In this work, we investigate airfoil tonal noise generation and reduction by the means of streak generators in form of cylindrical roughness elements. Roughness elements attenuate tones in the acoustic field for the case with chord base Reynolds number Re=80,000. Further, the coupling between structures generated by surface roughness and instability modes (Kelvin-Helmholtz) of shear layer has been identified through stability analysis, suggesting stabilisation mechanisms of Kelvin-Helmholtz instabilities by which the sound generation by the airfoil is reduced by the roughness elements.

We investigate the mechanisms underlying the generation of broadband trailing-edge noise from an airfoil, focusing on the role of hydrodynamic structures with nonzero spanwise wavenumbers and the development of a corresponding reduced-order model. A wall-resolved compressible large-eddy simulation (LES) of a NACA0012 airfoil at 3 degrees angle of attack, Reynolds number Re = 200 000, and Mach number M = 0.3 is performed. The simulation replicates the experimental setup of Demange et al. [AIAA Aviation Forum (2023)], including the tripping elements ensuring turbulent boundary layer development. The numerical domain has a large spanwise extent (43.75% of the chord) to allow for the investigation of large spanwise coherent structures. Validation against experimental data shows good agreement. The simulation results show that acoustic radiation from the trailing edge follows the scattering condition, whereby only hydrodynamic waves with spanwise wavenumbers lower than the acoustic wavenumber can generate propagative sound. Notably, the contribution from nonzero spanwise wavenumber modes becomes significant beyond the cut-on frequency at which the first nonzero spanwise acoustic mode becomes propagative. Through spectral proper orthogonal decomposition (SPOD), the associated flow structures are identified as streamwise-travelling wave packets convecting obliquely across the trailing edge. Reduced-order models based on acoustic-domain extended SPOD accurately reproduce the LES sound pressure levels within 2 dB, using only two modes per frequency-wavenumber pair. This compact representation paves the way for targeted noise-reduction strategies based on the control of the identified coherent structures.

The effect of a smooth surface hump on laminar–turbulent transition over a swept wing is investigated using direct numerical simulation (DNS), and results are compared with wind tunnel measurements. When the amplitude of incoming crossflow (CF) perturbation is relatively low, transition in the reference (without hump) case occurs near 53 % chord, triggered by the breakdown of type I secondary instability. Under the same conditions, no transition is observed in the hump case within the DNS domain, which extends to 69 % chord. The analysis reveals a reversal in the CF velocity component downstream of the hump’s apex. Within this region, the structure and orientation of CF perturbations are linearly altered, particularly near the wall. These perturbations gradually recover their original state further downstream. During this recovery phase, the lift-up mechanism is weakened, reducing linear production, which stabilises the stationary CF perturbations and weakens spanwise gradients. Consequently, the neutral point of high-frequency secondary CF instability modes shifts downstream relative to the reference case, leading to laminar– turbulent transition delay in the presence of the surface hump. In contrast, when the amplitude of the incoming CF perturbation is relatively high, a pair of stationary counterrotating vortices forms downstream of the hump. These vortices locally deform the boundary layer and generate regions of elevated spanwise shear. The growth of secondary instabilities in these high-shear regions leads to a rapid advancement of transition towards the hump, in agreement with experimental observations.

This study evaluates the potential of resolvent analysis to model broadband trailing-edge (TE) noise generated by a turbulent boundary layer over an airfoil. A compressible resolvent formulation is applied to the mean flow obtained from large-eddy simulation (LES) of a NACA0012 airfoil at a chord Reynolds number of Re=200,000. The resulting modes are validated against spectral proper orthogonal decomposition of the LES data, enabling direct comparison between a physics-based model and data-driven analysis. We identify regions in the frequency-spanwise-wavenumber spectrum where the flow exhibits low-rank behavior, notably for low Helmholtz numbers and spanwise wavenumbers, coinciding with peak acoustic emissions. In these regions, the leading resolvent mode captures the dominant hydrodynamic and acoustic structures: wave packets on the suction side exploiting the Orr mechanism. The resolvent framework isolates the structures responsible for sound generation from the complete turbulent dynamics and links them to the mean flow, offering a physics-based model suited for noise control. Unlike empirical models requiring full turbulent spectra or high-fidelity simulations with prohibitive cost, the resolvent approach provides a low-order, efficient alternative. These results demonstrate its potential as a foundation for future sensitivity-based design strategies targeting broadband TE noise reduction.

