Analysis and Design of a Leading Edge with Morphing Capabilities for the Wing of a Regional Aircraft—Gapless Chord- and Camber-Increase for High-Lift Performance

applied sciences

Article Analysis and Design of a Leading Edge with Morphing Capabilities for the Wing of a Regional Aircraft—Gapless Chord- and Camber-Increase for High-Lift Performance

Conchin Contell Asins 1, Volker Landersheim 1,*, Dominik Laveuve 1 , Seiji Adachi 2, Michael May 3 , Jens-David Wacker 1 and Julia Decker 1

1 Fraunhofer LBF, 64289 Darmstadt, Germany; [email protected] (C.C.A.); [email protected] (D.L.); [email protected] (J.-D.W.); [email protected] (J.D.) 2 Fraunhofer IBP, Nobelstraße 12, 70569 Stuttgart, Germany; [email protected] 3 Fraunhofer EMI, Ernst-Zermelo-Straße 4, 79104 Freiburg, Germany; [email protected] * Correspondence: [email protected]

Abstract: In order to contribute to achieving noise and emission reduction goals, Fraunhofer and Airbus deal with the development of a morphing leading edge (MLE) as a high lift device for aircraft.

Within the European research program “Clean Sky 2”, a morphing leading edge with gapless chord- and camber-increase for high-lift performance was developed. The MLE is able to morph into two

Citation: Contell Asins, C.; different aerofoils—one for cruise and one for take-off/landing, the latter increasing lift and stall Landersheim, V.; Laveuve, D.; angle over the former. The shape flexibility is realised by a carbon fibre reinforced plastic (CFRP) Adachi, S.; May, M.; Wacker, J.-D.; skin optimised for bending and a sliding contact at the bottom. The material is selected in terms Decker, J. Analysis and Design of a of type, thickness, and lay-up including ply-wise fibre orientation based on numerical simulation Leading Edge with Morphing and material tests. The MLE is driven by an internal electromechanical actuation system. Load Capabilities for the Wing of a introduction into the skin is realised by span-wise stringers, which require specific stiffness and Regional Aircraft—Gapless Chord- thermal expansion properties for this task. To avoid the penetration of a bird into the front spar of the and Camber-Increase for High-Lift wing in case of bird strike, a bird strike protection structure is proposed and analysed. In this paper, Performance. Appl. Sci. 2021, 11, the designed MLE including aerodynamic properties, composite skin structure, actuation system, 2752. https://doi.org/10.3390/ and bird strike behaviour is described and analysed. app11062752

Academic Editors: Rosario Pecora Keywords: morphing; leading edge; carbon ﬁbre reinforced plastic; gapless; CFD-analysis; high-lift and Sergio Ricci device; aircraft wing; actuation; bird strike

Received: 30 January 2021 Accepted: 15 March 2021 Published: 19 March 2021 1. Introduction Currently, the aviation industry is developing quieter and more efficient aircraft to Publisher’s Note: MDPI stays neutral reach the environmental goals of Flightpath 2050 [1,2]. New technologies are expected with regard to jurisdictional claims in to allow a 75% reduction in CO2 emissions per passenger kilometre and a 90% reduction published maps and institutional affil- in NOx emissions. The perceived noise emission of flying aircraft is to be reduced by iations. 65% [3]. The investigation and development of new high-lift devices provide not only an improvement of aerodynamics, but also a positive effect on the emission of noise and pollutants. Over the last years, the volume of air traffic has increased approx. 5% per annum. This means that air traffic will considerably increase within the next 20 years. Copyright: © 2021 by the authors. In order to sustain the CO2 and noise emissions, they have to be significantly reduced. Licensee MDPI, Basel, Switzerland. Traditional high-lift devices, such as Krueger flaps or slats, provide an increase of the This article is an open access article aerodynamic lift, but also an increase of the noise emission [4,5]. This increase in noise distributed under the terms and is due to the presence of a gap between the high-lift device and the main wing [6]. If a conditions of the Creative Commons gapless high lift device can be designed, this source of noise emission vanishes, which may Attribution (CC BY) license (https:// in turn reduce the noise footprint of the aircraft. One way forward is the development creativecommons.org/licenses/by/ of gapless, morphing high-lift devices. Additionally, gapless, morphing high-lift devices 4.0/).

Appl. Sci. 2021, 11, 2752. https://doi.org/10.3390/app11062752 https://www.mdpi.com/journal/applsci Appl. Sci. 2021, 11, 2752 2 of 28

could be an enabler for natural laminar flow (NLF) [7,8]. Therefore, for several decades, the aviation industry, as well as research institutions, have been investigating morphing high-lift devices. Some examples are briefly described below. Communier and co-workers experimentally and numerically demonstrated that morphing leading edge concepts can potentially replace traditional leading-edge devices in terms of their aerodynamic properties [9]. However, for the design of a morphing leading edge, the aerodynamic efficiency is not solely important. The MLE designed here was intended to be placed on an unmanned aerial vehicle (UAV). For this slow-flying UAV, the polymer skin was designed in a way that it did not contribute to the load carrying capability of the wing. However, for larger aircraft operating at higher velocities, the structural aspects of the skin, and in particular the fluid-structure interaction, should not be forgotten [8]. Within the scope of the European project “Smart high lift devices for next generation wings” (SADE) between 2008 and 2012 diverse concepts for deformable leading and trailing edge were considered. For the leading edge, approaches with aeroelastic effects, selective deformable structures with frames, as well as innovative actuator concepts were investigated [10]. Within SADE, Morishima tried to design a MLE with a carbon fibre reinforced composite skin for an A320 type aircraft. However, Finite Element simulations showed that the particular stacking sequences selected in the Cranfield design were not capable of withstanding the high aerodynamic loads imposed on the MLE. Therefore, the skin material was changed to aluminium [5]. Peel and co-workers developed a small-scale morphing leading edge demonstrator fea- turing a carbon fibre reinforced composite (CFRP) skin with stacking sequence [+10/−10]ns. The actuation concept makes use of filament-wound rubber muscles [11]. The German Aerospace Centre (DLR) has been investigating adaptive structures for several years, where the focal point is research on a suitable skin to reach the requested deformation. Within the project “Smart Leading Edge Device” (SmartLED), DLR realised a concept for a gapless leading edge in a 1:1 scale demonstrator. Glass fibre material with varied thickness of the laminate (ply drop-off) was used in conjunction with omega stringers, and an actuation kinematic was designed [10]. Within the framework of the “Smart Fixed Wing” platform of the European research programme “Clean Sky”, DLR developed flexible morphing skins composed of glass fibre-reinforced composite and a layer of elastomer [12]. Between 2011 and 2015, several morphing structures were developed, manufactured, and tested within the European project “Smart Intelligent Aircraft Structures (SARISTU)”. These included a morphing trailing edge [13], a morphing winglet [14,15], and a morphing leading edge [16,17]. The MLE developed within this framework follows the concept of an enhanced adaptive droop nose with a glass-fibre reinforced composite skin. The demonstrators were tested in the wind tunnel as well as subjected to bird impact loading to demonstrate the aerodynamic effectiveness and safety of the design. Besides the aerodynamic and structural design, the SARISTU project team also developed shape sensing systems based on Fiber Bragg Grating sensors [18], which have also been used by Ma and Zhen [19] as well as Swi˛ech[´ 20]. Additionally, load monitoring systems employing optical fibres and strain gauges have been developed within SARISTU [21]. These technologies help to verify that the MLE is operating within the operational window. TU Delft has also been working on the development of a morphing leading edge within the “Green Regional Aircraft” platform of the European Project “Clean Sky”. The approach followed by TU Delft foresees a deformable aluminium skin, which is actuated by a Linak LA23 series linear actuator [22]. It was demonstrated that the system is capable of reaching the target shapes with sufficient accuracy. Vasista et al. [23] developed a glass fibre reinforced plastics (GFRP) skin MLE, which is actuated by one harmonic drive rotational actuator per kinematic rib. Whilst the system showed to work satisfactorily, the authors commented that one of the big future challenges would be to reduce the structural mass Appl. Sci. 2021, 11, x FOR PEER REVIEW 3 of 29

