This paper presents a computational study of flow incidence effects on the aeroacoustics of a propeller operating at low blade-tip Mach numbers. The numerical flow solution is obtained by using the Lattice-Boltzmann/Very Large Eddy Simulation method, while far-field noise is computed through the Ffowcs-Williams & Hawkings' acoustic analogy applied on the propeller surface. The presence of an angular inflow leads to: (i) the radiation of tonal loading noise along the propeller axis; (ii) the increment/reduction of the sound pressure level in the region from/to which the propeller is tilted away/towards. However, contrarily to propellers operating at high blade-tip Mach numbers, the noise directivity change is found to be governed only by the rise of periodic unsteady loadings, with the modulation of the strength of the noise sources on the blade, associated to the periodic variation of the observer-source relative Mach number (in the blade reference frame), being negligible. Finally, thickness noise and turbulent boundary-layer trailing-edge noise did not show a significant directivity variation due to the propeller yaw angle change.

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The accurate and reliable prediction of the aerodynamic noise sources of open rotors and ducted-fans in electric Vertical Take-Off and Landing (eVTOL) and non-conventional aircraft configurations is a challenging task from a computational perspective. Indeed, such propulsive systems can often operate in highly distorted and non-homogeneous flows, with the rotating blades interacting with strongly non-uniform and turbulent flows; and/or experience phenomena related to low Reynolds numbers and boundary-layer transition, due to the relatively small diameters and blade tip speeds. While analytical, semi-empirical and low-fidelity numerical models can provide quick and computationally inexpensive predictions, their results are often not fully reliable and their state-of-the-art requires a further development step to properly address the problem of rotor noise prediction in non-conventional aircraft and rotorcraft. On the other hand, Navier-Stokes based scale-resolving approaches such as Large Eddy Simulation (LES) have the capability to capture most of the aforementioned phenomena, but at a prohibitive computational cost for a routine employment in the design stages of such innovative vehicles. In view of this, high-fidelity scale-resolving lattice-Boltzmann (LB) numerical simulations, coupled with the Ffowcs Williams & Hawkings' (FW-H) acoustic analogy, are extensively performed and validated in this thesis. A wide range of aeroacoustic problems, spanning from airfoil and small-scale propellers in transitional boundary-layer regime to open rotors in blade-vortex interaction conditions and ducted fans ingesting the airframe turbulent boundary-layer, are addressed with the aim of predicting, identifying and characterizing and the primary sources of aerodynamic noise associated to open rotors/propellers and ducted-fans in eVTOL and novel aircraft configurations by means of the hybrid LB/FW-H approach.@en

This paper presents an experimental investigation of a propeller operating at low Reynolds numbers and provides insights into the role of aerodynamic flow features on both propeller performances and noise generation. A propeller operating at a tip Reynolds number regime of 4.3 × 10⁴ − 4.38 × 10⁴ is tested in an anechoic wind tunnel at an advance ratio ranging from 0 to 0.6. Noise is measured by means of a microphone array, while aerodynamic forces are measured with load and torque cells. Oil-flow visualizations are used to show the flow patterns on the blade surface, whereas phase-locked stereoscopic particle image velocimetry (PIV) measurements are carried out to analyze the flow at 60% of the blade radius. The pressure field around the blade section has been computed from the PIV velocity data. Results reveal a complex flowfield with the appearance of a laminar separation bubble at the suction side of the blade. The separation bubble moves toward the leading edge and reduces in size as the advance ratio decreases. At an advance ratio equal to 0.6, the flowfield is characterized by a laminar separation without reattachment. This causes vortex shedding responsible for a high-frequency hump in the far-field noise spectra.

