24TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES

1

ABSTRACT

A study of wing control surface effectivenesswas carried out using numerical simulationswith the advanced Reynolds Averaged Navier-Stokes solver elsA. Non-coincident meshingtechniques were used as to make the meshgeneration process more flexible. The firstapplication attempts to predict an aileroneffectiveness using the patched grid meshingtechnique combined with a mesh deformationtool in order to operate the aileron deflection.The second one deals with spoiler deploymentand involves the Chimera technique, whichallows separating the spoiler meshing from thewing meshing and so avoiding a complete meshre-generation for each spoiler deflection.

NOMENCLATURE

α Angle of attackβ Sideslip angleδa Aileron deflection angleδs Spoiler deflection angleM0 Free stream mach numberRe Reynolds numberq/E Aero-elastic coefficient (q is the dynamic

pressure and E the Young’s modulus)CL Lift coefficientCD Drag coefficientCl Rolling moment coefficientCmc Hinge moment coefficientϕ Twist anglex/c Adimensioned coordinate x by the local

chord

ych Adimensioned coordinate y by half span,=(y-yroot)/b

chz Local load in z direction

1 INTRODUCTION

The correct prediction of handling qualitiesand hinge moments induced by the deploymentof wing control surfaces (spoilers and ailerons)is a crucial point in the general aircraft sizingprocess with a strong impact on the final aircraftweight. The complexity of the aerodynamicflows around deployed control surfaces and theimportance of the flight envelope to be coveredmade difficult the use of CFD in the elaborationof Aerodynamic Data. Until now, only verytime-consuming and costly wind tunnel testsand not very accurate semi-empirical methodswere used.

For a long time, CFD has been intensivelyused at Airbus for shape design andoptimization. As configurations are moderatelycomplex and studies focused on slightgeometrical variations, an efficient coincidentstructured mesh generation suite has been set uparound HEXA© mesher.

Thanks to the recent CFD progress inmeshing techniques, convergence accelerationand calculation performance, more and morenumerical simulations are involved inAerodynamic Data generation. However, theclassic coincident structured grid approach doesnot seem suitable for complex configurationssuch as deployed ailerons and spoilers.

NUMERICAL SIMULATIONS AROUND WINGCONTROL SURFACES

Guillaume FillolaAIRBUS, Toulouse, Fr

Marie-Claire Le Pape Marc Montagnac ONERA, Châtillon, Fr CERFACS, Toulouse, Fr

Keywords: Aileron, Spoiler, Chimera Mesh, Patched Grid, Wall Law

GUILLAUME FILLOLA

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Today, the following meshing techniquesappear to us as the most promising for controlsurfaces configurations:• The Patched-Grid technique, which allows

meshing independently on each side ofshared boundaries between two blocks, isthus more appropriate to build independentwing section grids.

• The Chimera technique is almost the sametechnique as the Patched Grid method but isfurther enriched by the overlapping gridcapability. Its principle is to meshindependently different bodies and then totake into account interactions between thedifferent components by interpolations.

• The Wall Law technique consists inapplying the “linear-logarithmic” law on thefirst cell, the size of which can be muchlarger than y+=1. Performance is thenimproved during calculation withoutdowngrading the solution accuracy.Moreover, it facilitates interpolation issuesfor Chimera technique.This paper first presents a brief summary

of the flow solver elsA, as well as the particulartechniques to be used. Following is a discussionaround two numerical simulations, withcomparison to experimental results. Aileroneffectiveness results using the Patched Gridtechnique will be presented first, followed byspoiler effectiveness results using the Chimeratechnique.

2 NUMERICAL METHODS

2.1 elsA solver

With the objective of federating allnational research teams and taking advantage ofolder functionalities implemented in separateCFD codes, ONERA has been developing a newgeneration solver called elsA since 1996, in co-operation with CERFACS since 2000. It hasbeen designed according to an Object Orienteddesign method and it is mainly coded with C++language, even though the most CPU-expensiveloops are coded with Fortran language for betternumerical efficiency. This innovative approach

leads to more upgradeable and inter-operableaerodynamic functions, and thus contributes to abetter integration of different development [1].Some applications made in Airbus France withthe elsA solver are described in [2].