Appl. Sci. 2021, 11, 2752 3 of 28

authors commented that one of the big future challenges would be to reduce the structural mass of the system. An alternative approach to classical actuation systems is the use of shapeof the memory system. Analloys alternative (SMA) for approach actuation, to as classical successfully actuation demonstrated systems is by the Ameduri use of shape [24]. memoryFraunhofer alloys (SMA) has also for investigated actuation, as efficient successfully aviation demonstrated structures bywithin Ameduri the “Green [24]. Re- gionalFraunhofer Aircraft” platform has also of investigated the European efﬁcient Project aviation “Clean structuresSky”. In this within context, the Fraunhofer “Green Re- LBFgional developed, Aircraft” platformmanufactured, of the Europeanand tested Project a 1:1 scale “Clean smart Sky”. morphing In this context, leading Fraunhofer edge demonstrator.LBF developed, It was manufactured, realised as a and full-scale tested morphing a 1:1 scale smartleading morphing edge of 3 leading m span edge for ground demon- demonstration.strator. It was realised The setup as a full-scalewas also validated morphing in leading a climate edge wind of 3 mtunnel span forunder ground icing demon- condi- tions.stration. Besides, The setup the electromechanical was also validated morphi in a climateng actuation wind system tunnel and under additional icing conditions. technol- ogiesBesides, were the integrated electromechanical in the demonstrator morphing actuation like optical system fibre and Bragg additional grating technologies sensors, for wereshapeintegrated reconstruction in the and demonstrator strain measurement like optical as well ﬁbre as Bragg an ice grating protection sensors, system for shapebased onreconstruction carbon nanotubes. and strain Figure measurement 1 shows the as wellleading-edge as an ice protectiondemonstrator system in the based climate on carbon wind tunnelnanotubes. [25]. Figure1 shows the leading-edge demonstrator in the climate wind tunnel [ 25].

Figure 1. Fraunhofer morphing leading edge demonstratordemonstrator in the climateclimate wind tunnel.tunnel.

Improvement of aerodynamic properties withoutwithout deterioration of acoustic properties can be achieved with adaptive and gapless high-lift devices. However, this kind of device is not not established established in in series series production production aircra aircrafts,fts, because because of ofthe the complex complex requirements requirements re- gardingregarding this this sort sort of of technology. technology. They They must must have have enough enough stiffness stiffness to to carry carry the aerodynamic loads as well as the ﬂexibilityflexibility to reach the differentdifferent desireddesired aerofoilsaerofoils forfor thethe wingwing [[26].26]. OfOf course, the leading edge (LE) design in these cases has to avoid a gapgap inin thethe connectionconnection with the wing. De Gaspari et al. [27] [27] have demonstrated demonstrated that the development of a morphing leading edge with relevance for the aeronautical industry is a very complex and thusthus challengingchallenging task covering several, sometimes rivalling, disciplines,disciplines, andand designdesign objectives.objectives. Some of the main aspects to be considered during the development of a morphingmorphing leading edge are summarized below: • Aerodynamic performance; • deformation capability of the skin material; • structural integrity including fatigue and impactimpact loading;loading; • actuation concept; • low structural weight; • mounting concept. Whilst there have been many efforts to addressaddress individualindividual aspects of aa morphingmorphing leading-edge design, design, only only a a few few works works have have addr addressedessed the the full full picture. picture. In these In these cases, cases, the skinsthe skins were were either either made made from from aluminium aluminium [5,23] [5 or,23 GFRP] or GFRP [17,27]. [17 However,,27]. However, none noneof these of broadthese broadapproaches approaches described described the development the development process process of a morphing of a morphing loading loading edge edgewith withCFRP CFRP skin. skin.Following Following a holistic a holistic design design approach, approach, Fraunhofer Fraunhofer LBF in LBF collaboration in collaboration with with Fraunhofer IBP, Fraunhofer EMI, and Airbus have developed a morphing leading edge (MLE) with gapless chord- and camber-increase for high-lift performance within Appl. Sci. 2021, 11, x FOR PEER REVIEW 4 of 29

Appl. Sci. 2021, 11, 2752 4 of 28 Fraunhofer IBP, Fraunhofer EMI, and Airbus have developed a morphing leading edge (MLE) with gapless chord- and camber-increase for high-lift performance within the framework of Clean Sky 2. In the present paper, the proposed design of the MLE is de- scribedthe framework and analysed. of Clean Section Sky 2.2 describes In the present the method paper, and the results proposed of the design CFD ofanalysis the MLE per- is formeddescribed on target and analysed. aerofoils Sectionfor cruise2 describes and high-lift. the method In Section and 3, resultsthe design of the of the CFD composite analysis structureperformed including on target skin aerofoils material for and cruise string anders high-lift.are described. In Section The actuation3, the design system of em- the ployedcomposite for morphing structure includingthe structure skin is materialdescribed and in Section stringers 4. areSections described. 5–8 address The actuation further relevantsystem employed aspects for for the morphing proposed the morphing structure device is described such as in fatigue Section test4. Sections of the skin5–8 material, address thefurther design relevant of a gap aspects closure for device, the proposed the propos morphinged mounting device process, such as fatigueand bird test strike of the analy- skin sis.material, the design of a gap closure device, the proposed mounting process, and bird strike analysis. 2. CFD Analysis 2. CFD Analysis 2.1.2.1. Analysed Analysed Airfoils Airfoils of of the the Morphing Morphing Wing Wing DependingDepending on on the the current current flight flight condition condition the the aerofoil aerofoil of of an an aircraft aircraft wing wing has has to to meet meet differentdifferent requirements. requirements. In In cruise, cruise, a a low low drag drag is is required required together together with with a a lift lift coefficient coefficient that that correspondscorresponds to to aircraft aircraft weight weight at at cruise cruise velocity. velocity. For For take-off take-off and and landing, landing, a a maximum maximum lift lift coefficientcoefficient is is required required in in order order to to reduce ta take-off/landingke-off/landing velocity velocity contributing contributing to to shorten shorten requiredrequired runway runway length length as as well well as as to to redu reducece perceived noise on the ground. Increasing thethe lift lift coefficient coefficient is is possible possible by by increasing increasing the the chord chord and and camber camber of of an an aerofoil, aerofoil, but but for for achievingachieving low low drag drag at at cruise, cruise, lower lower values values of of chord and camber are desired. The The MLE MLE allowsallows us us to to switch switch between between an an aerofoil aerofoil with with high high chord chord and and camber, camber, and and another another aerofoil aerofoil withwith lower lower drag drag at at cruise. cruise. The The aerofoils aerofoils and and flight flight conditions conditions considered considered here here are are for for a a regionalregional turboprop turboprop aircraft aircraft with with a a capacity capacity of of approx. approx. 40 40 passengers, passengers, 26 26 m m span, span, a a max. max. take-offtake-off weight weight of of 20 20 t, t, and and a a cruise cruise speed speed of 180 m/s. This This kind kind of of plane plane typically typically does does not not useuse high high lift lift devices devices at at the the leading leading edge edge up up to to now. now. ToTo confirm confirm the the aerodynamic aerodynamic effect effect of of the the MLE, MLE, four four aerofoil aerofoil profiles profiles having having different different LELE shapes shapes were were analysed, analysed, comparison comparison see see Figure Figure 22.. Aerofoil S4 is the targettarget aerofoilaerofoil forfor cruise,cruise, aerofoil aerofoil S5 S5 is is the the one one for for take-off take-off an andd landing landing with with increased increased chord chord and and increased increased camber.camber. Additionally, Additionally, it it was was analysed analysed whether whether intermediate intermediate ae aerofoils,rofoils, which which occur occur due due to theto theimplemented implemented actuation actuation syst systemem during during the morphing the morphing process, process, are aerodynamically are aerodynamically feasible.feasible. A monotonic A monotonic change change of lift of and lift anddrag drag is required is required during during the shape the shape transition. transition. For this For purpose,this purpose, the two the intermediate two intermediate baselines, baselines, called called BL and BL 050 and between 050 between S4 and S4 S5, and have S5, been have considered.been considered. Suffix Suffix HL implies HL implies that each that eachprofile profile has a has deployed a deployed trailing trailing edge edge (TE)(TE) flap flapfor analysesfor analyses in the in thehigh-lift high-lift condition. condition. For For CFD CFD analysis analysis in inthis this condition, condition, the the original original single single FowlerFowler flap flap of of the the prototype prototype wing, wing, which which the the MLE MLE development development is based on, was used.

