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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.

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.

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.

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.

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.

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.

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.

Drag reduction for hydrofoils is studied through thrust generation on foils plunging at low Strouhal numbers in order to simulate the action of the ocean waves. Force, deformation and flow field measurements are presented for a partially flexible plunging foil in water tunnel experiments. The foil is predominantly rigid with a short flexible trailing-edge plate of length: L=0.1c, 0.2c, or 0.3c. Using flexible plates, whose natural structural frequency is much higher than the frequency of the plunge oscillations, increases thrust compared to the rigid case. Flexibility is generally more effective for larger lengths of the flexible plate and smaller plunge amplitudes. The maximum observed is therefore for the largest length and smallest amplitude studied: L=0.3c and a=0.1c and equates to 28% more thrust than the rigid case. Optima are observed in the non-dimensional rigidity (λ) versus flap angle amplitude (δ, which is a measure of the relative deformation) parameter space. These occur at λ≈2 and δ≈7–13° for a wide range of flexible plate length and plunge amplitude. Whilst a satisfactory explanation of why there is an optimal flap amplitude remains unavailable, the case of optimal flap angle amplitude results in increased trailing-edge vortex circulation, giving a stronger reverse Kármán vortex street and thus a stronger time-averaged jet.

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.

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.

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 ...

In this paper the combined effect of two mechanisms for lift enhancement at low Reynolds numbers are considered, wing oscillations and wing flexibility. The force, deformation and flow fields of rigid and flexible low aspect ratio (AR=3) and high aspect ratio (AR=6) wings oscillating at a fixed post-stall angle of attack of 15° and amplitude of 15% of chord are measured. The force measurements show that flexibility can increase the time-averaged lift coefficient significantly. For low aspect ratio wings the maximum lift coefficient across all Strouhal numbers was Cl=1.38 for the rigid wing as opposed to Cl=2.77 for the flexible wing. Very similar trends were observed for the high aspect ratio wings. This increase is associated with significant deformation of the wing. The root is sinusoidally plunged with small amplitude but this motion is amplified along the span resulting in a larger tip motion but with a phase lag. The amount it is amplified strongly depends on Strouhal number. A Strouhal number of Src=1.5 was selected for detailed flow field measurements due to it being central to the high-lift region of the flexible wings, producing approximately double the lift of the rigid wing. For this Strouhal number the rigid wings exhibit a Leading Edge Vortex (LEV) ring. This is where the clockwise upper-surface LEV pairs with the counter-clockwise lower-surface LEV to form a vortex ring that self-advects upstream and away from the wing's upper surface. Conversely the deformation of the flexible wings inhibits the formation of the LEV ring. Instead a strong upper-surface LEV forms during the downward motion and convects close to the airfoil upper surface thus explaining the significantly higher lift. These measurements demonstrate the significant gains that can be achieved through the combination of unsteady aerodynamics with flexible structures.

Interaction of strake and wing vortices over a 70°/50° double delta wing were studied experimentally in a wind tunnel using particle image velocimetry (PIV) measurements. The upstream effect of the wing vortex on the formation of the strake vortex was identified. A dual-vortex structure of the strake vortices was observed before the wing vortex developed. Further downstream, wing and strake vortices rotated around each other slowly initially, and then faster with downstream distance, at an increasing rate with increasing incidence. Prior to vortex breakdown, both wing and strake vortices were found meandering in relatively small regions. The correlation between the instantaneous locations of the vortices increases if the vortices become sufficiently close to each other. The proper orthogonal decomposition (POD) analysis of the instantaneous velocity fields suggested that, for both wing and strake vortices, the most energetic mode was displacement in the first helical mode. The most energetic mode reveals out-of-phase displacements when the vortices are close to each other.

