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This study deals with the development of a numerical wire mesh screen model for the unsteady simulation of aeroacoustic configurations. Such devices are indeed of growing interest for practical app...

International audience; A Zonal Detached Eddy Simulation (ZDES) has been performed on the simplified LAGOON nose landing gear geometry using a Navier-Stokes solver on a fully unstructured grid. The attached boundary layers have been finely resolved using Y + values in the order of unity, while the high curvature zones have been intensively meshed in order to accurately solving adverse pressure gradients present in these regions. The mean and fluctuating flow fields have been compared with the experimental results, proving that both the mean flow field and the spectral content recorded at the wall are accurately reproduced. Following these comparisons, a detailed analysis of the topology of the flow has been carried out through the analysis of the skin friction coefficient and friction lines, coupled with three dimensional visualizations of the landing gear wake. The far-field acoustics, computed through the Ffowcs-Williams and Hawkings (FW-H) equation from the computed pressure on the landing gear skin, has been compared with the experimental results, obtaining a very good agreement for the different microphones and directions. Finally, the CFD methodology presented in this study proves to be a moderate cost approach, enabling an accurate flow and noise prediction for bluff bodies such as landing gears.

This study presents hybrid Reynolds-averaged Navier–Stokes/large-eddy simulations of the unsteady flow and noise-generation phenomena in the slat cove of a high-lift wing profile. These computations are part of a joint numerical/experimental aeroacoustics collaborative program dedicated to slat-flow analysis. A dedicated two-element wing profile (slat plus main body) has been designed to isolate slat noise from other possible sources (e.g., the flap), while minimizing mean flow deflection effects, to improve the fidelity of open-jet wind-tunnel measurements. The design of this two-element airfoil has been performed numerically, using an optimization process based on steady Reynolds-averaged Navier–Stokes calculations. This airfoil has been investigated experimentally at the Ecole Centrale de Lyon open jet facility. Unsteady zonal hybrid Reynolds-averaged Navier–Stokes/large-eddy simulations have been performed to provide a comprehensive description of the unsteady flow inside the slat cove, focusing on th...

International audience; In the SWAHILI (SWept Airfoil with HIgh LIft) Onera-DLR Common Research Project (CRP), both research centers are building an experimental database for the validation of CFD/CAA codes applied to the simulation of the unsteady flow and noise generation from a high-lift profile with deployed slat and flap and a global sweep angle of 30°. The project is based on DLR’s model F16, a two-dimensional airfoil (constant section in the span direction), with a 300 mm clean (retracted) chord. This project is a continuation of the LEISA2 (Silent Take-Off and Landing, 2010-2013) project, in which a similar database has been built with the same model, but without sweep angle. Since 2014, the LEISA2 test-case has been included by NASA and AIAA in the Benchmark for Airframe Noise Computations (BANC). In both projects LEISA2 and SWAHILI, the model has been tested in F2, an aerodynamic wind tunnel located in Onera-Le Fauga, with intensive aerodynamic measurements, including steady/unsteady wall pressure sensors, optical devices such as PIV and steady/unsteady LDV and a hot wire probe. Acoustic measurements were also achieved in F2 with a wall microphone array mounted in the windtunnel ceiling. During the SWAHILI project, Dassault-Aviation has partnered with Onera and DLR, contributing to the project with two additional configurations of flap side edge, with the same aerodynamic and acoustic measurements in F2. In LEISA2 the un-swept model has been also tested in AWB, an anechoic open-jet wind tunnel located in DLR-Braunschweig for acoustic data acquisition. In the SWAHILI context, similar tests are under progress with the SWAHILI swept model. The present paper focusses on the influence of the sweep angle and the flap side edge on the characteristics of the unsteady flow and on the radiated noise.