FigureFigure 2. 2. LELE (leading (leading edge) edge) comparison comparison of of profiles proﬁles S4HL, S4HL, BLHL, BLHL, 050HL, 050HL, and and S5HL S5HL [28]. [28]. Appl. Sci. 2021, 11Appl., x FOR Sci. PEER2021, 11REVIEW, 2752 5 of 29 5 of 28

2.2. Details of 2.2.CFD Details Analysis of CFD Analysis Stationary Reynolds-averagedStationary Reynolds-averaged Navier-Stoke Navier-Stokess (RANS) simulation (RANS) was simulation performed was performedus- using ing a k-ω sheara k-stressω shear transport stress transport(SST) turbulence (SST) turbulence model [29]. model The [ wall29]. Thefunctions wall functions were used were used to do to do so-calledso-called high Re-type high Re-type flow simulation. flow simulation. In the project, In the project, the cruise the cruisecondition condition was de- was defined as fined as 240 knots240 knots (134 (134m/s) m/s) flow flowvelocity velocity and andthe standard the standard atmosphere atmosphere at an at analtitude altitude of of 25,000 feet 25,000 feet (air(air density density ⍴ =ρ 0.549= 0.549 kg/m kg/m3). The3). The landing landing condition condition was was defined defined as 140 as 140knots knots (72 m/s) (72 m/s) flow flowvelocity velocity and the and sea-level the sea-level atmosphere atmosphere (⍴ = 1.23 (ρ = kg/m 1.23 kg/m3). The3). Reynolds The Reynolds number number per 1 m per 1 m amountsamounts to 1.42 to × 1.42107 and× 10 1.477 and × 10 1.477 for× cruise107 for and cruise landing and landingconditions, conditions, respectively. respectively. The 2D profilesThe in 2D Figure profiles 2 were in Figure first extr2 wereuded first by 1 extruded m along bythe 1 span-wise m along the direction. span-wise direction. Each of them Eachwas then of them placed was in then a computational placed in a computationaldomain of 80 m domain × 80 m × of 1 80m. mThe× peri-80 m × 1 m. The odic boundaryperiodic condition boundary is imposed condition on both is ends imposed in order on to both do 3D ends analysis in order by assuming to do 3D analysis by infinite span length.assuming In infinitethis domain, span length.a hexahedral-dominant In this domain, a hexahedral-dominantmesh was generated with mesh 9- was generated step refinementwith towards 9-step refinementthe aerofoil. towards The element the aerofoil. size was The approximately element size was reduced approximately from reduced 1 m to 2 mm.from The entire 1 m to surface 2 mm. Theof the entire aerofoil surface was of covered the aerofoil with was3 boundary covered layers. with 3 The boundary layers. total number Theof cells total was number about of 1.7 cells million. was about The 1.7y+ million.on the aerofoil The y+ onwas the confirmed aerofoil was to be confirmed to be between 30 andbetween 300 to 30let andthe 300wall to functions let the wall properly functions work. properly The generated work. The 3D generated CFD mesh 3D CFD mesh for profile S5HLfor profileis shown S5HL in various is shown view in variousangles in view Figure angles 3. The in Figureanalysis3. Thewas analysiscarried out was carried out using the incompressibleusing the incompressible stationary flow stationary solver simpleFoam flow solver provided simpleFoam by OpenFOAM. provided byOpenFOAM.

(a) (b)

(c)

Figure 3. 3DFigure mesh generated 3. 3D mesh around generated profile around S5HL. profile (a) CFD S5HL. mesh (a including) CFD mesh far including field. (b) Detailfar field. view (b) of Detail CFD mesh around the wing. (c)view Channel of CFD between mesh the around wing the main wing. part (c and) Channel the flap. between the wing main part and the flap.

2.3. Analysis of2.3. Lift Analysis and Drag of Lift and Drag

Lift coefficientsLift cl coefficientscalculated forcl calculated the high-lift for profiles the high-lift are plotted profiles as are functions plotted of as the functions of the angle of attackangle (AoA) of in attack Figure (AoA) 4’s left in Figurehand side.4’s left In handgeneral, side. the In c general,l increases the withcl increases the AoA. with the AoA. At a certain angle,At a certainstall occurs angle, and stall the occurs cl suddenly and the drops.cl suddenly As the drops.LE is lowered As the LE from is lowered S4HL from S4HL to S5HL, the stallto S5HL, is delayed the stall to ishigher delayed stall to angles. higher The stall maximum angles. The values maximum of lift coefficient values of lift coefficient cl,max and stall canglesl,max and of stallthe high-lif anglest of profiles the high-lift are listed profiles in Table are listed 1. By in changing Table1. Bythe changing profile the profile from S4HL tofrom S5HL, S4HL cl,max to is S5HL, increasedcl,max byis increased14% from by2.80 14% to 3.06. from The 2.80 stall to 3.06. angle The is stallalso angle is also increased fromincreased 14.2 to 16.4 from degrees. 14.2 to 16.4 degrees. Appl. Sci. 2021, 11, 2752 6 of 28 Appl. Sci. 2021, 11, x FOR PEER REVIEW 6 of 29

(a) (b)

Figure 4. CalculatedFigure 4.liftCalculated (a) and drag lift ( ba)) andcoefficients drag (b )for coefﬁcients high-lift profiles. for high-lift proﬁles.

Table 1. List of maximumTable 1. List lift of coefficients maximum and lift coefﬁcientsstall angles. and stall angles.

Profiles ProﬁlesBLHL BLHLS4HL S4HL050HL 050HLS5HL S5HL Max Lift Coeff.Max Lift Coeff. 2.80 2.80 2.69 2.69 2.91 2.91 3.06 3.06 Stall Angle/°Stall Angle/◦ 14.2 14.2 13.2 13.2 15.0 15.0 16.4 16.4

The stall can also be found in the plots of the drag coefficient cd in Figure 4’s right The stall can also be found in the plots of the drag coefﬁcient cd in Figure4’s right hand side; these suddenly increase at the respective stall angles. The difference in cd hand side; these suddenly increase at the respective stall angles. The difference in cd among d among the profilesthe proﬁles before beforethe stall the is stall minor. is minor. For the For AoA the AoA larger larger than than 4 degrees, 4 degrees, c iscd re-is reduced as the d duced as the LELE is is lowered. lowered. For For the the AoA AoA between between 0 0 and and 4 degrees, cd of S5HL is slightly higher than cd higher than cd ofof S4HL S4HL (2.4% (2.4% at AoA = 2 2 degrees). degrees).