Force and particle image velocimetry measurements were conducted on a NACA 0012 airfoil undergoing small-amplitude sinusoidal plunge oscillations at a poststall angle of attack and Reynolds number of 10,000. With increasing frequency of oscillation, lift increases and drag decreases due to the leading-edge vortices shed and convected over the suction surface of the airfoil. Within this regime, the lift coefficient increases approximately linearly with the normalized plunge velocity. Local maxima occur in the lift coefficient due to the resonance with the most unstable wake frequency, its subharmonic and first harmonic, producing the most efficient conditions for high-lift generation. At higher frequencies, a second mode of flowfield occurs. The leading-edge vortex remains nearer the leading edge of the airfoil and loses its coherency through impingement with the upward-moving airfoil. To capture this impingement process, high-fidelity computational simulations were performed that showed the highly transitional nature of the flow and a strong interaction between the upper and lower-surface vortices. A sudden loss of lift may also occur at high frequencies for larger amplitudes in this mode.

M ANNED fighter aircraft and unmanned combat air vehicles may require high angle-of-attack maneuvers where the roll dynamics become crucial. Small unmanned air vehicles (UAVs) and fixed-wingmicro air vehicles often have separated flows [1,2]. These small vehicles fly at relatively high angles of attack and close to the stall conditions due to the poor lift. Vertical gusts might induce roll instabilities at high angles of attack. This possibility is suggested based on the recent experiments in which rectangular, elliptical, and Zimmerman planforms exhibited self-induced roll oscillations, even before the stall angle [3,4]. Most of the knowledge on the aerodynamics of free-to-roll wings is on slender delta wings [5–7], which exhibit well understood freeto-roll oscillations about a zero mean roll angle. Free-to-roll aerodynamics of delta wings with lower sweep angles (or higher aspect ratios) have been investigated in recent experiments [8–14]. The existence of equilibrium at nonzero roll angles was reported for sweep angles 65 . The roll asymmetries are demonstrated in Fig. 1a for a simple delta wing with sweep angle of 55 (taken from [14]). The variation of themean roll angle as a function of angle of attack shows that nonzero roll angles are observed until the wing stalls. Here, we use the term “stall” to describe the onset of the flow regime for which there is no flow reattachment, and hence the roll angle becomes zero. Previouswork [12] confirms that the “stall angle of the free-to-roll wing” is near the stall angle of the fixed wing (at zero roll angle) at which the lift force is maximum. The stall of the free-to-roll wing is very sudden as the angle of attack is increased. In contrast, for a slender delta wing with sweep angle 70 , roll asymmetries mostly disappear, as shown in Fig. 1b. Contour plots in Fig. 2 are adapted from the data in [14], which show the mean roll angle and the standard deviation of roll angle in the free-to-roll experiments. The variation of the angle of attack at which the vortex breakdown appears at the trailing edge of the wing, as well as the stall angle of attack at which the lift is maximum (for zero roll angle), is added from another source [15] (shown as dashed lines). It is seen in Fig. 2a that, for wings with low sweep angles, roll asymmetries appear when the vortex breakdown is over thewing and becomes maximum near the stall angle. The magnitude of the mean roll angle decreases with increasing sweep angle and almost disappears for the most slender wing. The stall of the free-to-roll wing also becomes more gradual with increasing sweep angle. Figure 2b shows that, in a certain range of sweep angles and near the stall, self-induced roll oscillations are possible. In this Note, we extend our experiments to cropped delta wings with an initial leading-edge sweep angle of 55 , as shown in Fig. 3. We anticipate that the slenderness ratio, defined as the chord length divided by the span, may be important. It was shown [7] that if the distance from the vortex origin divided by the span is larger than two, the wing rock of aircraft configurations is likely. Similarly, this ratio may be important for the onset of the roll asymmetries for the nonslender free-to-roll wings. Table 1 lists the aspect ratio and slenderness ratio for the wings shown in Fig. 3. For comparison, we also added the simple delta wing with 70 into our study.

Aerodynamic characteristics of two-dimensional membrane airfoils were experimentally investigated in a wind tunnel. The effects of the membrane pre-strain and excess length on the unsteady aspects of the fluid–structure interaction were studied. The deformation of the membrane as a function of angle of attack and free-stream velocity was measured using a high-speed camera. These measurements were complemented by the measurements of unsteady velocity field with a high frame-rate Particle Image Velocimetry (PIV) system as well as smoke visualization. Membrane airfoils with excess length exhibit higher vibration modes, earlier roll-up of vortices, and smaller separated flow regions, whereas the membranes with pre-strain generally behave more similarly to a rigid airfoil. Measured frequencies of the membrane vibrations suggest a possible coupling with the wake instabilities at high incidences for all airfoils.