The main features and numerical functionsof elsA are listed below:• cell centered code dealing with structured

meshes.• classical central scheme for Euler model

(centered flux with scalar dissipation)• viscous flux computed from cell-centered

evaluations of velocity and temperaturegradients, with possible correction values atinterfaces.

• classical algebraic and transport equationsturbulence models (all of them followingBoussinesq’s assumption).

• backward-Euler time integration associatedwith the LU implicit method.

• Wall Law treatment possibility for wallboundary conditions [3].We describe below the specific techniques

used in this study for aircraft applications.

2.2 Patched Grid technique

Complex configurations are decomposed inmany structured meshes. These blocks areconnected to each other with shared interfacesthat impose constraints on the grid generation.Indeed, a local mesh refinement in one of theblocks spreads through the entire computationaldomain with classical coincident structuredgrids.

The Patched-Grid technique allowsmeshing two blocks independently on each sideof their shared boundaries. A local meshrefinement has then a lesser impact on the otherblocks. Thus the computer memory required andCPU time are reduced since the number of gridpoints decreases.

The most important characteristic of thistechnique is to provide the conservationproperty of the numerical scheme as shown in[4] and [5]. This paper uses the approachdescribed in [6]. Further details, complementsor other methods related to this technique can befound for example in [7], [8], [9] and [10].

3

NUMERICAL SIMULATION AROUND WING CONTROL SURFACES

A bi-dimensional overview of the patchedgrid technique is shown in Figure 1.

Figure 1 : Two blocks with a shared patchedgrid interface.

The indices (i,j) refer to the location ofcells in both blocks and the border surface islocated at the index _. For comprehensivepurposes, it is assumed that the border surface isplane that is to say there is neither overlappingnor gaps between cells that are next to thesurface border. The principle of the patched gridis explained with the example of the cell (i1,j1)of block 1. The numerical flux through theinterface AB can be written:

),()1(

),(

12,21,11

2,21,1111

1,2/1

+−

+==

jijiBM

jijiAMAB

WWF

WWFFFj

α

α(1)

with AB

AM=α and W represents the

conservative state vector or any other fields andα is the intersection area computed from the twoborder interfaces of the cells (i1,j1) and (i2,j2).The intersection surface of these two interfacesis obtained with the intersection polygon givenby a Sutherland-Hodgman polygon-clippingalgorithm [11].This treatment of spatial fluxes enables tomaintain the global conservation along thepatched grid border surface. Indeed, the above

operations are applied on each blockindependently and hence 21

AMFFAM = for planar

border surfaces. If the border surface is curved,then this technique is said quasi-conservativesince 1

AMF and 2AMF might slightly differ

according to the point projection process.The ghost cells are filled thanks to an α -weighted interpolation process of the statevector in order to keep the efficiency of implicittime-integration algorithms.

2.3 Chimera technique

The Chimera method is based on anoverset grid technique [12]. The principle is togenerate independent meshes around differentbody elements, and to solve the global flow byusing interpolation technique in the CFD solver.More precisely, on the one hand the mesh areasoverlapped by bodies are not computed by thesolver and body influence comes from a cellcrown around each body; on the other hand,domain influence goes through outflowboundaries. This technique allows almostindepend body meshing; meshes must onlysufficiently overlap to allow interpolations. Asfor the independent bodies, refined meshesmake this constraint respected. However, whenbodies are next to each other and because ofblanking, areas near junctions are notdiscretized. In order to by-pass this difficulty,the solution - used in the present work - is togenerate a mesh leaning on the body of the othermesh.

In the elsA software, the interpolation ispiecewise linear by tetrahedron, each cell beingdivided into 24 tetrahedrons. Bodies aremodeled by a great number of parallelepipeds.Interpolation cell search becomes efficient byusing a preconditioned Cartesian grid and otheracceleration techniques to find the interpolationtetrahedron [13]. In order to reduce overlappingconstraints and to avoid some points frombecoming orphan, extrapolation fromneighbouring cells is allowed; the numericalscheme is also degenerated on overlappingboundaries and around bodies, thus

Block 1 Block 2

GUILLAUME FILLOLA

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interpolation crown and boundaries have awidth of one cell [14].