2.4. Analysis of 2.4.Pressure, Analysis Flow, of Pressure,and cp Distributions Flow, andc p Distributions To estimate theTo largest estimate aerodynamic the largest load aerodynamic on the profiles, load on extreme the profiles, flight extremeconditions flight conditions were first definedwere based first definedon the load based factor on the Nz: load For factorcruise Nz:profiles For cruisewith retracted profiles withtrailing retracted trailing edge, three conditionsedge, three for conditionsNz = 2.5 (climb) for Nz and = 2.5 one (climb) for Nz and = 1 one(level for flight), Nz = 1 for (level high-lift flight), for high-lift profiles with extendedprofiles withtrailing extended edge, three trailing conditions edge, three for conditionsNz = 2.0 (climb). for Nz These = 2.0 (climb).are listed These are listed in Table 2. in Table2.

Table 2. List of definedTable 2. extremeList of deﬁned conditions extreme both conditionsfor cruise and both high-lift for cruise profiles. and high-lift proﬁles.

For Cruise ProfilesFor Cruise Proﬁles For High-Lift Profiles For High-Lift Proﬁles Conditions Conditions I II IIII IIIV III I’ IV II’ I0 III’ II0 III0 Air Speed/(m/s)Air Speed/(m/s) 130.6 153.7 130.6 192.1 153.7184.4 192.1 72.0 184.482.3 72.092.6 82.3 92.6 Angle of Attack/°Angle of Attack/ 14.4 ◦ 9.7 14.4 5.2 9.7 1.0 5.2 10.0 1.0 6.0 10.0 3.1 6.0 3.1

In these extremeIn conditions, these extreme 2D CFD conditions, analyses 2D were CFD carried analyses out. wereThe simulated carried out. pres- The simulated sure and flow pressure(left hand and side) flow and (left cphand distribution side) and orc pdimensionlessdistribution or pressure dimensionless normalized pressure normalized by dynamic pressureby dynamic (right pressure hand side) (right around hand cruise side) aroundprofile S5 cruise are shown profile in S5 Figure are shown 5. in Figure5. Those around high-liftThose around profiles high-lift S5HL are profiles shown S5HL in Figure are shown 6. The maximum in Figure6 .suction The maximum on the suction on upper side of fourthe upperprofiles side S4, ofS5, four S4HL, profiles and S5HL S4, S5, represented S4HL, and by S5HL cp and represented the actual bypres-cp and the actual sure in kPa arepressure listed in in Table kPa are3. These listed indata Table were3. These used datafor examining were used required for examining strength, required strength, stiffness, and actuationstiffness, forces and actuation for the morphi forces forng system the morphing being developed, system being see developed, Section 4.2. see Section 4.2. From these pressure distributions, the following conclusions were drawn. Suction on the upper side near the LE is increased along with AoA. Namely, the largest suction occurs in condition I for the cruise profiles and in condition I0 for the high-lift profiles. The suction peak is located at the leading edge where the highest velocity occurs due to the deflection of the flow direction induced by the aerofoil. At level flight or in condition IV, a large suction peak occurs just under the stagnation point on the lower side. This is especially prominent for profile S5HL having the fully lowered LE. This is again because of Appl. Sci. 2021, 11, 2752 7 of 28

large acceleration of the ﬂow turning around this point. On the other hand, the positive cp Appl. Sci. 2021, 11, x FOR PEER REVIEW 7 of 29 peak occurs at the stagnation point. The value of this peak, which is nearly 1, implying dynamic pressure, is independent of the AoA.

(a) Condition I

(b) Condition II

(d) Condition IV

Figure 5. Pressure and flow (left) and cp distribution (right) for cruise profile S5 in four extreme Figure 5. Pressure and ﬂow (left) and cp distribution (right) for cruise proﬁle S5 in four extreme conditions.conditions. Appl.Appl. Sci. Sci. 20212021, 11,,11 x ,FOR 2752 PEER REVIEW 8 of8 29 of 28

(a) Condition I’

(b) Condition II’

FigureFigure 6. Pressure 6. Pressure and and flow ﬂow (left (left) and) and cpc distributionp distribution (right (right) for) for high-lift high-lift profile proﬁle S5HL S5HL in in three extremeex- tremeconditions. conditions.

TableTable 3. List 3. List of ofsuction suction peak peak values values in inextreme extreme conditions. conditions.

Proﬁle Condition Max Pressure Coefﬁcient cp MaxMax Suction Suction (kPa) Profile Condition Max Pressure Coefficient cp I −4.53 (kPa)−21.2 I II −4.532.51 −−21.216.3 S4 II III −2.511.21 −−16.312.3 S4 IV −0.82 −7.7 III −1.21 −12.3 IV I −0.823.57 −7.716.7 II −2.22 −14.4 S5 I −3.57 −16.7 III −1.25 −12.7 II −2.22 −−14.4 S5 IV 2.62 24.5 III I0 −1.256.48 −−12.720.7 S4HL IV II0 −2.623.79 −−24.515.8 I’ III0 −6.482.45 −−20.712.9 S4HL II’ I0 −3.794.82 −−15.815.4 S5HL III’ II0 −2.453.14 −−12.913.1 0 I’ III −4.822.24 −−15.411.8 S5HL II’ −3.14 −13.1 III’ −2.24 −11.8 Appl. Sci. 2021, 11, 2752 9 of 28

3. Design of the Skin and the Stringers 3.1. Material Selection and Design of the Skin A crucial task in designing a morphing leading-edge system is the design of the skin including material selection and stringer layout. On the one hand, the skin needs to provide high deformability in order to be able to morph into the positions S4 (cruise) and S5 (high- lift), see Figure2. On the other hand, the skin has to meet competing requirements regarding strength and stiffness. As shown in Section 2.4, the leading edge is the region, where the highest aerodynamic loads occur, compared to other regions of the wing. Furthermore, the contour-length of S5 is 7.7% longer than that of S4. This elongation cannot be realised by elastic straining of conventional composite materials (or metals), especially since cyclic deformation has to be endured over the service life. Thus, an alternative solution using a sliding contact is selected in order to address the length difference, see Section6. It is positioned at the lower side of the MLE so that C2-continuity at the upper surface of the MLE is not affected. The contact position is carefully designed so that it is in the region of the favourable pressure gradient where flow separation is less likely to occur, and thus does not deteriorate the aerodynamic performance. Morphing from S4 to S5 was therefore expected to be realised by pure bending of the skin and without any membrane strain. Under these conditions, one surface of the skin is compressed (negative strain) while the other is stretched (positive strain), and the absolute values of strain on both sides are equal so that the overall length in any direction does not change. The change of curvature of the MLE’s skin due to morphing is given by the two target shapes S4 and S5. The bending stiffness of the skin should be as low as possible to limit the forces needed for actuation and maintain a low overall weight of the system. At the same time, however, its bending stiffness must be sufficient to limit the deviation from the target shapes under the relevant aerodynamic pressure distributions computed by 2D-CFD, see Section 2.4. While this is largely achieved by means of span-wise stringers, the areas between the stringers still require a certain bending stiffness so as not to bulge in- or outward under the pressure distribution. In a first step towards material selection, a simplified finite element model of the skin consisting of shell elements and assuming isotropic material was therefore used to identify a suitable distribution of the bending stiffness along the contour. The limit for acceptable displacement of the skin perpendicular to its original surface due to aerodynamic loads was set to 1 mm. The thickness in the different regions of the skin was adjusted to meet this requirement and to approximate the target shapes as closely as possible. After the distribution of the bending stiffness along the contour had thus been determined, generic composite laminates (different ply- materials and lay-ups) were analysed by means of Classical Laminate Theory (CLT) and Puck’s failure criterion [30] to determine their respective bending stiffness and bending limit, i.e., the allowable curvature change with respect to avoiding failure. Material data (elastic properties and uniaxial strengths) for typical glass- and carbon-fibre reinforced plies (unidirectional and woven; the database gave values for specific combinations of fibres, resins, and fibre-volume-fractions) were taken from the database of a commercial CLT-software (ESACOMP). Different laminates (all laminates symmetric and balanced, fibre orientations between +/−5◦ and quasi-isotropic) and with different thicknesses (achieved by repetition of the basic symmetric and balanced laminates) were defined for each ply material and thus analysed. The failure criterion considers the stresses occurring in each ply of the laminate using a failure envelope defined by interpolation of the uniaxial strength values. It yields the inverse reserve factor (IRF), which assumes values equal to or greater than zero. If the IRF reaches or exceeds 1, failure of the ply is assumed. The investigations described above show that laminates made of unidirectional plies reinforced by high-strength or intermediate-modulus carbon fibres provide the necessary combination of properties. However, since the “typical” materials mentioned in the database usually do not directly correspond to commercially available materials, a material type with similar properties then had to be identified that can actually be obtained from suppliers. This led to the selection of IM/8552 carbon fibre [28] for the design of the MLE. This is a “prepreg” Appl. Sci. 2021, 11, 2752 10 of 28