The effects of unsteady trailing-edge blowing on delta wing aerodynamics were investigated experimentally to understand the aerodynamics-propulsion interaction for dynamic thrust vectoring. Two models with sweep angles of A = 50 and 65 deg, representing nonslender and slender delta wings, respectively, were tested in a water tunnel. Flow visualization and velocity and force measurements were conducted at stall and poststall incidences. For the periodic blowing, it was found that the dynamic response of leading-edge vortex breakdown and wing normal force coefficient exhibit phase lags for both nonslender and slender delta wings. The estimated time constants are larger than those reported in the literature for unsteady wings undergoing pitching or plunging. In the case of transient blowing, the time delay for the decelerating jet is significantly larger than that for the accelerating jet. The variation of the circulation and the reattachment process near the wing surface were studied by means of velocity measurements. The range of the estimated time constants is similar at the stall and poststall incidences for both the slender and nonslender wings.

Experiments were conducted on various free-to-roll delta wings with sweep angles in the range of 40―70 deg in a wind tunnel. For low sweep angles, nonzero trim angles are observed until the wing stalls after which the trim angle becomes zero. The magnitude of the roll trim angle decreases with increasing sweep angle and almost disappears for the most slender wing. In a certain range of sweep angles, and just before the stall, large amplitude self-induced roll oscillations with nonzero mean were observed for thin wings with sharp leading edges for which separation point is fixed. Two different regions of angle of attack with free-to-roll oscillations were found for certain sweep angles. Velocity measurements suggest that the oscillations are initiated by the increasing proximity of the shear layers as the angle of attack is increased. Possible mechanisms of the self-induced roll oscillations based upon the phase-averaged velocity measurements have been proposed, incorporating varying vortex size and strength as well as the movement of the shear-layer reattachment locations. Hysteresis and time lag effects are significant even though typical dimensionless frequencies are very low.

The aerodynamics of thin, flat-plate wings of various planforms (rectangular, elliptical and Zimmerman) have been studied in free-to-roll experiments in a wind tunnel. Non-zero trim angles at low angles of attack, self-induced roll oscillations with increasing angle of attack and even autorotation in some cases were observed. The rectangular wings with round leading-edge had non-zero trim angles at low incidences due to the asymmetric development of the three-dimensional separation bubble at these low Reynolds numbers. With increasing angle of attack, the bubble increases in length and once reattachment is lost, large amplitude roll oscillations develop. The Strouhal number of the roll oscillations is of the order of 10−2, which is in the same range as those expected for small aircraft experiencing atmospheric gusts. Velocity measurements revealed that variations in the strength of the vortices drove the rolling motion. At the mean roll angle, because of the time lag in the strength of the vortices, an asymmetric flow is generated, which results in a net rolling moment in the direction of the rolling motion.

Computational simulations have been performed for a 50-deg-sweep delta wing with a sharp leading edge at a 15 deg angle of attack and moderate Reynolds numbers of Re = 2 x 10 5 , Re = 6.2 x 10 5 , and Re = 2 x 10 6 . A sixth-order compact-difference scheme with an eighth-order low-pass filter is used to solve the Navier-Stokes equations. Turbulence modeling has been accomplished using an implicit large eddy simulation method that exploits the high-order accuracy of the compact-difference scheme and uses the discriminating higher-order filter to regularize the solution. Computations have been performed on a baseline mesh of 11.3 x 10 6 grid points and a refined mesh of 35 x 10 6 grid points. An assessment of grid resolution showed that significantly finer-scale features of the flow could be captured on the refined mesh, providing a more accurate representation of the complex, unsteady, separated flow. Comparisons are also made with high-resolution particle image velocimetry images obtained for the two lower Reynolds numbers. The numerical results are examined to provide a description of the mean and instantaneous flow structure over the delta wing, including the separated vortical flow, vortex breakdown, surface flow features, and surface boundary-layer transition near the symmetry plane. The effect of Reynolds number on each of these features is assessed.