This paper presents numerical flow predictions of a two-wheel landing gear. Computations are based on the Lattice Boltzmann Method implemented in LaBS, a solver developed by a French consortium led by industrial companies and academic institutes. The landing gear geometry is one of the LEG (“Landing gear noise Evaluation Geometry ”) landing gear model configurations designed and tested by JAXA. This model includes geometric details such as the wheel-caps with cooling holes, brake calipers, and torque link. The unsteady flow simulations are achieved on “octree” Cartesian grids using cubic cells with refinement doubled along surfaces assigned by the user. Solid surfaces are treated via an Immersed Boundary Method. Two computations are currently achieved on grids with coarse and medium resolutions, and a final computation on a fine grid is still running. The time-averaged wall pressure on the wheels and velocity components in the landing gear wake show a global agreement with the steady pressure taps and PIV respectively. The main discrepancies appears in the flow deviation because the boundary conditions of the flowdiffers from open-jet condition in CFD to closed-test section in wind-tunnel. The medium grid compuations globally improves the time-averaged and unsteady aerodynamic predictions regarding the coarse grid results. Given their accuracy, these results already confirms the relevancy of the lattice-Boltzmann method for the landing-gear flow simulation., Cet article présente les prévisions numériques du bruit d'un train d'atterrissage à deux roues.Les calculs sont basés sur la méthode Lattice Boltzmann implémentée dans LaBS, un solveur développé par un consortium français dirigé par des entreprises industrielles et des instituts universitaires. La géométrie du train d'atterrissage LEG est un modèle de train d'atterrissage conçu et testé par la JAXA. Ce modèle comprend des détails géométriques tels que les jantes des roue avec des trous de refroidissement, des éléments de freinage et un compas. Les simulations d'écoulement instationnaire sont réalisées sur des grilles cartésiennes dites "octree" utilisant des cellules cubiques avec un raffinement doublé le long des surfaces assignées par l'utilisateur. Les surfaces solides sont traitées par une méthode des frontières immergées. Deux calculs sont actuellement effectués sur des résolutions grossières et moyennes, et un calcul final sur une grille fine est toujours en cours. La pression moyenne sur les roues et les composantes de vitesse dans le sillage du train d'atterrissage montrent un accord global avec les mesures réalisées en soufflerie. Ces résultats, bien que limité à un maillage de taille moyenne, confirme l'intérêt de la méthode LBM pour la simulation d'écoulement de train d'atterrissage.

The Benchmark for Airframe Noise Computations (BANC) has been initiated by the BECAN Technical Discussion Group under AIAA. This continuous framework, mainly impulsed by NASA LaRC, aims at evaluating numerical methods for the simulation of unsteady flows and aerodynamic noise radiated by airframe components such as airfoil trailing edges, landing gears and high lift devices. In this context, eight test-cases are proposed, with problem statements relying on extended experimental databases. The 2-wheel LAGOON landing gear is one of them. Originally designed and tested in the homonym project funded by Airbus-France, it has a simplified shape, compatible with a wide range of numerical methods, although involving complex physics. This paper summarizes seven submissions that were presented in this category at the Third BANC Workshop in Atlanta in June 2014. Researchers employed various block-structured, unstructured and embedded Cartesian (“octree”) grids and large computational resources to simulate the flow and radiated noise. The solutions are compared against each other and with experimental data gathered in two Onera’s windtunnels, F2 for aerodynamic data and CEPRA19 for acoustic data. Overall, all simulations captured the main features of the unsteady flow and radiated noise, including cavity resonances occurring between the wheels. The acoustic spectra and directivity diagrams of overall sound pressure levels are in fair agreement with experiment, although a significant dispersion can be observed between all contributions.

This study presents several unsteady computations of the 30P30N three-element high-lift airfoil, focused on slat noise prediction. These simulations rely on several RANS/LES approaches: two global hybrid approaches, DDES (Delayed Detached Eddy Simulation) and ZDES (Zonal Detached Eddy Simulation) and one zonal approach, the NLDE (Non-Linear Disturbance Equations). Global approaches compute all the geometry, while zonal simulations only consider an LES domain clustered around the slat to save computational resources. Two grid resolutions are considered, in order to perform a critical comparison of both the modeling and grid resolution effects for this kind of configuration. The obtained results are favorably compared to near-field results measurements obtained in JAXA low-speed Wind Tunnel (JAXA-LWT2): narrow band peaks present in the pressure spectra are well recovered by all the simulations, with a very good agreement with the measurements for the fine grid results, while the medium grid simulations tend to over-estimate the magnitude of these peaks. It is also observed that the DDES approach seems to delay the development of instabilities in the slat cusp shear layer, while other approaches lead to a faster break-up of the shear layer. The zonal NLDE approach leads to significant CPU time savings due to the reduced number of grid points, as well as to shorter transient times. However, such simulations do not account for the whole diffraction and reflection effects, as observed on the far-field acoustic field.

In the LEISA2 project (Silent Take-Off and Landing), ONERA and DLR have collaborated on an common program which main motivation was to build an experimental database for the validation of numerical codes devoted to the simulation of unsteady flows and noise generation by a high-lift wing configuration, data which are globally missing to the aeroacoustic community. Such database may also help explaining the physical mechanism of high lift airfoil noise generation and answer the general need for resolving uncertainties in high lift noise testing and code validation. The activity is based on testing the existing DLR model F16 in two test campaigns. In a first step, the model has been adapted by DLR to F2 aerodynamic windtunnel located in Onera-Le Fauga and intensive aerodynamic measurements were achieved using optical devices such as PIV and LDV, but also acoustic measurements using a wall microphone array. Then acoustic tests were conducted in AWB, an anechoic open-jet windtunnel located in DLR-Braunschweig. The present paper gives a detailed description of this database, with several comparisons of data measured in both windtunnels. The formatted and documented database is now available to the worldwide aeroacoustic community in the framework of the Benchmark for Airframe Noise Computations (Category 6) organized by NASA LaRC and sponsored by AIAA.