material, i.e., the fibres of the semi-finished product stacked to make the laminate are already impregnated (pre-impregnated) by a specific amount of resin. Apart from being a unidirectional intermediate-modulus carbon fibre material which is certified for aviation use, its material properties were readily available in the literature. All these aspects were relevant with respect to its selection. By applying the CLT and Puck’s failure criterion as described above and using the available material properties, it could easily be verified that the achievable combinations of bending stiffnesses and strengths of laminates made of the selected material are suitable for the present application. Tailoring the stiffness of the MLE skin along the contour was achieved by suitable composite laminate lay-ups, varying the skin thickness by ply drop-offs in the laminate. These were determined in another iterative process based on finite element analysis (FEA) of the structure. Models consisted of linear shell elements (ABAQUS element type S4R, i.e., with reduced integration) with displacement boundary conditions imposed at the top edges of the stringers. The upper edge (approximately parallel to the span-wise direction) of the skin-contour, where it would be fastened to the wing’s front spar, was fixed in all directions (translational and rotational), and the aerodynamic pressure distribution for the respective flight condition (cruise or high-lift) was applied to the outer surface of the skin. The solutions were obtained using an implicit solver and considering geometric non-linearity. This way, the reaction forces occurring at the locations of the displacement boundary conditions are equivalent to the forces applied by an actuation system holding the skin in the required shape under the respective aerodynamic pressure load. Lay-ups were manually modified, and the resulting stresses due to morphing and aerodynamic pressure loads were analysed by the FEM. Subsequently, the IRF according to Puck’s failure criterion as well as the deviation from the target shapes were determined from the FE- solution. Based on these, further modifications of the lay-ups were made, and the process Appl. Sci. 2021, 11, x FOR PEER REVIEW repeated until a suitable solution had been found. A simplified flow-chart11 of of 29 the process

for deﬁning the skin laminate is shown in Figure7.

Figure 7. FigureSimplified 7. Simplified flow-chart flow-chart of the process of the process result resultinging in definition in definition of the of the skin skin laminate. laminate.

Tailoring is essential,Tailoring because is essential, the bending because strain the bending occurring strain in occurringthe composite in the plies composite on plies on the inside and outsidethe inside surfaces and outside of the surfaceslaminate of is the proportional laminate is proportional to the laminate to the laminatethickness thickness as as well as the localwell change as the localof curvature change of curvaturebetween betweenS4 and S4S5. and Assuming S5. Assuming that thatthe thetarget target shapes are approximated as intended, increasing the stiffness by adding plies thus also increases shapes are approximated as intended, increasing the stiffness by adding plies thus also the ply stresses, which may then lead to failure. The iterative process described above increases the ply stresses, which may then lead to failure. The iterative process described above considers the IRF (strength related; becomes more critical as thickness increases because surface strains increase) as well as the approximation of the target shapes (stiffness related; becomes less critical as thickness increases because aerodynamic loads cause less deviation from the target shapes) so that both aspects are addressed adequately. The objective of stiffness tailoring therefore was to identify the laminate providing just enough bending stiffness to withstand deformation under the aerodynamic loads while being as thin and compliant as possible to reduce actuation forces, weight, and surface-strains. Moreover, fibres should be oriented mainly in the direction of the bending strains to provide the required strength. Since span-wise bending stiffness is provided by the stringers (see Section 3.2), this led to a skin laminate with fibres mostly oriented along the contour, i.e., approximately parallel to the flight-direction. Of course, fibre orientation varies slightly between the plies to prevent splitting of the laminate. The skin of the MLE was divided into seven regions depending on the necessary stiffness and deformation capabilities. Apart from providing span-wise bending stiffness, the stringers are also used as attachment points for the actuation system [28].

3.2. Selection and Design of the Stringers The aerodynamic surface of the MLE is provided by the above-mentioned skin of CFRP. Its laminate was designed with comparatively low bending stiffness in the circumferential direction, i.e., along the cross-section contour, to allow the change of shape required for the morphing function. In the span-wise direction, however, considerable bending stiffness is required to limit the deformation of the contour due to the aerodynamic loads. Otherwise, the target shapes would not be met and the aerodynamic performance of the MLE would not be acceptable. Therefore, the skin was stiffened by additional span-wise stringers. Depending on how the stringers are joined to the skin, they could lead to a local stiffening of the skin and thus to uneven deformation of the MLE, which would also reduce the aerodynamic performance. Figure 8 schematically shows the leading-edge contours in the cruise and high-lift configurations (Figure 8 left) as well as the implied change of curvature and positioning of the stringers along the cross-sections’ contour (Figure 8 right). The stringer locations Appl. Sci. 2021, 11, 2752 11 of 28

considers the IRF (strength related; becomes more critical as thickness increases because surface strains increase) as well as the approximation of the target shapes (stiffness related; becomes less critical as thickness increases because aerodynamic loads cause less deviation from the target shapes) so that both aspects are addressed adequately. The objective of stiffness tailoring therefore was to identify the laminate providing just enough bending stiffness to withstand deformation under the aerodynamic loads while being as thin and compliant as possible to reduce actuation forces, weight, and surface-strains. Moreover, fibres should be oriented mainly in the direction of the bending strains to provide the required strength. Since span-wise bending stiffness is provided by the stringers (see Section 3.2), this led to a skin laminate with fibres mostly oriented along the contour, i.e., approximately parallel to the flight-direction. Of course, fibre orientation varies slightly between the plies to prevent splitting of the laminate. The skin of the MLE was divided into seven regions depending on the necessary stiffness and deformation capabilities. Apart from providing span-wise bending stiffness, the stringers are also used as attachment points for the actuation system [28].