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The goal of this paper is to perform a detailed analysis of the hydrodynamic near field and acoustic far field of a NACA 64-618 airfoil with and without serrations. The impact of serrations is investigated across different airfoil angles of attack and serration flap angles. The natural boundary-layer transition cases 6 and 7 of the AIAA Benchmark Problems for Airframe Noise Computations V Workshop Category 1 are considered as benchmark for the straight edge cases. The numerical simulations are performed using the fully explicit, transient, and compressible lattice Boltzmann equation implemented in the computational fluid dynamics/computational aero-acoustics (CFD/CAA) solver SIMULIA PowerFLOW®. The acoustic far field is obtained by using the Ffowcs-Williams and Hawkings integral solution applied to the airfoil surface. A mesh resolution study is performed on the straight trailing-edge cases to demonstrate the grid independence of the numerical solution. The numerical results compare favorably against the experimental data. The impact of the serration flap angle on the effectiveness of sawtooth serrations in reducing noise is investigated by considering three different serration flap angles. It is found that the serration flap angle primarily affects the trailing-edge noise reduction through a modification of 1) the effective angle at which the turbulent structures are convected over the serrated edge; 2) the convection velocity and spanwise coherence length along the serration; and 3) the intensity of the hydrodynamic wall-pressure fluctuations that are scattered along the serrated edge. The first and last phenomena are expected to play the most important role on the far-field noise reduction.@en

This paper proposes a CFD/CAA-based approach to predict the aerodynamic performances and tonal/broadband noise radiation of low-Reynolds number propellers at engineering level. Broadband self-noise prediction of low-Reynolds number propellers is particularly challenging, due to the requirement for the employed computational method to emulate the complexity of the laminar/turbulent boundary-layer behavior on the blade. In this study, the numerical flow solution is obtained by using the Lattice-Boltzmann/Very Large Eddy Simulation method, whereas far-field noise is computed through the Ffowcs-Williams & Hawkings' acoustic analogy applied on the propeller surface. A zig-zag transition trip on the propeller blades is used in the numerical setup to reproduce resolved turbulent pressure fluctuations in boundary-layer for broadband noise computation at a relatively low computational cost. The effect of using a transition trip to simulate low-Reynolds number propellers, as well as the impact of its chordwise position on the calculation of performances and radiated noise, is outlined. The trip position marginally affects the thrust and to a slightly larger extent the torque prediction. Tonal noise at the blade-passing frequencies does not show a relevant sensitivity to it, whereas broadband noise is found to be slightly more influenced by the chordwise position of the trip, especially at high advance ratios. The low sensitivity of the numerical results to the trip location, as well as their good agreement with loads and noise measurements carried out in the A-Tunnel of TU-Delft, demonstrates the robustness of the proposed approach for industrial applications.

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Experimental and numerical results of a propeller of 0.3 m diameter operated at 5000 RPM and axial velocity ranging from 0 to 20 m/s and advance ratio ranging from 0 to 0.8 are presented as a preliminary step towards the definition of a benchmark configuration for low Reynolds number propeller aeroacoustics. The corresponding rotational tip Mach number is 0.23 and the Reynolds number based on the blade sectional chord and flow velocity varies from about 46000 to 106000 in the operational domain and in the 30% to 100% blade radial range. Force and noise measurements carried out in a low-speed semi-anechoic wind-tunnel are compared to scale-resolved CFD and low-fidelity numerical predictions. Results identify the experimental and numerical challenges of the benchmark and the relevance of fundamental research questions related to transition and other low Reynolds number effects.

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Experimental and numerical results of a propeller of 0.3 m diameter operated in quiescent standard ambient conditions at 5000 RPM and axial velocity ranging from 0 to 20 m/s and advance ratio ranging from 0 to 0.8 are presented as a preliminary step towards the definition of a benchmark configuration for low Reynolds number propeller aeroacoustics. The corresponding rotational tip Mach number is 0.231 and the Reynolds number based on the blade sectional chord and flow velocity in the whole radial and operational domain ranges from about 54000 to 106000. Force and noise measurements carried out in a low-speed semi-anechoic wind-tunnel are compared with scale-resolved CFD and low-fidelity numerical results. Results identify the experimental and numerical challenges of the benchmark and the relevance of fundamental research questions related to transition and other low Reynolds number effects.