The effects of active flow control by oscillatory blowing at the leading edge of a nonslender delta wing with a Λ=50° sweep angle have been investigated. Pressure measurements and Particle Image Velocimetry measurements were conducted on a half wing to investigate the formation of leading edge vortices for oscillatory blowing, compared to the stalled flow for the no blowing case. Stall has been delayed by up to 8, and significant increases in the upper surface suction force have been observed. Velocity measurements show that shear layer reattachment is promoted with forcing, and a vortex flow pattern develops. The time averaged location of the centre of the vortical region moves outboard with increased excitation. The near-surface flow pattern obtained from the PIV measurements shows reattachment in the forward part of the wing. There is no measurable jet-like axial flow in the vortex core, which seems to break down at or very near the apex. This highlights that unlike slender delta wings, vortex breakdown is not a limiting factor in the generation of lift for nonslender delta wings. Phase averaged measurements reveal the perturbation due to the pulsed blowing, its interaction with the shear layer and vortex, apparent displacement of the vortex core, and relaxation of the reattachment region. The flow in a phase averaged sense is highly three dimensional. Experiments indicate that unsteady blowing at Strouhal numbers in the region of St=0.5 to St=0.75, and in the region of St=1.25 to St=1.5 can be a highly effective. Reattached flow can develop from stalled flow after pulsing has been initiated with a time constant of tU/c=5 for unsteady blowing at St=0.75, and tU/c=7 for St=1.5. Experiments with excitation from finite span slots located in the forward half of the wing show that partial blowing can be more effective at low momentum coefficients. Force measurements of a full delta wing confirmed that the effectiveness of this method of flow control was not only confined to half delta wings.

Effects of small-amplitude wing oscillations on flow reattachment and vortex re-formation in the poststall region were studied experimentally. Low-aspect-ratio delta and cropped delta wings with sweep angles in the range of A = 20-50 deg were tested. Effects of forcing frequency, amplitude, wing sweep angle, and mode of the wing oscillations were investigated. It was shown that wing oscillations at an appropriate frequency promotes the reattachment of the otherwise separated flow at poststall incidences. Moreover, with increasing frequency, leading-edge vortices develop with axial flow in the cores, followed by their breakdown. For an optimum range of Strouhal numbers, Sr = 1-2, vortex breakdown is delayed to a maximum distance from the apex. This range of Strouhal numbers compares well with the dominant frequencies of the shear layer instabilities. For sweep angles larger than 20 deg, the reattachment process is generic for all nonslender wings and there is an optimum range of Strouhal numbers.

A water tunnel study of the effect of spanwise flexibility on the thrust, lift and propulsive efficiency of a rectangular wing oscillating in pure heave has been performed. The thrust and lift forces were measured with a force balance, and the flow field was measured with a Particle Image Velocimetry system. Introducing a degree of spanwise flexibility was found to be beneficial. For Strouhal numbers greater than 0.2, a degree of spanwise flexibility was found to yield a small increase in thrust coefficient, and a small decrease in power-input requirement, resulting in higher efficiency. In this case, a moderately stronger trailing-edge vortex system was observed. Introducing a far greater degree of spanwise flexibility, however, was found to be detrimental. A large phase delay of the wing tip displacement was observed, leading to the root and tip moving in opposite directions for a significant portion of the flapping stroke. Vorticity of opposing sign was observed to be shed from the root and tip, resulting in a weak and fragmented vorticity pattern. The thrust coefficient was observed to be significantly reduced, and the efficiency diminished. It is noted that the range of Strouhal numbers for which spanwise flexibility was found to offer benefits overlaps the range found in nature, of 0.2

Unsteady aerodynamics of flexible nonslender delta wings is investigated in an experimental study using various techniques. Dramatic fluid/structure interactions emerge with increasing wing flexibility and result in substantial lift enhancement in the poststall region. This recently discovered phenomenon appears to be a feature of nonslender wings. Self-excited antisymmetric vibrations of the wing promote reattachment of the shear layer, which results in the lift enhancement. These self-excited vibrations are not observed for a half-model. The Strouhal number of the dominant frequency of the structural vibration is on the order of unity for all nonslender wings, which also corresponds to the frequency of the shear-layer instabilities for the rigid wing. Velocity measurements demonstrate the striking difference between the flows over the flexible and rigid wings in the poststall region. The effect of flexibility is to promote the reattachment of the shear layer near or downstream of the apex, depending on the incidence. There are substantial effects on the vortical flow with increasing wing flexibility, which might lead to the axial flow developing within the reattached region. The time-averaged vorticity flux increases due to the oscillating leading edge, which leads to increased circulation.

