We investigate the turbulent boundary layer (TBL) over a NACA0012 airfoil at angle of attack 12 deg. A wall-resolved LES is performed for a chord based Reynolds number Re = 4 x 10(5) and freestream Mach number Re = 0.2 for which boundary layer tripping is applied near the leading edge. Mild, moderate and strong adverse pressure gradients (APGs) develop over the airfoil and, despite the strong APG, the mean flow remains attached. We observe a secondary peak in the Reynolds stress profiles arising in the outer layer and, for a strong APG, it overcomes the first peak observed in the inner layer. Due to the strong APGs on the suction side, the mean velocity profiles depict three inflexion points, the third being unstable under inviscid stability criteria. This promotes the formation of a shear layer in the outer region of the TBL which, in turn, leads to the occurrence of Kelvin-Helmholtz vortices captured by a spectral proper orthogonal decomposition (SPOD). The SPOD analysis also shows that streaks form along the airfoil suction side and, as the APG becomes stronger, they grow along the spanwise and wall-normal directions.

The effects of adiabatic and isothermal boundary conditions are investigated on the shock-boundary layer interactions (SBLIs) in a supersonic turbine cascade. Special attention is given to the characterization of the incoming boundary layers over the convex and concave walls of the blade and their impact in the SBLIs. Large eddy simulations (LES) are performed for an inlet Mach number of 𝐌_∞ = 2.0 and Reynolds number based on the axial chord 𝐑𝐞 = 200 000. For the isothermal condition, the wall to inlet temperature ratio is 𝐓_𝐰/𝐓_∞=0.75, representing a cooled wall. Different incident shock wave topologies occur on the suction and pressure sides of the airfoil. For the former, an oblique shock impinges on the boundary layer leading to a larger separation bubble. On the other hand, a normal shock from a Mach reflection induces a small separation region near the wall for the pressure side. Results are presented in terms of mean velocity and temperature contours, and the incoming boundary layers are characterized by looking at the Clauser parameter, shape factor, dilatation and turbulent kinetic energy (TKE) profiles. Inspection of the shape factors show that the adiabatic wall boundary layers are more prone to separate than the isothermal ones. This is indeed observed in the airfoil suction side, but not on the pressure side, where the flow separates in the same chord position regardless of the thermal boundary condition. This is a topic for investigation in the final version of the paper. An assessment of the dilatation and TKE profiles explains the disparities of the bubble sizes on the pressure and suction sides of the airfoil.

This study is the second in a two-part series, evaluating the ability of a lattice-Boltzmann high-fidelity simulation tool to resolve transitional flows with intermittent events. Here, the lattice-Boltzmann method (LBM) is employed to investigate the flow over a NACA0012 airfoil at an angle of attack of alpha = 3 degrees, freestream Mach number M-infinity = 0.3, and Reynolds number Re = 5 x 10(4). This flow configuration is typically found in the development of quiet air vehicles, making the study of noise generation mechanisms particularly relevant. For this case, a laminar separation bubble forms on the airfoil's suction side, leading to complex dynamics that include the shedding of coherent structures and the generation of trailing-edge tonal noise. For this second part, finite wing effects on the coherent structures are investigated and compared to the flow fields of a wing with periodic boundary conditions, as presented in the first part. The results indicate that the laminar separation bubble (LSB) on the suction side is significantly influenced by wing tip effects, forming much closer to the trailing edge than in the infinite wing. Additionally, an analysis of the vortex dynamics reveals distinct patterns of vortex shedding from the LSB, differing markedly from those observed in the spanwise homogeneous case. Overall, the finite wing impacts the pressure gradient along the wing, delaying the bubble formation and its subsequent vortex shedding and, hence, changing the acoustic noise generation.

We perform an assessment of pressure field reconstruction techniques for image velocimetry data. Three techniques are examined: a matrix omnidirectional integration method (OS-MODI), a Green's function integral approach (GFI) and a novel neural-network approach (NN). While the OS-MODI excels in noise-free conditions and structured meshes, the NN method, proposed in this study, offers superior performance in noisy environments, besides being mesh-free. For unstructured mesh analyses, challenges with ill-conditioned cells prompt sophisticated mesh generation strategies. Moreover, a simplistic application of the constant midpoint rule to compute the integrals of the boundary elements within the GFI method leads to issues with singularity. Despite obstacles, the OS-MODI and GFI show promise, with NN emerging as a robust solution for meshless noisy data.

Modal decomposition techniques are important tools for the analysis of unsteady flows and, in order to provide meaningful insights with respect to coherent structures and their characteristic frequencies, the modes must possess a robust spatial support. In this context, although widely used, methods based on singular value decomposition (SVD) may produce modes that are difficult to interpret when applied to problems dominated by intermittent and transient events. Fortunately, specific modal decomposition techniques have been recently developed to analyze such problems. However, a proper comparison between existing methods is still lacking from the literature. Therefore, this work compares two recent methods: the fast adaptive multivariate empirical mode decomposition (FA-MVEMD) and the multi-resolution dynamic mode decomposition (mrDMD). These techniques are employed here for the study of flow databases involving transient and intermittent dynamics. Specifically, the investigated problems include an SD7003 airfoil subjected to deep dynamic stall conditions, and a steady NACA0012 air-foil operating at a transitional Reynolds number. In the former case, the methods are employed to investigate the onset and evolution of the dynamic stall vortex (DSV), while in the latter case, intermittent vortex pairing is analyzed. We show that the combination of a multidimensional EMD with the Hilbert transform provides modes with superior spatial support when compared to the mrDMD, also allowing the characterization of instantaneous frequencies of coherent structures. Moreover, the EMD also condenses a larger amount of information within a single intrinsic mode function (IMF).