Surface roughness, such as forward-facing steps (FFS), significantly impacts the laminar-turbulent transition in swept-wing boundary layers. This study employs direct numerical simulations (DNS) to investigate the non-monotonic relationship between FFS height and transition location, complementing recent experimental findings. Two step heights and a baseline clean surface are analyzed under experimentally representative conditions. The simulations confirm a delay in transition for the shallow FFS and premature transition for the higher FFS, reproducing experimental trends while revealing some discrepancies in the abruptness of transition advancement for the larger step. DNS results provide detailed insights into the interaction of crossflow modes with the FFS, particularly near the step, highlighting the spanwise modulation and the formation of reverse flow regions.

This study examines the transition to turbulence downstream of fluttering and non-fluttering bioprosthetic aortic valves using global linear stability theory. During systole, increasing inflow velocities result in temporally evolving flow profiles downstream of the valve which are highly influenced by the leaflet kinematics. These profiles are time averaged at the sinotubular junction over successive windows and used as boundary conditions to obtain base flows for stability analysis. Three-dimensional global modes are computed for one design of each valve type across multiple time windows, revealing several unstable modes whose frequencies and growth rates increase over time. Notably, the non-fluttering valve exhibits higher growth rates than the fluttering valve. The resulting eigenspectra show that, for each case, the most unstable eigenvalues align along two distinct parabolic branches in the complex plane. For each valve case, the modes within each branch are found to have similar group velocities, suggesting that the unstable modes along a branch constitute a coherent structure. Motivated by this, a transient growth analysis is conducted to identify the optimal initial perturbations that maximise energy gain for a given time horizon. When superimposed onto the base flow, these perturbations generate vortical structures that closely resemble those observed in fully coupled nonlinear fluid–structure interaction simulations for a similar time scale as the one used to obtain the optimal perturbations. These results suggest that the optimal perturbations may initiate the shear-layer instabilities responsible for transition to turbulence, providing valuable insight into the underlying mechanisms in the flow fields downstream of bioprosthetic valve designs.

This work examines how inviscid interface-flux choices and Local Discontinuous Galerkin (LDG) viscous-flux parameters affect solution sensitivity and computational cost in high-order Discontinuous Galerkin Flux Reconstruction (DG-FR) simulations of compressible turbomachinery flows. A transonic low-pressure turbine cascade at M2is=1.22 and Re2is=1.5×105 is considered using a range of inviscid Riemann solvers, including Rusanov, HLLC-type solvers, RoeM and an Exact solver. The LDG discretization is varied through the penalty parameter τ and the directional parameter β. In addition to standard formulations, where the same inviscid flux is used on all interfaces, hybrid HF-R-X formulations are investigated, with Rusanov applied at wall boundaries and solver X applied on interior interfaces. The comparison is based on integrated aerodynamic forces, surface distributions of isentropic Mach number and skin friction, loss coefficients evaluated from inlet and outlet sampling planes, base pressure and normalized wall-clock time. The results show that, for the operating condition and resolution considered, lift and drag are only weakly sensitive to the inviscid Riemann solver, with variations remaining below approximately half a percent across the converged cases. Surface pressure and skin-friction trends are also nearly unchanged among the approximate solvers. Loss-related quantities are more sensitive to the numerical formulation and provide a clearer distinction between inviscid-flux and LDG-parameter choices. From a performance perspective, approximate inviscid solvers introduce only marginal overhead relative to the Rusanov baseline, whereas the Exact solver increases the wall-clock time more noticeably. The hybrid HF-R-E formulation reduces part of this cost penalty by avoiding Exact flux evaluations at wall boundaries. The dominant runtime sensitivity is associated with the LDG directional parameter: centered viscous paths with β=0 produce a substantial cost increase compared with biased paths with β=±0.5. Overall, the results indicate that inexpensive approximate inviscid fluxes are sufficient for integrated load prediction in this case, while loss prediction and computational efficiency are more strongly affected by the LDG viscous-flux construction and by the use of hybrid flux strategies.