3 AIRCRAFT APPLICATIONS

3.1 Prediction of aileron effectiveness usingthe Patched Grid technique

Aircraft wings are equipped with aileronsthat have several functions in terms of handlingqualities. They are usually activated to create arolling moment with a dissymmetricaldeflection; but they are also deflected in asymmetrical way in order to modify punctuallythe aircraft lift (in high lift configuration forexample). Aerodynamically, an aileron pulls upor down the aft part of a small wing section,modifying so the local load. As a result, theglobal lift and/or the rolling moment change.

This first application aims thus atpredicting the variation of global aerodynamiccoefficients due to an aileron deflection in atransonic flow field. The studied configurationis a half wing-body equipped with an outboardaileron separated from the wing by two lateralgaps; it is the model used in the HiReTTEuropean project [15]. The surface definition ofthe wing corresponds to 32.5.106 Reynoldsnumber and q/E=0.4395 wind tunnel conditions.Some calculations on the same configurationhave already been done in Airbus in theHiReTT framework [16] with a differentmeshing approach.

3.1.1 Grid generation process

With a traditional coincident mesh, theaileron deflection would lead to shear cellsconfined in the gap area. One way ofovercoming this problem consists in meshingindependently the aileron wing section byplacing two vertical patched grid plans in themiddle of each gap (Figure 2). In this way, theaileron deflection will only modify the areabetween the two patched grid plans notdegrading the cell skewness. The mesh topologyis a typical wing-body mesh of 6.500.000 nodes,with a CH grid around the fuselage, and a CH

grid around the wing completed by a pivot onthe wing tip.

Figure 2 : HiReTT wing-body with a δa=3°deflected aileron – inner gap overview

In order to speed up the mesh generationprocess, a mesh deformation procedure, calledRAiD (Rudder and Aileron Deflection)dedicated to control surfaces deflection has beendeveloped. It is composed of two independentprograms.

The first one generates a surface field(dx,dy,dz) simulating the aileron deflection. Itconsists in calculating the control surfacemovement with respect to a hinge axis and thento operate a smooth joining of this area to thewing box, preserving the continuity in positionand tangency. The second one, called VOLDEF,propagates this deformation in the mesh volumeusing an analytical method developed in Airbus[17]. It is based on the distinction between twokinds of surfaces:

• Modified surfaces Γm: which contain thesurface field

• Damping surfaces Γa: which are notdeformed

The deformation applied to a point M isthus given by the following formula:

∫

∫

Γ∈

Γ∈

Γ

Γ

⋅=

mP

mP

dPMd

dPMd

P

M

α

α

δ

ηδ

),(

1),(

)(

)(

r

r , α>2 (2)

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NUMERICAL SIMULATION AROUND WING CONTROL SURFACES

η is a damping function. This parameterhas a value of 1 near a modified surface andtends toward 0 near a damping surface.

β

κ

η

−

= da

dm

e (3)

with ( ))P,M(dmindmmP Γ∈

= and ( ))P,M(dmindaaP Γ∈

=

This procedure allows obtaining a newmesh with an aileron deflected in less than 3hours on a workstation and, thus, a wide rangeof deflected configurations can be easilyinspected.

Figure 3 depicts a bi-dimensionalapplication of the RAiD procedure applied on asimple OAT15A profile.

Figure 3 : aileron deflection by the meshdeformation procedure (left δa=-10°, right

δa=+10°)

3.1.2 Discussion of resultsOn the basis of the clean configuration,

three aileron settings (δa=-3°, 3°, 6°) areoperated with the RAiD procedure. Turbulentcalculations are then performed on each mesh,although we only have experimental results atδa=0° and 3°. The flow parameters are thefollowing:• M0=0.85, α=1.5°• Rec=32.5 106, q/E=0.4395• [SA] turbulent model in fully turbulent flow

Figure 4 presents results in terms of CLeffectiveness (experimental CL is measuredwith balances). A nearly linear behavior appearsat low deflections, δa=-3° and +3°, whereas anobvious non-linearity occurs at δa=6°. Thecomputed global lift on clean configurationapproximately matches the experimental value.However, the increment of lift induced by theaileron deployment is not well simulated. Thus,although errors between calculation and

experiment at both δa=0° and 3° are less than4% on global CL, this error reaches more than100% on ΔCL!