The interaction of thrust-vectoring jets with leading-edge vortices over stationary delta wings and its effects on the wing aerodynamics were investigated experimentally. Two models with sweep angles of A = 50 and 65 deg, representing nonslender and slender wings, respectively, were tested with rectangular and circular nozzles. Force, velocity measurements, and flow visualization were performed. It was found that undervortex blowing can significantly affect the aerodynamic forces on both nonslender and slender wings. The maximum lift change occurred near the stall angle, due to the earier reattachment of shear layer and delay of vortex breakdown. Force measurements revealed that the effect of nozzle geometry can be important, because the entrainment effect of the jet depends on it. The jet-vortex interact on, distortion of jet vortices, and merging of wing and jet vortices are more pronounced for the rectangular nozzle and have a larger influence on the delta wing aerodynamics. The effect of the jet yaw angle is small for the nonslender wing, whereas the aerodynamics of the slender wing is very sensitive to the jet yaw angle.

Vortex control concepts employed for slender and nonslender delta wings were reviewed. Important aspects of flow control include flow separation, vortex formation, flow reattachment, vortex breakdown, and vortex instabilities. The occurrence and relative importance of these phenomena strongly depend on the wing sweep angle. Various flow control methods were discussed: multiple vortices, control surfaces, blowing and suction, low-frequency and high-frequency excitation, feedback control, passive control with wing flexibility, and plasma actuators. For slender delta wings, control of vortex breakdown is achieved by modifications to swirl level and external pressure gradient acting on the vortex core. Effects of flow control methods on these two parameters were discussed, and their effectiveness was compared whenever possible. With the high-frequency excitation of the separated shear layer, reattachment and lift enhancement in the post-stall region is observed, which is orders of magnitude more effective than steady blowing. This effect is more pronounced for nonslender wings. Re-formation of vortices is possible with sufficient amplitude of forcing at the optimum frequency. Passive lift enhancement on flexible wings is due to the self-excited wing vibrations, which occur when the frequency of wing vibrations is close to the frequency of the shear layer instabilities, and promote flow reattachment.

The flow behind two side-by-side circular cylinders is experimentally investigated based on laser-illuminated flow-visualization, particle image velocimetry and hot-wire measurements. The flow is classified as three regimes: single street (the cylinder center-to-center spacing T/d 2.0). Special attention is given to the regime of 1.2 2.0, the flow structure change from the anti-phase to in-phase mode starts with a phase shift between gap vortices. The dynamical role of gap bleeding between cylinders for L/d

The vortex-induced structural vibration of an elastic square cylinder, on fixed supports at both ends, in a uniform cross flow was measured using fiber-optic Bragg grating sensors. The measurements are compared to those obtained for an elastic circular cylinder of the same hydraulic diameter in an effort to understand the effect of the nature (fixed or oscillating) of the flow separation point on the vortex-induced vibration. It is found that a violent vibration occurs at the third-mode resonance when the vortex-shedding frequency coincides with the third-mode natural frequency of the fluid-structure system, irrespective of the cross-sectional geometry of the cylinder. This is in distinct contrast to previous reports of flexibly supported rigid cylinders, where the first-mode vibration dominates, thus giving little information on the vibration of other modes. The resonance behavior is neither affected by the incidence angle (α) of the free stream, nor by the nature of the flow separation point. However, the vibration amplitude of the square cylinder is about twice that of the circular cylinder even though the flexural rigidity of the former is larger. This is ascribed to a difference in the nature of the flow separation point between the two types of structures. The characteristics of the effective modal damping ratios, defined as the sum of structural and fluid damping ratios, and the system natural frequencies are also investigated. The damping ratios and the system natural frequencies vary little with the reduced velocity at α=0deg, but appreciable at α⩾15deg; they further experience a sharp variation, dictated by the vortex-shedding frequency, near resonance.