In a combined experimental and numerical effort, we investigate the generation and reduction of airfoil tonal noise. The means of noise control are streak generators in the form of cylindrical roughness elements. These elements are placed periodically along the span of the airfoil at the mid-chord streamwise position. Experiments are performed for a wide range of Reynolds numbers and angles of attack in a companion work (Alva et al., AIAA Aviation Forum, 2023). In the present work, we concentrate on numerical investigations for a further investigation of selected cases. We have performed wall-resolved large-eddy simulations for a NACA 0012 airfoil at zero angle of attack and Mach 0.3. Two Reynolds numbers (0.8 × 10⁵ and 1.0 × 10⁵) have been investigated, showing acoustic results consistent with experiments at the same Reynolds but lower Mach numbers. Roughness elements attenuate tones in the acoustic field and, for the higher Reynolds number, suppress them. Through Fourier decomposition and spectral proper orthogonal decomposition analysis of streamwise velocity data, dominating structures have been identified. Further, the coupling between the structures generated by the surface roughness and the instability modes (Kelvin–Helmholtz) of the shear layer has been identified through stability analysis, suggesting stabilisation mechanisms by which the sound generation by the airfoil is reduced by the roughness elements.

The Actuator Line Method (ALM) is a technique that replaces the detailed airfoil geometry with distributed body forces to predict the flow field. ALM has been widely employed for simulating rotating blade wakes due to its flexibility and accuracy in the far field. In this study, the applicability of ALM for unsteady aerodynamics and acoustic field prediction is evaluated. The case study considered is the harmonic transverse oscillation of a thin airfoil in uniform flow. The ALM body forces are distributed over a few grid points following a Gaussian function, with a range of smearing ratio of epsilon/c. (smearing parameter over the chord length) between 0.4 and 1. These forces are computed using thin airfoil theory with the Prandtl-Glauert correction for compressible regime. Based on these computations, the compressible Navier-Stokes equations are numerically solved, yielding the velocity and pressure fields. ALM lift results are validated against unsteady aerodynamic theory in the subsonic regime. Moreover, results demonstrate an acoustic field consistent with a dipole distribution and a spectrum exhibiting a frequency corresponding to the plunging motion. Furthermore, the acoustic results are validated through an acoustic analogy approach, involving the prediction of the acoustic field via Green's function. The prediction of the acoustic far-field using ALM is expected to significantly reduce the computational cost of compressible simulations applied to propeller and wind turbine aeroacoustics.

In this work, Direct Numerical Simulation is performed on a low-pressure turbine blade with parallel end-walls, in a linear cascade environment at an exit Reynolds number of 1.5 · 10⁵. Our simulations are performed with Neko, a framework for high-order spectral elements for heterogeneous computing architectures. Secondary flow structures and associated losses are presented in configurations with and without free-stream turbulence and with a Blasius boundary layer inflow profile. Instantaneous and mean flow visualizations validate the classical secondary flow structures reported in the literature. The results highlight strong vortex cores at the outflow and large contributions to losses from the passage vortex and trailing shed vortex (or counter vortex). The application of turbulent structures at the inflow does not affect the formation of the horseshoe vortex nor the vortex cores at the outlet, but still suppresses the shedding at midspan. Proper Orthogonal Decomposition (POD) is applied to provide an overall picture of the flow structures in the entire domain. Without free-stream turbulence, the most energetic modes are found to be linked to the shedding at mid span and the secondary flow structures. Fourier analysis of the POD times series show low frequencies associated with the secondary structures. POD modes for the simulation with free-stream turbulence shows identical secondary flow structures, with additional streamwise-elongated streaky structures in the blade boundary layer and without any modes related to shedding.