Session H25: Aerodynamics: Fixed, Flapping and Rotating Wing II

Unsteady forces on the wing subjected to spanwise twisting

Force analysis and flow visualization on the wing subjected to spanwise twisting are studied at constant Reyolds number 18000. The wing is designed to change the shape in a spanwise direction with a flexion ratio of 0.6. The wing morphing time is kept at 1 second.  The spanwise twisting causes the spanwise variation in the angle of attack. Due to the spanwise variation of the angle of attack, there will be a vorticity gradient along the span of the wing. This type of wing morphing can be used to achieve higher stability during unsteady maneuvers. The maneuvers are inspired by animal locomotions. Most of the natural flyers and swimmers change the shape of the wing to get extra stability and higher maneuverability during their attack and escape maneuvers. Wing morphing not only controls the inertial forces but can be used to control circulatory forces by controlling the formation of edge vortices. In this work, we have demonstrated how we can quantitatively and qualitatively change both inertial and circulatory forces for the wing which can perform controoled spanwise angle of attack of variation.   

Determination of Flow Field Parameters using Inverse Interpolation Methods

Full flow field parameters cannot be directly measured in-flight and therefore, an inverse approach is required. This study utilizes surface data (e.g., sensors) and advances the inverse interpolation method to predict flow field characteristics for complex maneuvers (e.g., pitching, plunging, pitching-plunging, wind gusts) in real-time based on a pre-computed database. The inverse interpolation formulation is based on a parametric approximation of the surface conditions and is further constructed through a least-squares minimization of calculated and measured surface conditions. This procedure results in the governing system of linear algebraic equations, providing the unknown coefficients that accurately define the surface conditions in an efficient manner. Broadly, this inverse approach solves for the flow field based on surface conditions as opposed to solving it based on freestream conditions.   

Investigation of surface pressure fluctuations in airfoils subject to transverse gust: towards gust mitigation

One of the key issues in developing unmanned aerial vehicles (UAVs) capable of flying in adverse conditions is an understanding of their unsteady response when subject to fluid disturbances. Sudden gusts due to complex terrain or unfavorable weather cause large lift transients that can have detrimental consequences on UAV flight stability, efficiency, and maneuverability. To investigate the effect of a transverse gust, we perform large-eddy simulations of a gust impinging on a NACA 0012 airfoil at low Reynolds numbers and explore its influence on the pressure fluctuations at the airfoil surface. We find that these surface measurements may be used to inform predictions of the lift coefficient during gust encounters, paving the way for designing gust mitigation schemes that stabilize flight in UAVs.   

Unsteady Evolution of a Laminar Separation Bubble Subjected to Wing Structural Motion

Wind tunnel experiments and direct numerical simulations are employed to investigate the laminar separation bubble that forms on the suction side of a static and plunging wing section for a modified NACA64(3)−618 airfoil at a chord Reynolds number of Re=200k. A plunging motion with an amplitude of ℎ=6% chord and a reduced frequency of k=0.67 is imposed on the wing at zero degrees angle of attack. Surface pressure, 2D Particle Image Velocimetry and Infrared Thermography measurements are used to track the unsteady evolution of the bubble along the plunging cycle and its effect on the wing loading. Modal analysis of the flow, using Proper Orthogonal Decomposition, shows that the transition process in the separated shear layer is mainly due to the amplification of Kelvin-Helmholtz instabilities, followed by shedding of spanwise coherent structures that lead to turbulent reattachment. A hysteretic behaviour is observed for the bubble size and location during the cycle. No bubble bursting is observed at these conditions, thus having a small impact on the global lift and pitching moment coefficients. Active flow control using dielectric barrier discharge plasma actuators is explored for reducing separation and extending the laminar flow region.   

Impact of Surging Streamwise Flow on Laminar Separation Bubble Dynamics for a Finite-Span Wing

The impact of unsteady streamwise flows on a finite, semi-span, NACA 0015 wing is examined for mean chord-based Reynolds Numbers of 100,000 and 150,000. The unsteady lift response is primarily dominated by the coupling of the surging streamwise flow with the laminar separation bubble dynamics at low to moderate angles of attack, and with the vortex shedding characteristics at angles of attack post-stall. For low reduced frequencies at low angles of attack, the laminar separation bubble elongates at the minimum velocity, thereby increasing the suction pressure on the suction surface and enhancing lift. For low reduced frequencies at post-stall angles of attack, shedding of low pressure coherent structures from the leading edge convect more slowly and coalesce, creating lift plateaus during the lower velocity portion of the cycle, but does not have an impact on the phase averaged lift.   