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Aim of this paper is to investigate the effects of the turbulent flow developing over a fuselage on fan noise for BLI embedded propulsion systems. Such configurations can suffer from inlet flow distortions and ingestion of turbulence at the fan plane with consequent impact on both broadband and tonal fan noise. The analysis is performed on a modified version of the Low-Noise NASA SDT fan-stage integrated into the ONERA NOVA fuselage in order to reproduce the NOVA BLI configuration. The numerical flow solution is obtained by solving the explicit, transient and compressible lattice-Boltzmann equation implemented in the high-fidelity CFD/CAA solver Simulia PowerFLOW®. The acoustic far-field is computed by using the Ffowcs-Williams & Hawkings integral solution applied to a permeable surface. All simulations are performed for an operating condition representative of a take-off with power cut-back. Installation effects due to the BLI configuration are quantified by comparison with an isolated configuration of the modified Low-Noise SDT fan-stage at the same operating condition. It is found that the BLI fan-stage, which is not optimal, is characterized by strong azimuthal fan blade loading unsteadiness, less axisymmetric and coherent rotor wake tangential velocity variations and higher levels of in-plane velocity fluctuations compared to the isolated engine. This resulted in no distinct tonal components and higher broadband levels in the far-field noise spectra, as well as in an increment of cumulative noise levels up to 18 EPNdB. This study, which represents the first high-fidelity CFD/CAA simulation of a full-scale aircraft geometry comprehensive of a BLI fan/OGV, provides with a clear understanding of the change of the noise sources in BLI integrated configurations.

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The scope of this paper is to assess the accuracy of the Lattice-Boltzmann/Very Large Eddy Simulation Method to predict the aerodynamics and aeroacoustics of helicopter rotors in strong Blade-Vortex Interaction conditions, and to validate a computational approach to include the effects associated to the rotor blade deflections into the numerical setup. The numerical flow solution is obtained by solving the explicit, transient and compressible Lattice-Boltzmann equation implemented in the high-fidelity CFD/CAA solver Simulia PowerFLOW^R . The acoustic far-field is computed by using the Ffwocs-Williams & Hawkings integral solution applied to a permeable surface encompassing the whole helicopter geometry. The employed benchmark configuration is the 40% geometrically and aeroelastically scaled model of a BO-105 4-bladed main rotor tested in the open-jet anechoic test section of the German-Dutch wind tunnel in the framework of the HART-II project. In the present study, only the baseline operating condition of the HART-II test, without Higher-Harmonic Control enabled, is considered. Simulations are performed either assuming a fully-rigid blade motion or a computational strategy, based on a combination of a velocity wall boundary condition applied on the blade surface and fluid body-forces fields applied in proximity of the blade leading- and trailing-edge, to partially retrieve the effects related to the experimental blade flap and torsion deformations, respectively. The impact due to the inclusion of the blade elastic deformations into the computational setup on control settings, unsteady air-loads and noise footprint predictions is outlined.

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The aim of this paper is to assess the accuracy, capabilities and computational performances of the Lattice-Boltzmann/Very Large Eddy Simulation Method to predict the unsteady aerodynamic loads, the rotor wake development and the noise radiation of helicopter rotors in strong Blade-Vortex Interaction conditions. The numerical flow solution is obtained by solving the explicit, transient and compressible Lattice-Boltzmann equation implemented in the high-fidelity CFD/CAA solver Simulia PowerFLOW ® . The acoustic far-field is computed by using the Ffwocs-Williams & Hawkings integral solution applied to a permeable surface encompassing the whole helicopter geometry. The employed benchmark configuration is the 40% geometrically and aeroelastically scaled model of a BO-105 4-bladed main rotor tested in the open-jet anechoic test section of the German-Dutch wind tunnel in the framework of the HART-II project. In the present study, only the baseline operating condition of the experimental campaign, without Higher-Harmonic Control enabled, is considered. All simulations are performed by assuming a rigid blade motion, but a computational strategy based on a combination of a rigid blade pitching motion and a transpiration velocity boundary condition applied on the blade surface is employed to take into account the blade elastic deformation motion measured during the experiments. As expected, modeling the blade elastic deformation leads to more accurate predictions of control settings, unsteady air-loads and noise footprint. The effects of the computational grid on the aerodynamic and aeroacoustic prediction is documented as well. @en