Figure 4: Aileron effectiveness, M0=0.85,α=1.5°

Figure 5 : Cp distribution for 2 deflectedconfigurations, δa=0° and 3°, M0=0.85,

α=1.5°

On the Figure 5, Cp-distributions aroundtwo wing sections (one located on the middle ofthe wing, the other one at half span of theaileron) are represented, for two deflectedconfigurations, δa=0° and 3°. (Unfortunately,three pressure probes, located on aileron lowersurface, failed during the W/T test on thedeflected configuration.) First of all, we observethat the structure of the flow is quite well

η=83.7%

η=61.1%

η=83.7%

η=61.1%

• HiReTT W/T δa=0°• HiReTT W/T δa=3°__ elsA δa=0°__ elsA δa=3°

GUILLAUME FILLOLA

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simulated on the clean configuration. Indeed,CFD predicts well the double shock topologyand sets it roughly at the good location. Themain aileron deployment effects are also wellcaptured by the calculation: the lower surfacepressure increases and a suction peek appearsalong the hinge line.

The shock location and the supersonic arealevel express the major discrepancy betweencalculation and experiment. We can see onFigure 6 that δa=3° calculation predicts a too-far-aft shock location, as well as an over-estimation of the supersonic plateau.Nevertheless, the comparison between non-deflected/deflected calculations highlights aconsistent behaviour of the flow with a bi-dimensional approach (see [18]). On thecontrary, in the experiment the shock movesforward and the supersonic plateau decreasesmaking us thinking about an aero-elastic effect.

Figure 6 : Cp distribution on the aileronsection, zoom on the supersonic area

In the HiReTT project framework, directaero-elastic calculations have been done on thisstudied configuration, at the RWTH University[19]. It allows estimating the aerodynamicdeformation only due to an aileron deflection.These results have been provided to us in termof twist angle and wing bending (Figure 7).They enlighten a slight reduction of the twistangle (ϕmax~0.3°) in aileron sections (η=0.7 to0.95). This is mainly due to the elevation of aft-load implying aileron sections to pitch down.

This deformation has been applied to ourrigid deflected mesh using the VOLDEFprocedure (Figure 8) and another e l s A

calculation has been carried out with the sameflow parameters.

Figure 7 : Wing bending and twist angle dueto a δa=3° aileron deflection.

Figure 8 : Aero-elastic effect on the aileronsection

Figure 9 : Cp-distribution on the δa=3°configuration, flexible effect, M0=0.85, α=1.5°

The reduction of the twist angle in aileronsections involves a reduction of the effectiveangle of attack. As the consequence the shockmoves forward, the supersonic plateaudecreases, and the calculated flow bettermatches to experimental values (Figure 9). Thismovement results in a slight decrease of thedeflected configuration lift, and thus, a good

__ δa=0°__ δa=3°__ δa=3° + flexible effect

• HiReTT W/T δa=0°• HiReTT W/T δa=3°__ elsA δa=0°__ elsA δa=3°

ϕ

η=83.7%

η=61.1%

• HiReTT W/T δa=3°__ elsA δa=3°__ elsA δa=3° + flexible effect

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NUMERICAL SIMULATION AROUND WING CONTROL SURFACES

agreement is found in term of ΔCL(ΔCLflexible/ΔCLexp=90%, Figure 10).

Figure 10 : Aileron effectiveness, flexibilityeffect, M0=0.85, α=1.5°

3.2 Prediction of spoiler effectiveness usingthe Chimera technique

This second application intends to simulatethe influence of spoiler deployment on globalaerodynamic coefficients in a transonic flowfield. A spoiler is a control device located in themiddle of the upper wing, which produces flowseparation when it is deflected. It causestherefore the lift to decrease and the drag toincrease. When they are symmetricallydeflected, spoilers are used as airbrakes andplay a part in landing or during an emergencydescent. With a dissymmetrical deflection, theyintervene to accentuate aileron action and thusto create a high level of rolling moment

The studied configuration represents atypical AIRBUS wing body with externalspoilers deployed. The wing shape correspondsto wind tunnel condition (5.106 Reynoldsnumber) and was assumed as rigid whenspoilers are deflected.