An inverted wing in extreme ground effect produces downforce with a static peak at h/c~10%. A sinusoidally heaving wing at reduced frequency of k=0.5 has less average downforce than the static mean. Dynamic downforce is approximated from analytical/state-space models based on static chordwise separation.

An inverted single element was subjected to a sinusoidal heaving motion in both free flight and extreme ground effect, with the ground-effect simulations oscillating in various states of interaction with the peak lift ride height of the wing. Peak negative lift during the heaving cycle was greater than the static values at the same ground clearances, time, and ensemble averaging showed an overall reduction in the lift coefficient of 10–22%. An analytical model combining potential flow lift predictions and a new variation of the Goman–Khrabrov state-space model predicts the lift behavior of the wing-in-ground effect based on reduced frequency and ground clearance.   

Visualization and Transport Analysis of Non-slender Delta Wing Vortices

The flow physics of non-slender delta wings are complex and relatively poorly understood in comparison with their slender counterparts.  A vorticity transport analysis on a low-Reynolds number (O(10⁴)), sharp-edged delta wing, with sweep angle of 50 degrees, showed that the vorticity in the leading-edge vortices was generated primarily near the apex, with convective and diffusive vorticity transport processes governing LEV evolution along the leading edges [Buchholz, Wabick, Manazir, and Snider, APS DFD 2019].  In the present work, we examine a similar delta wing geometry in a wind tunnel at Re = O(10⁵) using volumetric particle tracking velocimetry.  A vorticity transport framework is used to characterize leading-edge vortex evolution under these conditions.  Whereas the vorticity transport budget, which quantifies various physical mechanisms governing LEV strength, is in essence a reduced order model of the flow, we also investigate these transport mechanisms in the context of flow field modal decomposition.  Reynolds number effects are also considered.   

Not Participating

The vortex shedding from a square bluffy body has been seen to impact the aerodynamic performance of a NACA-0012 when placed upstream of the airfoil, at a Reynolds number of 40 x 10³. This study aims to understand the impact of the vortex shedding from a bluff body on the aerodynamic performance of the NACA-0012 airfoil. Aerodynamic load experiments were conducted in a water tunnel facility at a Reynolds number of 40 x 10³. Particle image velocimetry (PIV) was used to measure the velocity flow field around the bluff body and over the airfoil. Proper Orthogonal Decomposition (POD) was used to identify the coherent structures and their impact on the aerodynamic performance of the airfoil. The shed vortices from the wake of the bluff body influence the flow behavior over the airfoil and impacting the aerodynamics performance of the airfoil.   

Wake bi-modal switching and its dependence on the upstream bluff body boundary layers

The wake of the squareback Ahmed body exhibits a random bi-modal switching in one of the crossflow directions. This study investigates numerically the sensitivity of wake bi-modality to upstream disturbances related to the boundary layers (BL) on the longitudinal Ahmed body surfaces. Close to the body nose, BL on the top and the side surfaces separate and reattach before reaching the base, where a massive separation occurs forming the wake. Hairpin vortices are generated close to the reattachment points. These vortices grow along the surfaces reaching a critical point -beyond which their interaction with the freestream leads to vortex breakdown. The resulting smaller vortices from all the surfaces interact periodically as they convect downstream. The associated flow disturbances are suggested to be the main trigger of the wake switching. Here, we investigate suppressing the near-nose BL separation using a suction velocity. When only the top-surface BL is suppressed, bi-modal switching still occurs. The BL suppression interrupts the interaction of the vortices upstream of the base and increases the fluctuations of the wake around its asymmetric position. Suppressing BL on both sides is found to completely suppress bi-modal switching and lock the wake in an asymmetric position.   

Wake of an Oscillating Cylinder in Noisy Flow

The wake of a circular cylinder executing undamped transverse vortex-induced vibrations in stochastic noisy flow is numerically analyzed at Reynolds number, Re = 150. The mass ratio (m*) of the cylinder is 2 and the range of reduced speed (U*) is between 1 and 15. A Gaussian noise is added to mean flow velocity at the inlet and the computations are performed using a discrete forcing Immersed Boundary Method (IBM) based in-house fluid flow solver. The vortex-shedding and arrangement of vortices near and away from the oscillating structure is analysed for various values of U* in pre-synchronization, synchronization and post-synchronization zones of the cylinder vibration. For an interesting case in the synchronized zone (U* = 4), in the region close to the cylinder surface, the number of vortices shed in each oscillation cycle is different. Another important observation was the absence of vortex-shedding for certain oscillation cycles. Unlike the vibrations under a uniform flow field at the inlet, the arrangement of vortices in the noisy inflow are highly dynamic and standard modes of vortex shedding are not always recognised. However, in the farwake, rich vortex interaction behaviour like vortex-merging, vortex-pairing and vortex breakdown are observed.