A fiber-optic Bragg grating (FBG) sensor is proposed to measure the local static temperature θ s and the fluctuating temperature θ s on the surface of a heated circular cylinder subjected to a cross-flow. In order to validate the new technique, a type- K thermocouple and a single hot wire are used to measure θ s and the near-wake streamwise fluctuating velocity u , respectively. The FBG sensor measurement of θ s agrees well with that simultaneously obtained by the thermocouple. The θ s -spectrum exhibits a prominent peak at the vortex shedding frequency, f s ∗ =f s d/U ∞ ≈0.2 ( U ∞ is the free-stream velocity and d is the cylinder diameter), consistent with the u -spectrum. In fact, the θ s and u signals are almost perfectly correlated at f s ∗ . Furthermore, the measured θ s and root mean square value of θ s are agreeable with previously reported Nusselt number data. These results demonstrate that the FBG sensor can be used to measure reliably both static and fluctuating temperatures. It is expected that the FBG sensor, because of its uniqueness in many aspects and being at an affordable cost, has an excellent prospect in temperature measurements.

Interference effects on vortex-induced vibrations of two side-by-side elastic cylinders, fixed at both ends (with no deflection and displacement) in a cross-flow, were experimentally investigated. The dynamic responses of the cylinders were measured using two fiber-optic Bragg grating (FBG) sensors. Simultaneously, a single hot wire was used to measure the velocity in the wake. It has been previously observed that violent resonance occurs when transverse cylinder spacing ratio, T/d, is either large (>2.0) or small (2.0 or T/d

This note reports an experimental study of the fluid damping of a long slender cylinder, fixed at both ends (no rotation and displacement), in a cross air-flow. The structural dynamic strain was measured in the lift direction over a range of reduced velocity Ur using a fibre-optic Bragg grating sensor. An auto-regressive moving average technique was used to deduce the effective damping ratios (including structural and fluid damping) from the strain data. The modal damping ratios corresponding to the first-, second- and third-mode vibrations over a range of Ur from 3 to 45 have been characterized and discussed. The results show that the value of fluid damping varied significantly at resonance when the vortex-shedding frequency coincides with one of the natural frequencies of the combined fluid–structure system.

Free vibrations of two side-by-side cylinders with fixed support (no rotation and displacement) at both ends placed in a cross-flow were experimentally investigated. Two fibre-optic Bragg grating sensors were used to measure the dynamic strain, while a hot wire and flow visualization were employed to examine the flow field around the cylinders. Three T/d ratios, 3.00, 1.70 and 1.13, were investigated, where T is the centre-to-centre cylinder spacing and d is the diameter; they give rise to three different flow regimes. The investigation throws new light on the shed vortices and their evolution. A new interpretation is proposed for the two different dominant frequencies, which are associated with the narrow and the wide wake when the gap between the cylinders is between 1.5 and 2.0 as reported in the literature. The structural vibration behaviour is closely linked to the flow characteristics. At T/d = 3:00, the cross-flow root-mean-square strain distribution shows a very prominent peak at the reduced velocity Ur ≈ 26 when the vortex shedding frequency fs, coincides with the third-mode natural frequency of the combined fluid–cylinder system. When T/d < 3:00, this peak is not evident and the vibration is suppressed because of the weakening strength of the vortices. The characteristics of the system modal damping ratios, including both structural and fluid damping, and natural frequencies are also investigated. It is found that both parameters depend on T/d. Furthermore, they vary slowly with Ur, except near resonance where a sharp variation occurs. The sharp variation in the natural frequencies of the combined system is dictated by the vortex shedding frequency, in contrast with the lock-in phenomenon, where the forced vibration of a structure modifies the vortex shedding frequency. This behaviour of the system natural frequencies persists even in the case of the single cylinder and does not seem to depend on the interference between cylinders. A linear analysis of an isolated cylinder in a cross-flow has been carried out. The linear model prediction is qualitatively consistent with the experimental observation of the system damping ratios and natural frequencies, thus providing valuable insight into the physics of fluid–structure interactions.