3.2. Selection and Design of the Stringers The aerodynamic surface of the MLE is provided by the above-mentioned skin of CFRP. Its laminate was designed with comparatively low bending stiffness in the circumferential direction, i.e., along the cross-section contour, to allow the change of shape required for the morphing function. In the span-wise direction, however, considerable bending stiffness is required to limit the deformation of the contour due to the aerodynamic loads. Otherwise, the target shapes would not be met and the aerodynamic performance of the MLE would not be acceptable. Therefore, the skin was stiffened by additional span-wise stringers. Depending on how the stringers are joined to the skin, they could lead to a local stiffening of the skin and thus to uneven deformation of the MLE, which would also reduce the aerodynamic performance. Appl. Sci. 2021, 11, x FOR PEER REVIEW 12 of 29 Figure8 schematically shows the leading-edge contours in the cruise and high-lift conﬁgurations (Figure8 left) as well as the implied change of curvature and positioning of the stringers along the cross-sections’ contour (Figure8 right). The stringer locations were were defineddeﬁned where where the change the change in curvature in curvature is minimal, is minimal, limiting limiting the the effect effect of of any any local local stiffening stiffening theythey may may cause. cause.

Figure 8. CurvatureFigure analysis 8. Curvature for placement analysis offor stringers placemen (schematic);t of stringers (a (schematic);) abscissa: Location (a) abscissa: along Location the MLE-contour; along blue: Curvature alongthe contour MLE-contour; S4, red: Curvatureblue: Curvature along contouralong contour S5, grey: S4, Difference red: Curvature between along S4 andcontour S5 curvatures, S5, grey: Differ- grey squares: Selected locationsence for between stringers S4 2 and to 5 (regionS5 curvatures, of stringer grey 1 squares: does not Selected change curvaturelocations for very stringers much anyhow); 2 to 5 (region (b) Blue of curve is S4, red curve isstringer S5 (not to1 does scale). not change curvature very much anyhow); (b) Blue curve is S4, red curve is S5 (not to scale).

As shown in Figure 9, the stringers are formed by the inner part of the skin-laminate itself. While this is more complicated with respect to manufacturing than adhesive bonding of separate stringers, it avoids much of the local stiffening that would be caused by the bonding flanges.

Figure 9. Stringers formed from the inner part of the skin; the final decision on specific fillet material has not been made yet.

However, the skin’s laminate mostly contains fibres oriented approximately in the circumferential direction. To provide sufficient span-wise bending stiffness, the stringers Appl. Sci. 2021, 11, x FOR PEER REVIEW 12 of 29

were defined where the change in curvature is minimal, limiting the effect of any local stiffening they may cause.

Figure 8. Curvature analysis for placement of stringers (schematic); (a) abscissa: Location along the MLE-contour; blue: Curvature along contour S4, red: Curvature along contour S5, grey: Differ- ence between S4 and S5 curvatures, grey squares: Selected locations for stringers 2 to 5 (region of Appl. Sci. 2021, 11, 2752 12 of 28 stringer 1 does not change curvature very much anyhow); (b) Blue curve is S4, red curve is S5 (not to scale).

As shown in Figure9 9,, thethe stringersstringers areare formed formed by by the the inner inner part part of of the the skin-laminate skin-laminate itself.itself. WhileWhile thisthis isis moremore complicated complicated with with respect respect to to manufacturing manufacturing than than adhesive adhesive bonding bond- ofing separate of separate stringers, stringers, it avoids it avoids much much of the of localthe local stiffening stiffening that that would would be caused be caused by the by bondingthe bonding ﬂanges. flanges.

FigureFigure 9.9. StringersStringers formedformed from from the the inner inner part part of of the the skin; skin; the the final fina decisionl decision on on specific specific fillet fillet material material has has not beennot been made made yet. yet. However, the skin’s laminate mostly contains fibres oriented approximately in the circumferentialHowever, the direction. skin’s Tolaminate provide mostly sufficient contains span-wise fibres bendingoriented stiffness,approximately the stringers in the thereforecircumferential require direction. further reinforcement.To provide sufficient In principle, span-wise this bending could readily stiffness, be achievedthe stringers by adding CFRP-plies with fibres oriented along the stringers. However, the span-wise coefficients of thermal expansion (CTE) of skin and stringers would then be very different. In the temperature range considered for the MLE (i.e., down to −45 ◦C), this would lead to unacceptable transverse (i.e., span-wise) residual stresses in the skin. This is demonstrated by finite element analyses of the skin and CFRP-stringers in conjunction with applying Puck’s failure criterion (see Section 3.1). The skin consists of a symmetric and balanced laminate with fibres oriented mostly in the +5◦ and −5◦ direction (relative to the flight direction). Its span-wise CTE is ca. 18 × 10−6 1/K. The reinforced CRFP-stringers assumed initially possessed the following lay-up: [−85◦, 85◦, −45◦, 45◦, ◦ ◦ ◦ ◦ 5 , −5 , −5 , 5 ]S (expressed in the same coordinate system as the skin laminate, i.e., 90◦ corresponds to the span-wise direction, which is longitudinal to the stringers and perpendicular to the flight direction). The corresponding CTE in the span-wise direction is only 1.15 × 10−6 1/K. As the temperature is reduced, the skin itself would contract in the span-wise direction much more than the stringers. Since it is connected to the stringers and its contraction would thus largely be prevented, residual stresses would occur in the span-wise direction, i.e., roughly perpendicular to its fibres. Evaluation of the above mentioned IRF shows that these residual stresses would be critical with respect to causing ply-failure, especially when additional stresses due to morphing also occur. The obvious solution to this problem was to form the stringers from the skin laminate, as mentioned above, but to reinforce them not by additional CFRP-layers, but using a material with a span-wise CTE matching that of the skin as closely as possible. Con- veniently, the span-wise CTE of the skin laminate is very close to that of certain steels (e.g., X10CrNi18-8). Reinforcing the stringers by steel parts therefore reduced thermal residual stresses to acceptable levels. This is shown in Figure 10 for a preliminary version of the MLE that has since been superseded by later iterations of the design (the number of stringers was changed in this process). The current design of the stringers and the reinforcing steel-profiles is shown in Figure 11. Since the material (steel) and therefore the Appl. Sci. 2021, 11, x FOR PEER REVIEW 13 of 29

therefore require further reinforcement. In principle, this could readily be achieved by adding CFRP-plies with fibres oriented along the stringers. However, the span-wise coefficients of thermal expansion (CTE) of skin and stringers would then be very different. In the temperature range considered for the MLE (i.e., down to −45 °C), this would lead to unacceptable transverse (i.e., span-wise) residual stresses in the skin. This is demonstrated by finite element analyses of the skin and CFRP-stringers in conjunction with applying Puck’s failure criterion (see Section 3.1). The skin consists of a symmetric and balanced laminate with fibres oriented mostly in the +5° and −5° direction (relative to the flight direction). Its span-wise CTE is ca. 18 × 10-6 1/K. The reinforced CRFP- stringers assumed initially possessed the following lay-up: [−85°, 85°, −45°, 45°, 5°, −5°, −5°, 5°]𝑆 (expressed in the same coordinate system as the skin laminate, i.e., 90° corresponds to the span-wise direction, which is longitudinal to the stringers and perpendicular to the flight direction). The corresponding CTE in the span-wise direction is only 1.15 × 10−6 1/K. As the temperature is reduced, the skin itself would contract in the span-wise direction much more than the stringers. Since it is connected to the stringers and its contraction would thus largely be prevented, residual stresses would occur in the span-wise direction, i.e., roughly perpendicular to its fibres. Evaluation of the above mentioned IRF shows that these residual stresses would be critical with respect to causing ply-failure, especially when additional stresses due to morphing also occur. The obvious solution to this problem was to form the stringers from the skin laminate, as mentioned above, but to reinforce them not by additional CFRP-layers, but using a material with a span-wise CTE matching that of the skin as closely as possible. Conven- iently, the span-wise CTE of the skin laminate is very close to that of certain steels (e.g., X10CrNi18-8). Reinforcing the stringers by steel parts therefore reduced thermal residual Appl. Sci. 2021, 11, 2752stresses to acceptable levels. This is shown in Figure 10 for a preliminary version of the 13 of 28 MLE that has since been superseded by later iterations of the design (the number of stringers was changed in this process). The current design of the stringers and the reinforcing steel-profiles is shown in Figure 11. Since the material (steel) and therefore the span-wise CTE of the span-wisecurrent stiffeners CTE of are the unchanged, current stiffeners there areis no unchanged, relevant effect there on is nothe relevant thermal effect on the residual stressesthermal by this. residual stresses by this.

