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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.

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.

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.

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 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.

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.

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.

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 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.

This paper addresses the numerical simulation of the acoustic propagation through an aerodynamic flow and its interaction with solid bodies. First, a general conservative formulation is derived from Euler's equations using a small perturbation hypothesis. In a second time, a discrete form is obtained with the help of high-order explicit schemes. Then, boundaries treatments and curvilinear or filtering aspects are examined. Finally, as an illustration and a validation of the E3P code that has been developed on these theoretical basis, an extensive set of numerical simulations is presented. It should be showing the wide range of possible applications that this tool, taken alone or coupled with others, could offer.

This study, that is relevant from the turbofan engines noise prediction/reduction, aims at CAA‐computing the aft fan noise propagation/radiation of a realistic full‐3D exhaust (with pylon and internal bifurcations), the latter being affected of (i) typical in‐flight (take‐off) thermodynamic conditions and of (ii) a representative fan noise modal content. As for previous studies conducted over baseline geometries, this CAA computation is conducted following the usual hybrid process, where a preliminary aerodynamic calculation provides a heterogeneous steady mean flow on which an acoustic calculation is then conducted A RANS computation is first performed, delivering the stationary jet mean flow characterizing the 3D exhaust in its typical 'take‐off flight' (M∞ = 0.25). A CAA grid (22 blocks, 28 millions cells) is then derived from the CFD one, before the RANS steady jet mean‐flow is interpolated on it. After what the CAA computation is computed, a fan noise mode (26, 1) being emitted at a reduced frequency of kR = 30 (1 BPF) in the upstream of the engine's secondary exhaust, and numerically propagated along and outside the latter. Finally, a Kirchhoff post‐treatment provides the far‐field radiation characterizing these engine geometry, modal content and thermodynamic conditions.