This paper presents a theoretical analysis and two applications of a high-order overlapping grid method for coupling Cartesian and curvilinear grids, developed in order to simulate aerodynamic noise. First, the overlapping grid method based on Lagrange interpolating polynomials is described and a theoretical analysis of the interpolation operator is then carried out. It shows that the interpolation generates spurious modes that depend on the wavenumbers of the signal. Besides it also gives the optimal conditions in which interpolation can be applied. Then an application to the simulation of the aeroacoustic noise generated by the vortex shedding behind a cylinder is presented. During this simulation, it appears that interpolation can create some spurious acoustic modes in regions where hydrodynamic fluctuations are significant, as predicted by the theoretical analysis. It is shown that these spurious modes disappear when a refined Cartesian grid is used (26 points per wavelength of the vortex shedding were found to be adequate in this study). At last, the simulation of the aerodynamic noise of a three element high-lift wing profile has then been carried out. For this application, the main acoustic source at the slat trailing edge is represented analytically. The propagation of the generated acoustic wave is simulated with a mean flow at rest and with a steady turbulent mean flow computed by RANS. The first application allows us to assess the method by comparing the results to a reference solution. The second one shows that the influence of a non-uniform mean flow on the directivity of an acoustic source can be observed in complex geometries. This application therefore shows that the proposed coupling method is well adapted to complex geometries that are usually met in industrial applications.

This paper describes some significant steps made towards the numerical simulation of the noise radiated by the high-lift devices of a plane. Since the full numerical simulation of such configuration is still out of reach for present supercomputers, some hybrid strategies have been developed to reduce the overall cost of such simulations. The proposed strategy relies on the coupling of an unsteady nearfield CFD with an acoustic propagation solver based on the resolution of the Euler equations for midfield propagation in an inhomogeneous field, and the use of an integral solver for farfield acoustic predictions.

This study presents a highly-resolved Large Eddy Simulation of a three-element highlift airfoil with deployed slat and flap, which involves a large amount of CPU resources (2.6 billions of grid cells and 6 millions of CPU hours). The aim of this simulation is double: the first objective is to perform a detailed flow analysis and identify the physical events responsible for noise generation as well as the implied flow mechanisms; the second objective is to provide an important numerical database that may be used in the future to assess more affordable numerical approaches based for instance on hybrid RANS/LES methods. The assessment of the present LES relies on a recent experimental campaign performed in Onera F2 wind tunnel in the framework of a joint Onera/DLR project named LEISA2.

The filtered Navier-Stokes equations are used to perform the large-eddy simulation of the unsteady incompressible flow around the blunt trailing edge of a thick flat plate. The computed flow exhibits a three-dimensional vortex shedding mechanism. The frequency domain of this mechanism is in agreement with experiments and theory. In that frequency domain normalized wall-pressure levels are favorably compared to spectra measured at the blunted trailing edge of an airfoil. The far-field radiated noise is first computed via Curle's formulation, the solution of Lighthill's equation for flows embedding solid bodies. Then the theory developed by Ffowcs Williams and Hall is considered. This formulation expresses the noise generated by turbulence passing over the edge of an infinite half-plane. The suitability of this theory to the case for a thick plate is discussed. The normalized spectra of the radiated noise predicted by both methods are compared, in the frequency domain of the vortex-shedding mechanism, to airfoil noise measurements in an anechoic facility.

The LAGOON program currently supported by Airbus Operation SAS and involving ONERA and DLR (the French and German national aerospace research centres) and Southampton University, focus on the validation of aeroacoustics numerical tools for landing gear applications. An experimental aeroacoustic program have been performed in ONERA’s aerodynamic and anechoic windtunnels F2 and CEPRA19 with a generic landing gear configuration. The present work follows the previous CFD/CAA coupling of the baseline landing gear model configuration at Mach number M = 0.23. Using the same computational process but run at a Mach number M = 0.18, CFD is achieved with Zonal-Detached Eddy Simulation (ZDES) using ONERA’s code elsA and provides the data for the FfowcsWilliams and Hawkings surface integral methods (Onera’s code KIM). Both solid and porous surfaces are used and compared to experiment. In addition, the errors probably induced by the acoustic reflections on the CFD floor and the wake impinging the porous surfaces are studied. These errors are found to be minimal in the upstream direction where the porous surface results are in best agreement with experiment. In all directions, the solid surface results are in good agreement with experiment, predicting an overall sound pressure level with an error lower than 1.5 dB.