3.1.1 Grid generation processThe complex geometry of the studied

configuration leads us to use the Chimeratechnique. Actually, one multi-block grid of3.500.000 nodes is dedicated to the wing + bodywhereas another one of 500.000 nodes, totallyindependent, is dedicated to the spoilers and a

wing portion (Figure 11). Thus, the meshingprocedure is made simpler and a good meshquality can be easily ensured.

Because the spoiler is joined to the wing,we have chosen to define a part of the wing bothin the spoiler mesh and in the wing-body mesh,in order to avoid the presence of orphan pointsat the intersection between the wing and thespoiler. Then, the wing grid is holed at thespoiler location in order to take into account thespoiler presence.

This double discretization of the wing wascausing interpolation problems in the originallow Reynolds mesh because of the surfacecurvature. Actually, cells closed to one body canbe located inside the second one. This problemis well known and Schwarz proposes a solutionin [20] that should be soon integrated in elsA.Waiting for this, we overcome the limitation byusing wall law technique.

Figure 11 : Airbus wing-body with a δs=20°deflected spoiler

Another interesting feature of the Chimeratechnique is that the spoiler deflection will onlyaffect the spoiler domain, and so a completemesh re-generation is not necessary for everyspoiler setting.

3.1.2 Discussion of resultsTo validate our calculations, we rely on

experimental results coming from a wind tunnel

GUILLAUME FILLOLA

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test campaign with few Cp measurements. Themock-up is a complete aircraft with pods, flap-trap-fairings, horizontal and vertical tail planes(H/VTP), and we will compare it to ourcalculations around a simple wing-body. Thesedifferences between the two configurationsinvolve several aero dynamical discrepanciesthat we will describe below.

The HTP increases the CLα slope becausethe local HTP angle of attack increases when αincreases (despite a reduction due to wingdownwash). The pods induce a negative ΔCL(almost constant with α), mainly due to aforward location of the shock. However, theseimpacts are the same on the clean wing and onthe wing with a deployed spoiler; that is why wewill only express our results in term of variationwith respect to the clean wing characteristics.

The following calculation matrix has beencarried out for three deflected configurations,δs=0°, 20° and 45°:

αδs=0°,20°, 45° 0 2.5 3.5 4.0 5.0 6.0 7.0

0.7 x x x x xM

0.85 x x x x x

Table 12 : calculation matrix

Lift Effectiveness

Figure 13 : CL spoiler effectiveness, leftδs=20°, right δs=45°, M0=0.7, 0.85 (line :

experiment, symbols : calculations)

A good agreement between experiment andnumerical calculations is found on ΔCL(α)(Figure 13). This curve is characterized by tworegimes. At low angles of attack, ΔCL is almostconstant (this level is quite well simulated by

CFD at M0=0.85). At high angles of attack, anon-linearity appears, bringing about a drop ofΔCL. This phenomenon occurs because, for agiven α, the wing with a deflected spoiler stayslinear whereas first non-linearity appears onclean wing (Figure 14).

Figure 14 : CL(α) δs=0°,45°, M0=0.85

Figure 15 : -Cp distribution on δs=0°, 20°,45° deflected wing, M0=0.85, α=2.5°

On Figure 15 a visualization of the flowthrough –Cp distribution, around three deflectedwings is given. The spoiler deployment pushesforward the shock wave and thereby causes theloss of lift. On the spoiler upper surface thepressure increases with the spoiler deflection

δs=0°

δs=20°

δs=45°

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NUMERICAL SIMULATION AROUND WING CONTROL SURFACES

whereas a strong acceleration appears on theinner spoiler lower surface. It creates highdifferences between the spoilers’ hingemoments.

Spoiler Hinge MomentHinge moment coefficients on each spoiler

are plotted on Figure 16. Because of the lowersurfaces over-speed, the inner spoiler Cmc isquite more negative than the other spoiler’sones.

At high angles of attack, a shock separationappears and encompasses the spoilers,beginning at the outer spoilers (Figure 17, thered surfaces show the separation area). Thiscauses these spoilers Cmc to increase until theybecome positive.

