vortex forms a pair with the vortex from the downstream wing, and then exhibits large displacements around the wing tip. This has been found to cause the largest lift fluctuations on the downstream wing. The meandering of the vortex pair occurs at the natural meandering frequency of the isolated vortex, with a low Strouhal number, and is not affected by the frequency of the wing bending oscillations if the downstream wing is flexible. While the displacement of the leading vortex is larger than that of the trailing vortex, time-dependent variations of the core radius, core vorticity and circulation are evident for the trailing vortex. In contrast, co-rotating vortices do not exhibit any increased meandering.

The work presented within this study is sponsored by the United States Airforce Office for Scientific Research (AFOSR) under grant number FA9550-14-1-0126.

Active and passive control of leading-edge vortices over a 70°/50° double delta wing by means of jet blowing and bleed were studied experimentally in a closed-loop water tunnel using particle image velocimetry (PIV) measurements. Depending on the jet location and yaw angle, the strake vortices and the wing vortices could be separated further away from each other thus resulting in weak interactions or the interactions between the strake and wing vortices could be intensified thus leading to an earlier merge. Ingestion of jet turbulence into the vortices appears to promote merging. Evidence suggests that significant forces/moments could be produced even at low momentum coefficients. Passive bleed with estimated momentum coefficients of the order of 0.1% can also promote vortex merging. Secondary vortex can be strengthened and alter the vortical flow structure over the delta wing. It was found that bleed is more effective at lower wing incidences.

Experiments have been undertaken to study the formation of afterbody vortex flows from cylindrical bodies with a slanted base, whose upsweep angle was varied between 24° and 32°. Vortex roll-up is mostly completed in the first half of the upswept section, where the vortex causes largest suction on the surface. Towards the trailing-edge the vortices become more axisymmetric and stronger with increasing upsweep angle. Although there is some delay in vortex roll-up at lower Reynolds number, the main features of the vortex flow are similar to those at higher Reynolds number. The strength of the vortices at the trailing-edge was proportional to the time-averaged drag coefficient, which increased by nearly 50% in the range of upsweep angles tested. The vortex was more coherent with reduced meandering and a smaller core radius towards the trailing-edge. This reduction in meandering along the streamwise direction had not been observed previously with other external vortex flows in aerodynamics. Proper Orthogonal Decomposition revealed that the helical displacement mode with azimuthal wavenumber m = 1 was the dominant mode towards the trailing-edge, suggesting that the afterbody vortices bear much similarity with the more widely studied wing tip vortices and delta wing vortices. The instantaneous vortex pair exhibits time-dependent asymmetry; however, there is virtually no correlation between the displacements of the vortex centers.

Attenuation of self-excited roll oscillations of low-aspect-ratio wings using acoustic excitation was studied in a wind tunnel. For a rectangular flat-plate wing with an aspect ratio of 2, roll oscillations can be completely suppressed, and the onset of the roll oscillations can be delayed with external acoustic excitation. Similar results were also obtained for wings with two different airfoil profiles. Velocity measurements indicated that acoustic excitation could restore a symmetric vortex flow over the free-to-roll wings, thus eliminating the self-excited roll oscillations. The effect of excitation is most noticeable for the side of the wing that sees a larger effective angle of attack due to the rolling motion. Acoustic excitation energizes the shear-layer instabilities and results in reattachment or smaller separated flow region closer to wing surface, thus in turn suppressing the roll oscillations.

There is great interest in small aircraft known as Micro Air Vehicles and mini Unmanned Air Vehicles due to the wide range of possible applications. This article reviews recent work that aims to exploit the flexibility of the wing structure in order to increase lift and thrust, and delay stall. Wing flexibility has often been considered to be unwanted for large conventional aircraft and measures are taken to limit the deformation. In contrast, very small aircraft flying at low speeds are not necessarily subject to the same limitation. This approach is only applicable to small aircraft because the frequencies of the wing structure and fluid flow instabilities are close to each other. Consequently, small amplitude and high-frequency motions will be considered. We first start with rigid airfoils and wings in forced plunging motion, which mimics the bending oscillations. The main advantage of this approach is the freedom to vary the frequency within a wide range. Two mechanisms of high-lift production on the oscillating rigid airfoils are discussed. In the first one, leading-edge vortex dynamics and different modes of vortex topology play an important role on the time-averaged lift and thrust at post-stall angles of attack. Existence of optimal frequencies and amplitudes are demonstrated, and their relation to other phenomena is discussed. In the second mechanism of high-lift, trailing-edge vortex dynamics leads to bifurcated/asymmetric flows at pre-stall angles of attack. Deflected wakes can lead to time-averaged lift coefficients higher than those for the first mechanism. Some aspects of lift enhancement can be sensitive to the airfoil shape. For three-dimensional finite wings, lift enhancement due to the leading-edge vortices and existence of optimal frequencies are similar to the two-dimensional case. Vortex dynamics of the leading-edge vortex and tip vortex is discussed in detail. Leading-edge sweep is shown to be beneficial in the reattachment of the separated flows over oscillating wings. Oscillating flexible wings can provide much higher lift coefficient than the rigid ones. Amplitude and phase variation in the spanwise direction result in much stronger leading-edge and tip vortices. Self-excited vibrations of flexible wings, including membrane wings, can excite shear layer instabilities, and thus delay stall and increase lift. Finally, thrust enhancement or drag reduction can be achieved by employing chordwise and spanwise flexibility. The effects of wing flexibility on the vortices and thrust/drag are discussed in relation to the characteristics of wing deformation.