In the context of the evaluation and reduction of fan noise in modern turbofan engines, the possibility to compute the acoustic transmission through a rotor-stator system, accounting for the rotation of the rotor, is evaluated. The objective is the development of a numerical method to simulate the propagation of acoustic waves in a global fixed region which includes a partial moving zone, with equations solved in a fixed reference frame. In order to insure a high order resolution these CAA developments and computations were performed using Onera’s CAA solver sAbrinA-V0. Firstly, a full or partial 2D moving zone was employed as a reference test case. It was insured that the moving grid acts only as a resolution support and does not affect the propagation of the waves by its rotation. Monopole and planar sources were considered. In order to connect the two (fixed and rotating) zones, spline cubic interpolation functions were employed in the ghost cells rows. Secondly, a cross-shaped rigid obstacle with zero thickness walls was embedded in the moving grid and the interaction between this obstacle and a fixed monopole was computed. Note that, despite the CAA solver can handle any non-homogeneous mean flow, this case was addressed accounting for a medium at rest. The field of application of the method is restricted to non-viscous interactions between boundary layer and acoustics, so the considered mean flows have to respect this condition. Finally, a 3D test case is presented, based on a simplified rotor with four twisted blades with zero thickness in a perfect annular duct. The acoustic interaction of this rotating blades with an helicoidal modal source injected upstream the rotor within a medium at rest was evaluated. Recent developments in sAbrinA-v0 on improved inlet/outlet acoustic boundary conditions and conclusions of gust-fixed airfoil interactions will allow a proper evaluation of CAA simulations of moving surfaces.

This study presents numerical simulations of the unsteady flow and noise generation phenomena in the slat cove of a highlift wing profile by mean of a zonal hybrid RANS/LES approach. These computations are part of a joint numerical/experimental aeroacoustics collaborative program, dedicated to slat flow analysis. For that purpose, a dedicated twoelement wing profile (slat+main body) has been designed in order to isolate the slat noise sources from other possible spurious sources (flap for instance), while minimizing mean flow deflection effects in order to improve the reliability of open-jet wind tunnel measurements. The design of this two-element airfoil has been done numerically using an optimization process based on steady RANS calculations. This airfoil has been investigated experimentally in Ecole Centrale de Lyon open jet facility, therefore providing unsteady wall-pressure measurements in the slat cove area and farfield acoustic measurements. Unsteady zonal hybrid RANS/LES simulations have been performed to provide a comprehensive description of the unsteady flow inside the slat cove, with focus being made on the noise generation processes. A fine physical unsteady analysis of the flow inside the slat cove is presented, as well as a comparison of numerical results with available experimental ones. The pressure spectra associated to the slat cove flow appear to be characterized by several tonal peaks emerging from a global broadband content. The existence of such peaks is discussed and attributed to a feedback loop involving the main shear layer inside the slat cove. A theoretical law is proposed and assessed at the end of the paper to predict the associated tonal frequencies.

In the context of the on-going European research project OPENAIR, the acoustic potential of the concept of scarfed nozzles for coaxial turbofan is evaluated. The objective is to reduce the fan noise radiation through the engine nozzle towards the ground, without significant losses in the aerodynamic performances. This evaluation relies on CAA computations achieved with Onera’s CAA solver sAbrinA-V0 . The nozzle configuration is typical of a coaxial turbofan with large bypass ratio, including 3D effects from the internal bifurcation and external pylon. From a reference configuration, a scarfed configuration has been designed by SNECMA with the constraint of minimizing the loss in aerodynamic performance. The acoustic computations rely on the Random Phase Multimodal Injection (RPMI), an innovative technique based on the optimization of the modal phases in order to obtain, with a minimum number of CAA computations, the contribution, in an un-correlated way, of all cut-on modes with evenly distributed acoustic power. The noise propagation accounts for (i) the 3D geometric details of the secondary duct, including the bifurcation and (ii) the refraction effects due to the large velocity gradients in the coaxial nozzle. For this purpose, non-homogeneous RANS mean flows were computed by SNECMA for both the reference and the scarfed configurations, allowing to check their respective aerodynamic performances. The CAA computations provide acoustic fields on a surface enclosing the engine and pylon, then the farfield directivity of the isolated engine is evaluated using a Kirchhoff integral method. Both configurations are compared at large distance in flyover and sideline directions corresponding to certification points. The acoustical benefit of the scarfed nozzle is demonstrated, especially in the flyover direction. Abbreviations

