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1. Effect of Reynolds Number on Traveling Wave Flow Control

   https://doi.org/10.2514/6.2023-2137  

   Ogunka, Uchenna E.; Akbarzadeh, Amir; Borazjani, Iman (January 2023, AIAA SCITECH 2023 Forum)

   Large-eddy simulations (LES) of the fluid flow over a NACA0018 airfoil at AOA =20 degrees angle of attack are performed to investigate the effect of surface morphing oscillations on the aerodynamic performance of the airfoil over a wide range of Reynolds numbers (Re = 5,000 to 500,000). These oscillations are in the form of low amplitude backward (opposite to the airfoil's forward motion) traveling wave actuations on the upper surface of the airfoil. The sharp interface curvilinear immersed boundary (CURVIB) method is used to handle the moving surface of the airfoil. The nondimensional amplitude is a*=0.001 (a*=a/L; a: amplitude, L: chord length of the airfoil) and reduced frequency (f*= fL/U; f is the frequency and U is the freestream velocity) is chosen to match the leading edge vortex shedding frequency. The results of the simulations at the post-stall angle of attack (AOA =20 degrees) show that the lift coefficient increases more than 20% and the drag coefficient decreases more than 40% within the Reynolds number range of Re = 50,000-500,000 for traveling wave actuation of amplitude, a*=0.001, and frequency, f*=8. However, the lift and drag coefficients of the actuated airfoil were similar to the baseline airfoil for Re = 5,000. 

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2. Controlling Flow Separation on a Thick Airfoil Using Backward Traveling Waves

   https://doi.org/10.2514/1.J059428  

   Akbarzadeh, A. M.; Borazjani, I. (June 2020, AIAA Journal)

   Large-eddy simulations (LESs) of low-Reynolds-number flow (Re=50,000) over a NACA0018 airfoil are performed to investigate flow control at the stall angle of attack (15 deg) by low-amplitude surface waves (actuations) of different types (backward/forward traveling and standing waves) on the airfoil’s suction side. It is found that the backward (toward downstream) traveling waves, inspired from aquatic swimmers, are more effective than forward traveling and standing wave actuations. The results of simulations show that a backward traveling wave with a reduced frequency f∗=4 (f∗=fL/U, where f is frequency; L, chord length; and U, free flow velocity), a nondimensional wavelength λ∗=0.2 (λ∗=λ/L, where λ is dimensional wavelength), and a nondimensional amplitude a∗=0.002 (a∗=a/L, where a is dimensional amplitude) can suppress stall. In contrast, the flow over the airfoil with either standing or forward traveling wave actuations separates from the leading edge similar to the baseline. Consequently, the backward traveling wave creates the highest lift-to-drag ratio. For traveling waves at a higher amplitude (a∗=0.008), however, the shear layer becomes unstable from the actuation point and creates periodic coherent structures. Therefore, the lift coefficient decreases compared with the low-amplitude case. 

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

   https://doi.org/10.2514/1.J062449  

   Ogunka, Uchenna E; Akbarzadeh, Amir M; Borazjani, Iman (April 2023, AIAA Journal)

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

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4. Flow Control and Separation Delay in Morphing Wing Aircraft Using Traveling Wave Actuation

   https://doi.org/10.1115/SMASIS2020-2355  

   Olivett, A.; Corrao, P.; Karami, M. A. (September 2020, ASME 2020 Conference on Smart Materials, Adaptive Structures and Intelligent Systems)

   null (Ed.)

   This study examines the biomimicry of wave propagation, a mode of locomotion in aquatic life for the use-case of morphing aircraft surfaces for boundary layer control. Such motion is theorized to inject momentum into the flow on the upper surface of airfoils, and as a consequence, creates a forcible pressure gradient thereby increasing lift. It is thought that this method can be used to control flow separation and reduce likelihood of stall at high angles of attack. The motivation for such a mechanism is especially relevant for aircraft requiring abrupt maneuvers, and especially at high angles of attack as a safety measure against stalling. The actuation mechanism consists of lightweight piezoelectric ceramic transducers placed beneath the upper surface of an airfoil. An open-loop system controls surface morphing. A two-dimensional Fourier Transform technique is used to estimate traveling to standing wave ratio, which is verified analytically using Euler Bernoulli beam theory, and experimentally using a prototype wing. Propagating wave control is tuned and verified using a series of scanning laser vibrometry tests. A custom two-dimensional NACA 0018 airfoil tests the concept in a low-speed wind tunnel with approximate Reynolds Number of 50,000. Both traveling waves and the changes in lift and drag will be experimentally characterized. 

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5. Numerical investigation of the effect of airfoil thickness on onset of dynamic stall

   https://doi.org/10.1017/jfm.2019.235  

   Sharma, Anupam; Visbal, Miguel (July 2019, Journal of Fluid Mechanics)

   Effect of airfoil thickness on onset of dynamic stall is investigated using large eddy simulations at chord-based Reynolds number of 200 000. Four symmetric NACA airfoils of thickness-to-chord ratios of 9 %, 12 %, 15 % and 18 % are studied. The three-dimensional Navier–Stokes solver, FDL3DI is used with a sixth-order compact finite difference scheme for spatial discretization, second-order implicit time integration and discriminating filters to remove unresolved wavenumbers. A constant-rate pitch-up manoeuver is studied with the pitching axis located at the airfoil quarter chord. Simulations are performed in two steps. In the first step, the airfoil is kept static at a prescribed angle of attack ( $$=4^{\circ }$$ ). In the second step, a ramp function is used to smoothly increase the pitch rate from zero to the selected value and then the pitch rate is held constant until the angle of attack goes past the lift-stall point. The solver is verified against experiments for flow over a static NACA 0012 airfoil. Static simulation results of all airfoil geometries are also compared against XFOIL predictions with a generally favourable agreement. FDL3DI predicts two-stage transition for thin airfoils (9 % and 12 %), which is not observed in the XFOIL results. The dynamic simulations show that the onset of dynamic stall is marked by the bursting of the laminar separation bubble (LSB) in all the cases. However, for the thickest airfoil tested, the reverse flow region spreads over most of the airfoil and reaches the LSB location immediately before the LSB bursts and dynamic stall begins, suggesting that the stall could be triggered by the separated turbulent boundary layer. The results suggest that the boundary between different classifications of dynamic stall, particularly leading edge stall versus trailing edge stall, is blurred. The dynamic-stall onset mechanism changes gradually from one to the other with a gradual change in some parameters, in this case, airfoil thickness. 

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