◦ Figure 10. InverseFigure reserve 10. Inverse factors duereserve to thermal factors residual due to thermal stresses residual in the undeformed stresses in skin the forundeformed cooling to skin−45 forC; cool- (a) ABAQUS- model (linear shelling to elements) –45 °C; (a of) ABAQUS-model the CRFP-only stringers (linear shell largely elements) shows of inverse the CRFP-only reserve factor stringers (IRF)-values largely shows (even without morphing) of ca.inverse 0.35 reserve (green) factor and above; (IRF)-values (b) ANSYS-model (even withou (lineart morphing) shell elements—here of ca. 0.35 (green) displayed and above; with (b their) AN- speciﬁed thickness) showsSYS-model IRF-values (linear mostly shell around elements—here 0.1 and below displayed (also without with their morphing); specified upperthickness) right shows corner: IRF-values Design principle Appl. Sci. 2021, 11, x FOR PEER REVIEW 14 of 29 of the stringer-reinforcementmostly around 0.1 assumed and below for the(also ANSYS-model without morphing (this); has upper since right been corner: superseded Design byprinciple the design of the shown in Figure 11). stringer-reinforcement assumed for the ANSYS-model (this has since been superseded by the design shown in Figure 11).

Figure 11. SteelFigure profiles 11. Steel for additional profiles for stiffening additional of the stiffening stringers; of stringersthe stringers; 1, 2, 4,stringers and 5 are 1, reinforced2, 4, and 5 byare a rein- pair of C-section forced by a pair of C-section profiles, each, while higher loads at stringer 3 require box-section profiles, each, while higher loads at stringer 3 require box-section profiles; cylindrical parts are fittings for the kinematic profiles; cylindrical parts are fittings for the kinematic system; stringer 1 is not connected to the system; stringer 1 is not connected to the kinematic system and therefore does not possess fittings; the CFRP-portion of each kinematic system and therefore does not possess fittings; the CFRP-portion of each stringers (not stringers (notshown shown here) here) is is sandwiched sandwiched between between the the two two profiles. profiles.

4. Actuation 4.System Actuation System Several optionsSeveral exist optionsfor actuating exist forthe actuating MLE system. the MLE Besides system. classi Besidescal actuators classical like actuators like pneumatic, hydraulic,pneumatic, and hydraulic, electromechanical and electromechanical systems, which systems, are well whichestablished are wellfor air- established for craft applications,aircraft several applications, novel actuation several novel concepts actuation for morphing concepts forsystems morphing are discussed systems are discussed in the literaturein the like literature shape memory like shape alloys memory [31] alloysand piezoceramic [31] and piezoceramic actuators actuators [32]. For [ 32the]. For the MLE MLE system systempresented presented in this paper, in this an paper, electromechanical an electromechanical actuation actuation system systemhas been has se- been selected lected becausebecause their reliability their reliability for aircraft for aircraft applications applications can be can quantified be quantiﬁed based based on experi- on experience [33], ence [33], andand they they are areconsistent consistent with with a more a more electric electric aircraft aircraft approach, approach, which which is expected is expected to play a to play a majormajor role rolein future in future aircraft aircraft design design [34]. [34 ]. For deployingFor the deploying morphing the leading morphing edge, leading the stringers edge, the2, 3, stringers 4, and 5 2,are 3, positioned 4, and 5 are positioned by the drive mechanism shown in Figure 12. by the drive mechanism shown in Figure 12. The drive mechanism is driven by a linear electromechanical actuator. Its orientation in spanwise direction allows to design an actuation system that leaves approx. 50% of the cross section as free space for cables, tubes, and other systems to be placed in the leading edge of the wing. A bell crank deﬂects the linear movement of this actuator into the desired direction of movement for stringer 3, which is an oblique movement pointing downwards and to the front. Stringers 2, 3, and 4 are connected to a crank by links. Stringer 5 is connected to the support structure by a link. The geometry of these links and the crank

Figure 12. Kinematic system design of the morphing leading edge [28].

The drive mechanism is driven by a linear electromechanical actuator. Its orientation in spanwise direction allows to design an actuation system that leaves approx. 50% of the cross section as free space for cables, tubes, and other systems to be placed in the leading edge of the wing. A bell crank deflects the linear movement of this actuator into the desired direction of movement for stringer 3, which is an oblique movement pointing downwards and to the front. Stringers 2, 3, and 4 are connected to a crank by links. Stringer 5 is connected to the support structure by a link. The geometry of these links and the crank together with the stiffness properties of the skin define the relative position of the stringers, resulting in an aerodynamically feasible aerofoil for any actuator position. Appl. Sci. 2021, 11, x FOR PEER REVIEW 14 of 29

Figure 11. Steel profiles for additional stiffening of the stringers; stringers 1, 2, 4, and 5 are reinforced by a pair of C-section profiles, each, while higher loads at stringer 3 require box-section profiles; cylindrical parts are fittings for the kinematic system; stringer 1 is not connected to the kinematic system and therefore does not possess fittings; the CFRP-portion of each stringers (not shown here) is sandwiched between the two profiles.

4. Actuation System Several options exist for actuating the MLE system. Besides classical actuators like pneumatic, hydraulic, and electromechanical systems, which are well established for aircraft applications, several novel actuation concepts for morphing systems are discussed in the literature like shape memory alloys [31] and piezoceramic actuators [32]. For the

Appl. Sci. 2021, 11, 2752 MLE system presented in this paper, an electromechanical actuation14 system of 28 has been selected because their reliability for aircraft applications can be quantified based on experience [33], and they are consistent with a more electric aircraft approach, which is expected to play a major role in future aircraft design [34]. together with the stiffnessFor deploying properties the of morphing the skin deﬁne leading the relativeedge, th positione stringers of the2, 3, stringers, 4, and 5 are positioned resulting in an aerodynamicallyby the drive mechanism feasible shown aerofoil in for Figure any actuator12. position.

Appl. Sci. 2021, 11, x FOR PEER REVIEW 15 of 29

Figure 12. KinematicFigure system12. Kinematic design system of the morphing design of leadingthe morphing edge [ 28leading]. edge [28]. 4.1. Deﬁnition4.1. of an Definition Intermediate of an Baseline Intermediate Shape Baseline Shape The drive mechanism is driven by a linear electromechanical actuator. Its orientation Wheneverin amorphing spanwiseWhenever systemdirection a morphing is designed,allows system to design the is baseline designed, an actuation shape the needs baselinesystem to bethat shape deﬁned, leaves needs i.e.,approx. to be 50%defined, of the the manufacturingi.e.,cross the geometry, sectionmanufacturing as which free space isgeometry, stress for freecables, which as long tubes, is stress as an nod loadsfree other as are systemslong applied. as noto beloads placed are applied.in the leading An optionedge forAn the of option baselinethe wing. for shape the A bellbaseline is usingcrank shape onedeflects of is the using the target linear one shapes, of movement the target either ofshapes, the this aerofoil actuator either the into aerofoil the defor cruise (S4)for orsired forcruise take-off/landingdirection (S4) or offor movement take-off/landing (S5), see for Figure stringer (S5), 13 .3,see which Figure is an13. oblique movement pointing downwards and to the front. Stringers 2, 3, and 4 are connected to a crank by links. Stringer 5 is connected to the support structure by a link. The geometry of these links and the crank together with the stiffness properties of the skin define the relative position of the stringers, resulting in an aerodynamically feasible aerofoil for any actuator position.