We investigate the effects of adverse pressure gradients (APGs) on extreme and intermittent events in the turbulent boundary layer (TBL) of a NACA0012 airfoil at 12 deg. angle of attack. A wall-resolved large-eddy simulation (LES) is performed for a Reynolds number Re = 4 x 10(5) and freestream Mach number M = 0.2. Boundary-layer tripping is enforced near the leading edge to produce bypass transition. Despite the high angle of attack, the mean flow remains attached throughout the airfoil suction side, although the adverse pressure gradient exhibits a steep rise towards the trailing edge. Results of a quadrant analysis for the Reynolds shear-stress distribution show that sweeps are predominant near the wall, while ejections dominate in the outer region. The former are the main contributor to the inner peak of turbulence production. Due to the strong APG, a secondary peak of production arises in the outer layer, and the combination of both sweeps and ejections contribute to such peak. A backflow characterization is performed, demonstrating that as the APG increases, the magnitude of the friction coefficient decreases, leading to a higher probability of such events near the trailing edge.

The dynamic stall features of a simplified vertical axis wind turbine (VAWT) configuration are analyzed. Wall-resolved large eddy simulations (LES) are performed for a NACA0018 airfoil operating as a single blade VAWT, and the operational parameters are set to tip speed ratio lambda = 3, Reynolds number Re = 50, 000 and Mach number M = 0.1. The numerical model is validated by performing a grid convergence study as well as an analysis of the airfoil spanwise domain. Results suggest that a computational domain with 40% of the airfoil chord in the spanwise direction is sufficient to capture the underlying flow dynamics when compared with span lengths of 10% and 80% of the chord. A good agreement with literature results is observed in terms of the main flow features for similar setups both in terms of integral quantities and the overall flow behavior.

Convolutional neural network (CNN) models are developed to predict the aerodynamic response from images of the flowfield of an airfoil under dynamic stall. Here, we take the aerodynamic coefficients and pressure distribution as examples. The networks are capable of identifying relevant flow features present in the images and associate them to the airfoil response, while effectively interpolating and extrapolating between flow parameters. This suggests that flow imaging may offer a promising alternative for sensors in experimental campaigns and for building robust surrogate models of complex unsteady flows.

The present study consists in the first of a two-part joint paper series, where we assess the capability of a lattice-Boltzmann high-fidelity simulation tool to resolve transitional flows with intermittent events. Here, the lattice-Boltzmann method (LBM) is employed to investigate the spanwise-periodic flow over a NACA0012 airfoil at an angle of attack of alpha = 3 degrees, freestream Mach number of M-infinity = 0.3, and Reynolds number Re = 5 x 10(4). For this particular case, a laminar separation bubble forms on the airfoil suction side, being responsible for a complex dynamics including the shedding of coherent structures that generate trailing-edge tonal noise. A grid convergence study is performed for the LBM and results are validated against large eddy simulations (LES) of the Navier-Stokes equations. Good comparisons are observed between the LBM and LES results in terms of time-averaged flow fields. Similar separation and reattachment locations are found for the suction side laminar separation bubble (LSB). An analysis in terms of the vortex dynamics is also presented, and different patterns of vortex shedding from the LSB are found, in agreement with LES results. In summary, the LBM approach is able to represent similar flow physics compared to the Navier-Stokes equations, but with around 10% of the computational cost of the LES.

Hypersonic flow over a cylinder is modeled using the finite volume method to solve the Navier-Stokes equations, including Park's two-temperature model for chemical dissociation. The main focus of this work is to carry out a comparative analysis of the thermodynamic nonequilibrium properties along the flow stagnation line, and obtain the infrared spectrum of radiative heat flux at the stagnation point of the cylinder using a line-by-line approach. The spectrum of radiative heat flux at the stagnation point of the cylinder are obtained using NEQAIR numerical code, considering the NEQAIR and HITRAN transitions. In hypersonic flow conditions, it is possible to observe the occurrence of thermodynamic nonequilibrium through the magnitude difference of the translational-rotational and vibrational-electronic temperature modes inside the shock layer forming upstream of the cylinder. Analyses of the thermodynamic nonequilibrium effects are performed, considering the excitation state of the temperature modes, in addition to the chemical effects of dissociation and exchange of molecules and atoms present in the mixture.

An investigation of streamwise vortices is performed in a supersonic turbine cascade by post-processing wall-resolved large eddy simulation data. For the present flow, the inlet Mach number is M-infinity = 2.0, and the Reynolds number is set as Re = 395, 000. Previous research highlighted the role of near-wall large-scale streaks in driving the separation bubble breathing motion for the current configuration. The present study aims to investigate how streamwise vortices affect the transport of near-wall streaks and, subsequently, the contraction of the separation bubble. For this purpose, Lagrangian coherent structures are extracted from the velocity fields at the leading and trailing edges of the bubble (prior and after the incident shock) by computing the finite-time Lyapunov exponent. As the bubble contracts, there is a noticeable increase in mixing, with the fluid being transported by the streamwise vortices from the high-speed region over the bubble, in the shear layer, towards the wall. This process occurs at specific spanwise locations, resulting in the bubble contraction or even local flow reattachment.