Landing gear noise computations require the calculation of unsteady flow fields around complex geometries. Recent experience was gained at Onera in CFD computations on LG simplified geometries, based on an aerodynamics solver which uses block structured grids ( els A). However, simulations of more complicated configurations are now also envisaged, and there are questions tha t a structured approach may not succeed in handling complex geometries, without requiring excessive grid works. The present study describes first attempts to assess th e DDES (Delayed Detached Eddy Simulation) method of Onera’s unstructured flow solver CEDRE, with respect to the simulation of landing gear aeroacoustics. This work is part of an Onera internal effort that is conducted in the framework of the Benchmark for Airframe Noise Computations (BANC). The paper firstly presents results obtained on the supercritical tandem cylinders configuration, an academic test case prop osed in the BANC framework. As a reference, computations are first run with a classi cal structured grid, which is easy to construct for such a configuration. This computatio n is followed by a calculation performed on an unstructured grid, which objective is to assess the sensitivity of the results to unstructured grids. For both calculation s, the farfield noise is then extrapolated from the local CFD results, using an i ntegral FW-H (Ffowcs WilliamsHawkings) method. Finally, the solver CEDRE is applied to a more complex test case representing a 1/4 th model of a partially dressed nose landing gear, al so proposed in the

rance This paper presents results of a CFD-CAA coupling approach for the prediction of the aerodynamic noise radiated by a generic landing gear model. These results are compared to experimental aerodynamic/aeroacoustic data gathered in Onera’s aerodynamic and anechoic windtunnels F2 and CEPRA19. This work is achieved in the framework of the LAGOON program ( LA nding Gear NOise database for CAA validatiON ), currently supported by Airbus and involving Onera and DLR (the French and German national aerospace research centres) and Southampton University. In the present work, the simulation of the unsteady flow is achieved with Zonal-DES using Onera’s code elsA . The CFD provides the surface pressure fluctuations which are the entry d ata for a Ffowcs-Williams and Hawkings (FW-H) surface integral solver used to predict the far field noise. Two FW-H formulations are used and compared, based either on the model rigid surface or from a porous surface surrounding the landing gear. Numerical unsteady aerodynamics are presented firstly and compared with experiment results, then acoustic predictions are detailed and compared with measurements from microphones. The noise predicted from the solid surface input is in accurate agreement with the experiment. The acoustic prediction from the porous surface input globally overestimates the experimental noise levels. In general, the agreement with experimental acoustic data depends on the sound direction, upstream or downstream and flyover or sideline. Tonal noise peaks, which may b e attributed to cylindrical cavities located on the inner side of wheels, are well predicted by the simulations.

The main objective of this paper is the characterization of the fan noise radiated in the aft direction for a coaxial turbofan installed in RFN (Rear Fuselage Nacelle) position. The challenge is to validate a numerical hybrid methodology, which has been developed for several years by Onera and Airbus, to provide the acoustic far field in the presence of the aircraft. This activity refers to the NACRE European project, in which a dedicated isolated/installed fan noise experiment was conducted with an aircraft model and a Turbine Powered Simulator (TPS). The numerical approach combines near field CAA computations, achieved by Onera with the sAbrinA-V0 solver, and farfield BEM computations achieved either by Airbus with the ACTIPOLE solver or by Onera with the BEMUSE solver. The fan noise source consists in azimuthal/radial modes of an infinite annular duct with uniform mean flow. All “cut-on” modes are computed individually, and then the un-correlated mode contributions are summed with amplitudes derived from measurements performed inside the by-pass duct with a circular array of pressure sensors. It is shown that the simulation of the internal propagation must take into account the three-dimensional geometry of the bypass duct, including the bifurcation, to explain the strong azimuthal oscillations of the RMS pressure inside and outside the TPS. Onera and Airbus’s BEM solvers are used to compute the installation effects generated by the rear part of the aircraft equipped with an H-tail plane. The agreement between both solvers is excellent, and the comparison with the experiments is satisfying.

This paper presents early results of an experimental aeroacoustic program recently performed in ONERA’s open-jet anechoic windtunnel CEPRA19 with a generic landing gear configuration. This program is the continuation of steady/unsteady flow measurements achieved in 2007 in ONERA’s F2 aerodynamic windtunnel (closed test section). Both experiments constitute the experimental phase of the LAGOON program (LA nding Gear NOise database for CAA validatiON ), currently supported by Airbus and involving ONERA and DLR (the French and German national aerospace research centres) and Southampton University, with the general purpose of evaluate up-to-date CFD/CAA techniques for airframe noise simulation, validated against an extensive experimental aerodynamic/acoustic database. The first point addressed in the present paper is the flow identification between both facilities, relying on limited aerodynamic measurements performed with a 5-hole probe and with onboard static pressure taps and unsteady pressure transducers. This flow identification was supported by a CFD study of the 3D flows in both windtunnels, performed by DLR. It is shown that there exist tiny differences between both flows. The second addressed point is the “signal-to-noise” ratio, or the ratio of the aerodynamic noise radiated by the model, to the background noise measured in the windtunnel without the model. It is shown that this ratio is globally satisfying, especially since this landing gear model with smooth shape is expected to be more silent than an actual landing gear with the same scale.