Large amplitude roll oscillations are inherent to fixed-wing MAVs. This paper reviews our recent works on the suppression of the self-induced roll oscillations of LAR rectangular wings using active flow control techniques. Both acoustic excitation and synthetic jet blowing have been used to attenuate the self-excited roll oscillations. Three rectangular wings, of flat plate, NACA0012 and SD7003-085-88 profiles were tested. It was found that roll oscillations can be completely suppressed and the onset of the roll oscillations can be delayed by active flow control approaches. PIV measurements indicated that the excitations could restore a symmetric vortex flow over free-to-roll wings thus stabilize the roll oscillations.

To identify the role of flexibility in lift generation, the force, deformation and flow fields of rigid and flexible low aspect ratio (sAR = 1.5) wings plunging with a fixed post-stall angle of attack of 15° and amplitude of 15% of chord were experimentally measured. It was shown that spanwise flexibility can significantly enhance lift. The peak increase in lift coefficient over the stationary case is more than three times larger for the flexible wing compared to the rigid wing. This increase is associated with significant deformation of the wing. The root is sinusoidally plunged with small amplitude but this motion is amplified along the span resulting in a larger tip motion. The amount it is amplified strongly depends on Strouhal number. A Strouhal number of Src = 1.5 was selected for detailed flow field measurements due to it being associated with high lift, near the natural frequency, and comparable with a previous study for high aspect ratio wings. For this Strouhal number the rigid wing exhibits a LEV dipole. This is where the clockwise upper-surface LEV pairs with the counterclockwise lower-surface LEV to form a vortex ring that self-advects upstream and away from the wing's upper surface. Conversely the flexible wing experiences deformation resulting in tip amplitude 1.53 times greater than the root amplitude and tip phase lag of 98°. This deformation inhibits the LEV dipole. Instead a strong upper-surface LEV forms during the downward motion and convects close to the airfoil upper surface thus explaining the significantly higher lift.

Experiments were performed to investigate the effects of amplitude and depth on the drag reduction of a NACA 0012 airfoil plunging near a free surface for a range of frequencies. Beyond the effect of the free surface, at low Strouhal numbers based on amplitude, Sr A, the drag reduction follows a parabolic trend with greater effect for greater amplitude, similar to the Garrick predictions. At Sr A ≈ 0.08, larger amplitudes break from this trend due to leading-edge vortex formation. As a result, smaller amplitudes become preferable for Sr A > 0.12. In addition, for the first time, vortex lock-in is documented experimentally. The effect of depth is twofold; firstly with decreasing depth, there is a general departure from the Garrick trends. Secondly, a reduction in thrust is observed around a constant unsteady parameter of τ = U ∞ 2πf/g ≈ 0.25; around this value significant free-surface waves form that detract from thrust creation. For depths greater than two chord lengths, there is negligible free-surface effect.

Force and particle image velocimetry measurements were conducted on a NACA 0012 aerofoil undergoing small-amplitude high-frequency plunging oscillation at low Reynolds numbers and angles of attack in the range 0–$2{0}^{\ensuremath{\circ} } $. For angles of attack less than or equal to the stall angle, at high Strouhal numbers, significant bifurcations are observed in the time-averaged lift coefficient resulting in two lift-coefficient branches. The upper branch is associated with an upwards deflected jet, and the lower branch is associated with a downwards deflected jet. These branches are stable and highly repeatable, and are achieved by increasing or decreasing the frequency in the experiments. Increasing frequency refers to starting from stationary and increasing the frequency very slowly (while waiting for the flow to reach an asymptotic state after each change in frequency); decreasing frequency refers to impulsively starting at the maximum frequency and decreasing the frequency very slowly. For the latter case, angle of attack, starting position and initial acceleration rate are also parameters in determining which branch is selected. The bifurcation behaviour is closely related to the properties of the trailing-edge vortices. The bifurcation was therefore not observed for very small plunge amplitudes or frequencies due to insufficient trailing-edge vortex strength, nor at larger angles of attack due to greater asymmetry in the strength of the trailing-edge vortices, which creates a preference for a downward deflected jet. Vortex strength and asymmetry parameters are derived from the circulation measurements. It is shown that the most appropriate strength parameter in determining the onset of deflected jets is the circulation normalized by the plunge velocity.