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.

Noise produced by aerodynamic interaction between a circular cylinder (rod) and an airfoil in a tandem arrangement is investigated numerically using incompressible large eddy simulations. Quasi-periodic shedding from the rod and the resulting wake impinges on the airfoil to produce unsteady loads on the two geometries. These unsteady loads act as sources of aerodynamic sound and the sound radiates to the far-field with a dipole directivity. The airfoil is set at zero angle of attack for the simulations and the Reynolds number based on the rod diameter is Re_d = 48 K. Comparisons with experimental measurements are made for (a) mean and root mean square surface pressure on the rod, (b) profiles of mean and root mean square streamwise velocity in the rod wake, (c) velocity spectra in the near field, and (d) far-field pressure spectra. Curle’s acoustic analogy is used with the airfoil surface pressure data from the simulations to predict the far-field sound. An improved correction based on observed spanwise coherence is used to account for the difference in span lengths between the experiments and the simulations. Good agreement with data is observed for the near-field aerodynamics and the far-field sound predictions. The straight leading edge airfoil is then replaced with a test airfoil with a serrated leading edge geometry while maintaining the mean chord. This new configuration is also analyzed numerically and found to give a substantial reduction in the far-field noise spectra in the mid- to high-frequency range. Source diagnostics show that the serrations reduce unsteady loading on the airfoil, reduce coherence along the span, and increase spanwise phase variation, all of which contribute to noise reduction.

Numerical analysis of airfoil geometries inspired by the down coat of the night owl is presented. The bioinspired geometry consists of an array of 'finlet fences', which is placed near the trailing edge of the baseline (NACA 0012) airfoil. Two fences with maximum nondimensional heights, [Formula: see text] and [Formula: see text] are investigated, where [Formula: see text] is the displacement thickness at 2.9% chord upstream of the airfoil trailing edge. Wall-resolved large eddy simulations are performed at chord-based Reynolds number, [Formula: see text], flow Mach number, [Formula: see text], and angle of attack, [Formula: see text]. The simulation results show significant reductions in unsteady surface pressure and farfield radiated noise with the fences, in agreement with the measurements available in the literature. Analysis of the results reveals that the fences increase the distance between the boundary layer turbulence (source) and the airfoil trailing (scattering) edge, which is identified to be the mechanism behind high-frequency noise reduction. These reductions are larger for the taller fence as the source-scattering edge separation is greater. Two-point correlations show that the fences reduce the spanwise coherence at low frequencies for separation distances greater than a fence pitch (distance between two adjacent fences) and increase the coherence for smaller distances, the increase being higher for the taller fence. This increase in coherence and the reduced obliqueness of the leading edge of the fence are hypothesized to be responsible for the small increase in farfield noise at low frequencies observed in the simulations with the taller fence.

Computational fluid dynamics (CFD) simulations of ship airwakes are discussed in this article. CFD is used to simulate the airwakes of landing helicopter assault (LHA) and landing platform dock-17 (LPD-17) classes of ships. The focus is on capturing the massively separated flow from sharp edges of blunt bodies, while ignoring the viscous effects. A parallel, finite-volume flow solver is used with unstructured grids on full-scale ship models for the CFD calculations. Both steady-state and time-accurate results are presented for a wind speed of 15.43 m/s (30 knot) and for six different wind-over-deck angles. The article also reviews other computational and experimental ship airwake research.