Aim of this paper is to investigate the effects of the turbulent flow developing over a fuselage on fan noise for Boundary Layer Ingestion (BLI) embedded propulsion systems. Such engine configurations can suffer from inlet flow distortions and ingestion of turbulence at the fan plane with consequent impact on both broadband and tonal fan noise. The analysis is performed by considering a modified version of the Low-Noise configuration of the NASA Source Diagnostic Test (SDT) integrated into the Nextgen ONERA Versatile Aircraft (NOVA) fuselage in order to reproduce the NOVA BLI configuration. The numerical flow solution is obtained by solving the explicit, transient and compressible lattice-Boltzmann equation implemented in the high-fidelity CFD/CAA solver Simulia PowerFLOW ^R®. The acoustic far-field is computed by using the Ffwocs-Williams & Hawkings integral solution applied to a permeable surface encompassing the fan-stage. Simulations are performed for an operating condition representative of a take-off with power cutback. Installation effects due to the BLI configuration are quantified by comparison with an isolated configuration of the modified Low-Noise SDT fan-stage geometry at same operating conditions. Comparisons are carried out in terms of fan-stage intake/interstage velocity fields, fan blades section air-loads and far-field noise; correlations between the fan-stage velocity field and noise emission for the BLI configuration are outlined. It is found that the BLI fan-stage is characterized by strong azimuthal fan blade loading unsteadiness, less periodic and coherent rotor wake tangential velocity variations and higher levels of in-plane velocity fluctuations compared to the isolated engine, resulting in far-field noise spectra with no distinct tonal components and higher broadband levels. This study represents the first high-fidelity CFD/CAA simulation of a full-scale aircraft geometry comprehensive of a BLI fan/Outlet Guide Vane (OGV) stage.

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The scope of this paper is to perform a detailed analysis of the unsteady flow properties in proximity of the trailing-edge of a lifting free transition NACA 64-618 extruded airfoil. The natural transition cases 6 and 7 of the AIAA workshop on Benchmark Problems for Airframe Noise Computations (BANC-V Category 1) are considered as references. The numerical flow solution is carried out by using the fully explicit, transient and compressible lattice-Boltzmann equation implemented in the CFD/CAA solver Exa PowerFLOWR. The acoustic far-field is obtained by using the Ffowcs-Williams and Hawking integral solution applied to the wing surface. In addition, the validity of Roger and Moreau's trailing- edge model, fed with Schlinker-Amiet's and Rozenberg's wall-pressure spectrum model, is checked by comparison with both experimental and numerical results. The influence of the computational grid on the turbulent boundary layer statistics at the trailing-edge is documented as well. As conclusive effort, the effectiveness of sawtooth serrations on noise reduction is investigated by considering two different flap angles at a fixed airfoil incidence.

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The aim of this work is to evaluate the accuracy and the computational performances of the CFD/CAA solver PowerFLOW®, developed and distributed by Exa Corporation, to predict the unsteady aerodynamic loads, the rotor wake development and the noise radiation of helicopters in Blade-Vortex Interaction conditions. The employed benchmark configuration is the 40% geometrically and aeroelastically scaled model of a BO-105 4-bladed main rotor tested in the open-jet anechoic test section of the German-Dutch wind tunnel in the framework of the HART-II project. In the present study, only the baseline operating condition of the test matrix, without higher harmonic control, is considered. All simulations are performed by assuming a rigid blade motion, but a computational strategy is employed to take into account the effective elastic deformation motion of the blade measured during the experiments. As expected, modeling the elastic blade motion leads to more accurate predictions of both unsteady air-loads and noise footprint. The effects of the mesh resolution on the aerodynamic and aeroa-coustic prediction is investigated. As a conclusive effort, the effects of fuselage scattering on the noise footprint are evaluated by using the same computational model to simulate two additional configurations: the isolated rotor of the HART-II configuration and the same rotor installed on a different helicopter fuselage. Significant far-field noise scattering effects are observed.

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