Experiments in a water tunnel have been carried out on low-aspect-ratio rectangular wings undergoing a small-amplitude harmonic plunge motion at Reynolds numbers 10,000 and 20,000. A series of measurement techniques have been used, including force measurements, hot film, particle image velocimetry, and volumetric velocimetry measurements, to study the lift enhancement as a function of forcing frequency. Multiple peaks in the time-averaged lift have been observed, occurring at frequencies in the order of natural vortex-shedding frequencies of the stationary wings. It is postulated that interaction between the leading-edge and trailing-edge vortices contributes to the selection of the optimal frequencies for the time-averaged lift. At a specific Strouhal number, an adverse interaction between the vortices results in a vortex dipole that directs flow upstream. A comparison between a NACA 0012 and flat-plate profile provides further insight into the advantages and disadvantages of using a thinner profile in l...

Active control of self-excited roll oscillations of a rectangular flat plate wing with an aspect ratio of two was studied experimentally in a wind tunnel, using synthetic jet excitation near the leading edge. It was found that, by activating the synthetic jet excitation at an optimum frequency of St = 1, large amplitude roll oscillations could be attenuated and the onset of the oscillation can be delayed by up to Δαmax = 3.5° for extremely small values of momentum coefficient. High frame-rate Particle Image Velocimetry (PIV) measurements revealed a strong resonance between the synthetic jet excitation and shear layer instabilities. The resonance energizes the shear layer separated from the leading edge and results in a local flow field that is more typical of lower wing incidence, thus effectively suppressing roll oscillations.

Experiments were conducted on rectangular wings with an aspect ratio of two, at a chord Reynolds number Re c = 46 , 000 , in order to understand how membrane wings behave in a prescribed transient motion. As a reference, high-speed Digital Image Correlation (DIC) and Particle Image Velocimetry (PIV) measurements were performed on stationary membrane wings in the range 0 ° α 25 ° . Four distinct incidence regions were identified based on the mode shape, frequency and amplitude of the surface vibrations. Flow-field measurements indicated that the different regions were caused by the interactions of the leading-edge separated shear layer with the membrane. Then the membrane wing was subjected to transient sinusoidal pitching manoeuvres at starting incidences in each of these regions. The membrane deformation characteristics were seen to vary considerably with starting angle; both in terms of time-averaged and instantaneous quantities. In some cases there was significant time-lag in the response of the membrane. The PIV flow field measurements for this case showed signs of hysteresis between the pitch-up and pitch-down parts of the wing motion, which was not the case for a rigid wing with identical planform.

Two high-lift mechanisms, namely, convected leading-edge vortices and stable deflected jets, have previously been identified for an airfoil undergoing small-amplitude plunging oscillations. In this paper the effect of geometry is investigated through direct comparison of the forces and flowfields associated with small-amplitude plunging oscillations of a NACA 0012 airfoil and a flat plate for 0 deg and a poststall angle of attack of 15 deg with a Reynolds number of 10,000. For 0 deg at high Strouhal numbers, the NACA airfoil experiences stable deflected jets responsible for very large lift coefficients, whereas the flat plate experiences deflected jets that are prone to periodic oscillation in direction, resulting in oscillation of the lift coefficient with a period on the order of 100 cycles. It is postulated that this jet switching is driven by the leading-edge vortex (LEV). At 15 deg angle of attack, the flat plate is shown to produce a comparable increase in lift up to a Strouhal number of unity, but ...