Figure 16 : Spoilers hinge moment (Cmc),δs=20°, M0=0.85 (line : experiment, symbols :

calculations)

Figure 17 : visualization of separation area(in red) – spoiler in blue – M0=0.85, α=5.0°

Drag EffectivenessNear field drag has also been investigated.

Results in term of ΔCD are presented in Figure18. The main effects of spoiler deflection arewell captured by CFD: ΔCD decreases when αincreases until becoming negative at high anglesof attack (meaning that spoiler deploymentreduces drag instead of increasing it).

Figure 18 : CD spoiler effectiveness, δs=20°,M0=0.7, 0.85 (line : experiment, symbols :

calculations)

Rolling Moment EffectivenessThe rolling moment effectiveness analysis

highlights a discrepancy between calculationand experiment that does not appear on the lifteffectiveness curve. ΔCl is overestimated (inmodulus, Figure 20) while ΔCL better matchesexperiment (Figure 13).

Taking advantage of the Chimera methodflexibility, a new mesh representing a completedissymmetrical aircraft with a spoiler deflectedon the left wing, has been rapidly built. OnFigure 19, the wings loads of this completeaircraft are compared to the half aircraft ones.The differences between the two wing loadshighlight an interaction between the left and theright wing. Right wing inner load decreaseswhile left wing inner load increases; as a result,the rolling moment decreases and comparesfairly to experimental values (Figure 20). TheΔCL(α ) curve is quite the same as the halfaircraft one because of a compensationphenomenon between right and left wing’s lift.

GUILLAUME FILLOLA

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Another result of the complete aircraftcomputation is the creation of a sideslip field ofabout 1° at the fin location (see Figure 21).Airbus Clβ data modeling indicates that thisimpact would cause a positive Cl responsible ofan additional ΔCl around 0.01.

Figure 19 : Computed wing load - interactionleft wing/ right wing – M0=0.85, α=2.5°

Figure 20 : Cl spoiler effectiveness – δs=20° -M0=0.85 (line : experiment, symbols :

calculations)

Figure 21 : β field visualization on the finarea due to a δs=20° deflected spoiler on left

wing

In light of these numerical results, aconclusion is that the use of a complete mock upseems to be necessary for the correct assessmentof rolling moment in wind tunnel test.

4 SUMMARY AND CONCLUSIONS

Two numerical applications have beencarried out with elsA solver in order to studywing control devices effectiveness. Throughouttheir results, the non-coincident structured gridapproach has proved its efficiency and shouldbe integrated in the future Aero-Data generationprocess.

In the first application, aileron deflectionsare studied, using the Patched Grid techniquecoupled to a mesh deformation package, RAiD.Aileron deflection of a rigid wing has beensimulated at first, showing an over-estimation ofaileron effectiveness explained by an aero-elastic effects induced by aileron deflection.Even if bending and twist effect have beentaken from HiReTT outcomes, this aero-elasticeffect will be estimated by elsA solver as soonas the flexible matrix is built.

The second application uses a Chimeraapproach and deals with spoiler deployment. Itpermits to adequately estimate the spoilers’hinge moments and effectiveness (in terms oflift, drag and rolling moment), while keeping asimple mesh. Thanks to the Chimera methodflexibility, a complete dissymmetrical aircraftmesh has been carried out, and an interaction

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NUMERICAL SIMULATION AROUND WING CONTROL SURFACES

right wing/left wing has been pointed out. Thesepromising results should be improved in a nearfuture, with the integration by ONERA intoelsA solver of a multi-grid convergenceacceleration technique compatible with theChimera method [21].

In the prospect of speeding up even morethe wing control surface meshing process, anautomatic procedure based on HEXA© mesher,is thus in development in Airbus. Its goal willbe to extract the spoiler geometry from the cleanwing, to deploy it and, then, to generate thespoiler domain around.

ACKNOWLEDGEMENTS

The authors acknowledge the AirbusMethods and Tools team for their technicalsupport, and especially Luis Barrera whofollows all the study with a great attention.

CONTACTS

• guillaume.fillola@airbus.com• Marie-Claire.Le_Pape@onera.fr• Marc.Montagnac@cerfacs.fr

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