The EC research program NACRE (New Aircraft Concepts Research in Europe) studies innovative aircraft concepts, including radical integration of the engines. This paper describes numerical simulations of fan noise installation effect for a turbofan engine with a semi-buried inlet. The main objective is to evaluate the influence of the “offset level” (the vertical distance between the airframe surface and the lowest position of the fan, divided by its diameter). Two different configurations with offset levels of 8 % (Shape 1) and 15 % (Shape 2) are studied. 3D acoustic computations are performed using ONERA's CAA sAbrinA solver based on the non-linear Euler equations in perturbation form, using specific structured multi-block acoustic grids. The entry data are the mean flow RANS simulations computed by ONERA. Acoustic modes (cut-on) generated in the engine fan plane are prescribed by MTU (frequency, azimuthal and radial orders). These acoustic computations account for the acoustic refraction effect through the inlet non-homogeneous mean flow. First results are presented at the BPF for one single particular mode ( m = 0, n = 4), then additional “ broadband ” simulations are achieved for both shapes by combining all prescribed cut-on modes with equal energy. All computations show that noise levels radiated by Shape 2 are globally lower than the noise radiated by Shape 1. However, this difference could be attenuated by an exact estimation of the amplitudes considered for the modes generated by the inflow distorsion.

In the context of the AEROCAV French Research Program, Large-Eddy Simulations were conducted to compute the turbulent structures and the noise generated by the flow over realistic cavities, similar to those present on air craft fuselages and wings. The radiated acoustic field was computed using a Kirchhoff integ ral method. Two geometrical configurations are considered, both with cylindrical shapes but with different Depth/Diameter aspect ratio (H/D = 1 and H/D = 0.5 ) embedded in a subsonic flow ( Mach = 0.2 ) and for a Reynolds number based on the length of the cavity equal to 4.6·10 5 . Dedicated experimental aerodynamic and aeroacoustic measurements databases are used to validate the numerical computations. Experimental and numerical flows are similar except that the boundary layer upstream the cavity is turbulent in the experiments and laminar for the CFD. The aerodynamic fields in the two sets of results are in good agreement, whereas the acoustic field is largely influenced by the install ation effects. These effects are analyzed through computations based on Boundary Element Method and CAA and then corrections factors are provided. The acoustic emission for the H/D = 1 case is characterized by a discrete tonal mode, as a superposition of structur al and Rossiter resonance phenomena, while the H/D = 0.5 case presents a directive and highly dissymmetrica l broadband noise emission. A numerical process is also proposed to c ontrol this dissymmetry.

The present work is devoted to the numerical simulation of acoustic emissions characterizing turbojet engines, a subject that is relevant of the more general purpose of aircraft noise prevision and reduction. More precis ely, we explore here the ability of a structured CAA (Computational AeroAcoustics) method/solver to address complicated problems of engine noise prediction. With that end, and by using the ONERA’s CAA solver sAbrinA.v0 , we conduct realistic calculations of aft fan nois e emission, which involve both a full-3D exhaust geometry (with its pylon / internal bifurcations) and typical fan noise modal contents (high azimuthal order / frequency). The re sults highlight how far the installation / refraction effects induced by the complex geometry / flow of an engine can affect its fan noise emission. Results also tend to demonstrate that bot h the here used CAA method and solver are mature enough to face out industrial-like engin e noise problems.

Soucieux de remedier a l'impact societal du trafic aerien, l'ensemble des acteurs du monde aeronautique s'efforce d'ameliorer toujours plus la discretion acoustique des avions commerciaux. Dans ce contexte, l'ONERA s'est employe a acquerir puis a accroitre ses capacites en termes de simulation numerique en aeroacoustique. Recemment, dans le cadre d'une collaboration avec Airbus, ces moyens ont permis d'investiguer numeriquement le probleme, complexe, de la caracterisation du bruit aval de soufflante.

In the framework of airframe noise, computational aeroacoustic methods give results which have to be compared with experimental measurements. Such a validation process requires to perform aerodynamic measurements in high quality flows, and acoustic measurements in an quiet and anechoic environment. The French Aerospace Lab (ONERA) launched the EXAVAC experimental programme in order to constitute a complete aeroacoustic database. Measurements were carried out in two ONERA’s test facilities : the F2 aerodynamic wind tunnel and the CEPRA19 acoustic wind tunnel. The test model is a NACA0012 airfoil with a blunted trailing edge. This paper describes the experimental programme and its principal results. The growing interest for aerodynamic noise generated by civil aircrafts in approach, primarly generated by high-lift systems and landing gears, strongly stimulates the development of Computational Fluid Dynamics (CFD) methods and Computational AeroAcoustic (CAA) methods. Whereas CFD methods aim at reproducing unsteady flows around airfoils, CAA methods are intended to simulate the acoustic radiation of aeroacoustic sources through mean flow, which is non-uniform near airfoils and uniform in the far field. Today, these numerical methods give results which have to be compared with experimental measurements. Therefore, the need for fundamental experiments intended to a validation process of CAA methods is critical, both for aerodynamicists and acousticians. Such a validation process requires aerodynamic measurements in high quality flows, and acoustic measurements in an quiet and anechoic environment. The best way would be to carry out all measurements in the same test facility, but such installations are very rare and inevitably lead to compromises. Indeed, aerodynamic wind tunnels are often noisy and incompatible with any acoustic measurement, whereas anechoic wind tunnels are seldom convenient for local flow measurements using traditional probes (pitot tubes, hot wires, etc.) or laser velocimetry (particule image velocimetry, laser doppler velocimetry, etc.). Another approach consists in successively performing aerodynamic measurements and acoustic measurements in two different test facilities. Of course, that supposes to check flow reproducibility between installations, and thus to carry out a minimum of aerodynamic measurements in the acoustic test facility to correctly align the model with respect to the flow. If this aspect is fulfilled, the approach makes it possible to choose test facilities whose equipments and characteristics are perfectly adapted to each type of measurements. EXAVAC programme (Experimental Aerodynamic and Acoustic programme for the Validation of CFD/CAA methods applied to airfoil noise prediction) aims to constitute a complete database on an academic airfoil in terms of unsteady aerodynamics and acoustics. The test model is a NACA0012 airfoil with a chord length of 0.5 m, a span length of 1.4 m, and a blunted trailing edge with a thickness of 2.5 mm. Measurements were carried out in two ONERA’s test facilities : the F2 aerodynamic wind tunnel and the CEPRA19 acoustic wind tunnel. This paper describes the experimental programme and its principal results.

This paper deals with the numerical study of installation effects, a subject which is relevant of the more general purpose of the aircrafts noise prevision and reduction. As an example of installation effects, we study here the potential shielding effect that an empennage airfoil could offer to the reduction of the aft fan noise produced by a coaxial engine. Because the numerical simulations constitute an as powerful as cheap tool of investigation, we use for it a solver developed at ONERA, the sAbrinA CFD/CAA platform. In particular, we take benefits from an innovative overlapping method recently implemented in it - a method which allows handling more easily the solid obstacles to be considered. As an illustration of both the methodology and the tool, a full-3D aft fan noise propagation is then computed over an installed (over an airfoil) engine.

This paper presents an application of a Computational AeroAcoustics (CAA) hybrid process to the three-dimensional numerical prediction of the downstream fan noise of coaxial engines. This process may include the acoustic refraction effects resulting from the propagation through the highly inhomogeneous (subsonic hot) jet mean flow associated to the nozzle. The CAA hybrid methodology associates two different acoustic methods. Firstly, the near and mid field propagation through the (possibly inhomogeneous) mean flow is computed with an Euler, high order, finite differences solver. Then, the far field noise is calculated by use of a classical Kirchhoff integration. First computations are performed over a classical axi-symmetric coaxial nozzle in a fluid at rest, and are validated against BEM simulations. Then the simulations are performed including the jet viscous mean flow previously computed with a RANS solver. These computations show that the acoustic directivity pattern of fan noise is strongly modified when the jet mean flow is included in the simulation. Finally, a last computation is performed over a modified engine presenting a scarf of it secondary engine, this highlighting the abilities of the CAA process to handle full 3D problems, as well as underlining the potential shielding effect of such a structural modification of the nozzle.

This paper addresses the numerical simulation of the acoustic scattering and propagation through non-uniform mean flows using a high-order, finite-difference, full-conservative Euler solver and, more precisely, the specific problems raised by multi-block structured grids when complex geometries are considered. Two examples of two-dimensional computations of this type are presented. The first case, a benchmark problem from the 4 th CAA Workshop, is the simulation of the acoustic scattering from multiple rigid circular cylinders, without mean flow, of the sound generated by a spatially distributed, axisymmetric source. The proposed solution uses conformal multi-domain grids associating body-fitted grids near solid walls and quasicartesian grids elsewhere. The solutions are favorably compared to analytical data. The second example is the simulation of the acoustic scattering from a high-lift wing section with deployed slat and flap, from a source located in the slat cove region. An acoustic grid is derived from a CFD grid previously used to compute the non-uniform viscous mean flow around the wing. A first acoustic computation, performed with the fluid at rest, is favorably compared to a Boundary Element Method solution. Then a second computation is performed on the non-uniform mean flow. The comparison of both solutions highlights the influence of the flow on the directivity diagram of the sound field.

This paper describes a new step in the development of a Computational AeroAcoustics (CAA) process whose general long term objective is the numerical prediction of the aerodynamic sound radiated by the airframe of large aircraft at approach, and especially the noise generated by deployed high lift devices such as slats and flaps. The proposed 3-step hybrid process combines CFD (Computational Fluid Dynamics) techniques and acoustic numerical methods, each one being adapted to a particular domain in which specific physical fluid mechanisms are simulated solving an adequate set of equations. In a first step, the nearfield unsteady flow is computed via a compressible three-dimensional LES (Large Eddy Simulation). In a second step, LES-computed perturbations are injected at the inner boundary of a larger domain in which the outward propagation of small perturbations over a non-uniform mean flow is simulated using LEE (linearized Euler equations). In the third and last step, the acoustic field radiated at the external boundary of the LEE domain becomes the entry data of a Kirchhoff integration which provides the noise radiated in the far field. The critical point of the process is the coupling, via an interface, of the LES with the LEE. This process has been carefully studied using analytical fields, an acoustic point source monopole and a convected Eulerian vortex. It has been found that the correct injection of such fields requires severe conditions in terms of space resolution, conditions which are especially difficult to meet for purely vortical fields. In a former study, the LES of the unsteady flow around a NACA0012 airfoil has formed the basis of numerical noise predictions using acoustic integral methods. In the present paper, the same LES is used as a basis for the 3-step CAA process. First results revealed the generation of non-physical noise at the boundary interface where the airfoil's turbulent wake is injected in the Euler domain. Additional tests based on the injection of an analytical vortex suggest that this problem was most probably caused by the under-resolution of the injected vortical structures. This difficulty was not solved, but by-passed by using a LES/LEE interface which did not intercept the airfoil's wake. The final result integrates the three components, including the nearfield LES, the midfield noise propagation using LEE and the farfield noise radiation using the Kirchhoff integral. __________________________________________________ Copyright © 2002 by ONERA. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission INTRODUCTION The general context of this paper is the numerical prediction of the aerodynamic noise generated by the high lift devices HLD, slats and flaps of large airliners, an important contributor to the total radiated airframe noise, especially in approach configuration. It is commonly admitted that the design of new low-noise HLD concepts incorporating specific noise reduction devices, although still relying on necessary experiments, will take growing advantage of the numerical simulation in terms of lower costs and shorter delays, especially considering the spectacular continuing progress of Computational AeroAcoustics (CAA) methods. The problem of the numerical simulation of HLD noise is still beyond the capabilities of complete Direct Numerical Simulation (DNS), so hybrid methods are used in most practical cases. Figure 1 sketches the possible numerical strategies, showing how the nearfield turbulent flow and the farfield noise are computed separately. The idea is to divide the physical space into several domains, in which specific physical mechanisms are simulated using the most adequate set of equations with the cheapest discretization strategy. Computational Fluid Dynamics (CFD) techniques are used to simulate the nearfield flow which contains the aerodynamic noise sources. Available techniques include steady ReynoldsAveraged Navier-Stokes (RANS) computations, in conjonction with stochastic models of the wavenumber-frequency spectrum of the turbulence [1-3], unsteady RANS methods [4-5], and Large Eddy Simulation (LES) [6-9]. This local flow solution has to be coupled to an acoustic numerical technique for the prediction of farfield noise. The most practical formulations are the integral methods such as Lighthill's analogy [7] [10] (including the Ffowcs WilliamsHawkings (FW-H) equation [4, 5, 11, 12]), the Boundary Element Method (BEM) [13] and the Kirchhoff integral. In a former study, the compressible LES of the unsteady flow around a symmetrical NACA0012 airfoil with a blunted trailing edge has formed the basis of airfoil aerodynamic noise predictions. A detailed analysis of the nearfield unsteady flow showed that the local aeroacoustic characteristics were correctly simulated, including the local acoustic field. This suggested to define a control surface around the airfoil, on which the acoustic nature of the pressure field was established. The pressure field and normal derivative on this surface where used to compute the farfield noise via a 3D Kirchhoff method. In a second step, another noise prediction based on the Ffowcs WilliamsHawkings equation was performed using the same LES data. 8th AIAA/CEAS Aeroacoustics Conference & Exhibit Fire 17-19 June 2002, Breckenridge, Colorado AIAA 2002-2573 Copyright © 2002 by the author(s). Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.