Figure 13.FigureDerived 13. Derived intermediate intermediate baseline baseline together together with target with shapes target S4 shapes and S5. S4 and S5.

However, deformationHowever, from deformation S4 to S5 orfrom vice S4 versa to S5 leads or vice to large versa strains leads into thelarge skin. strains In in the skin. the context ofIn the the work context presented of the work here, thepresented material’s here, stress the material’s exposure isstress quantified exposure using is quantified us- the multiaxialing stress the failuremultiaxial criterion stress according failure criterion to Puck according [30]. According to Puck to [30]. this According criterion, to this crite- an inverse reserverion, factoran inverse (IRF) reserve is computed. factor (IRF) An IRF is computed. of 1 or higher An wouldIRF of indicate1 or higher static would indicate failure of the compositestatic failure structure. of the composite For cyclic structure. loading, For an IRFcyclic below loading, 0.5 is an acceptable. IRF below For 0.5 is acceptable. deformation from S4 to S5, an IRF of f = 0.752 (inter-fibre fracture with weakening) was For deformation from S4E1 to S5, an IRF of fE1 = 0.752 (inter-fibre fracture with weakening) computed bywas finite computed element (FE) by finite analysis, element which (FE) is criticallyanalysis, high.which is critically high. In order to reduce material strain and required actuation forces, an “intermediate baseline” between the geometries S4 and S5 was derived, see Figure 13. From this “intermediate baseline” the MLE can be moved up into aerofoil S4 for cruise or down to aerofoil S5 for take-off and landing. This intermediate baseline needs to be compatible with the movement of the kinematic system. Hence, the intermediate baseline was derived from a simulated intermediate shape of the skin resulting from an intermediate position of the drive mechanism. Using the intermediate baseline, the material stress exposure in terms of IRF is reduced by 38% to a value of fE1 = 0.4633, see Figure 14.

Appl. Sci. 2021, 11, 2752 15 of 28

In order to reduce material strain and required actuation forces, an “intermediate baseline” between the geometries S4 and S5 was derived, see Figure 13. From this “intermediate baseline” the MLE can be moved up into aerofoil S4 for cruise or down to aerofoil S5 for take-off and landing. This intermediate baseline needs to be compatible with the movement of the kinematic system. Hence, the intermediate baseline was derived from a simulated intermediate shape of the skin resulting from an intermediate position of the Appl. Sci. 2021, 11, x FOR PEER REVIEWdrive mechanism. 16 of 29 Appl. Sci. 2021, 11, x FOR PEER REVIEW 16 of 29 Using the intermediate baseline, the material stress exposure in terms of IRF is reduced by 38% to a value of fE1 = 0.4633, see Figure 14.

FigureFigure 14. 14. DistributionDistribution of inverse of inverse rese reserverve factor factor in inthe the skin skin of ofthe the morphing morphing leading leading edge. edge. Figure 14. Distribution of inverse reserve factor in the skin of the morphing leading edge. c t TheThe model model parameters parameters for for calculation calculation of ofthe the IRF IRF are are s s= =m m = =0.5, 0.5, 𝑝p⊥k, ,𝑝p⊥k= 0.35= 0.35 and and The model parameters for calculation of the IRF are s = m = 0.5, ∥𝑝 , ∥𝑝 = 0.35 and 𝑝 p=c 𝑝= ,p t𝑝 ,, p𝑝c , p𝑝t ,andpc 𝑝and s pandt sm and are mthe are weakening the weakening parameters parameters∥ expressing∥ expressing re- 𝑝⊥⊥= 𝑝 ⊥⊥, ∥𝑝 , ⊥k∥𝑝 , ⊥k𝑝 and⊥⊥ 𝑝 s ⊥⊥and m are the weakening parameters expressing reduction of inter∥ fibre∥ fracture (IFF) strength due to longitudinal stress. The material ductionreduction of ofinter inter fibre ﬁbre fracture fracture (IFF) (IFF) strength strength due due to to longitudinal longitudinal stress. stress. The The material material strengthstrength parameters parameters used used can can be befound found in in[35]. [35 ]. strength parameters used can be found in [35]. 4.2.4.2. Finite Finite Element Element Modeling Modeling of the of the Actuation Actuation System System 4.2. Finite Element Modeling of the Actuation System TheThe mechanical mechanical behaviour behaviour of ofthe the morphing morphing leading leading edge edge system system was was analysed analysed using using The mechanical behaviour of the morphing leading edge system was analysed using anan FE FE model. model. It includes It includes the the flexible ﬂexible composite composite structure structure of ofthe the skin skin as aswell well as asall all moving moving an FE model. It includes the flexible composite structure of the skin as well as all moving partsparts of ofthe the kinematic kinematic system, system, which which were were mo modelleddelled using using kinematic kinematic constraints. constraints. It Italso also parts of the kinematic system, which were modelled using kinematic constraints. It also includesincludes the the sliding sliding contact contact which which closes closes the the gap gap at the at the bottom bottom of ofthe the MLE MLE (see (see Section Section 6).6 ). includes the sliding contact which closes the gap at the bottom of the MLE (see Section 6). TheThe FE FE model model is depicted is depicted in inFigure Figure 15. 15 It. represents It represents a MLE a MLE section section of of1.3 1.3 m mspan. span. The FE model is depicted in Figure 15. It represents a MLE section of 1.3 m span.

FigureFigure 15. 15. FEFE (finite (ﬁnite element) element) model model of ofthe the morphing morphing leading leading edge edge (MLE) (MLE) system system shown shown from from two two different different perspectives: perspectives: Figure 15. FE (finite element) model of the morphing leading edge (MLE) system shown from two different perspectives: (a) side view, (b) rear view. (a) side view, (b) rear view. (a) side view, (b) rear view. In Figure 15a, all components of the drive mechanism are highlighted in red, which In Figure 15a, all components of the drive mechanism are highlighted in red, which move in the x-z-plane. In Figure 15b, those components of the drive mechanism are move in the x-z-plane. In Figure 15b, those components of the drive mechanism are highlighted in red, which move in the plane of the bell crank and the linear actuator. The highlighted in red, which move in the plane of the bell crank and the linear actuator. The hinges with a fixed position are supposed to be connected to the support structure. All hinges with a fixed position are supposed to be connected to the support structure. All parts of the drive mechanism are modeled as rigid parts, whereas skin and stringers are parts of the drive mechanism are modeled as rigid parts, whereas skin and stringers are deformable with the local material properties and ply stack sequence assigned. The deformable with the local material properties and ply stack sequence assigned. The actuator is represented by a “translator” feature, changing its length in order to move the actuator is represented by a “translator” feature, changing its length in order to move the MLE into the desired position. This movement drives the bell crank on the left hand side MLE into the desired position. This movement drives the bell crank on the left hand side Appl. Sci. 2021, 11, 2752 16 of 28

In Figure 15a, all components of the drive mechanism are highlighted in red, which move in the x-z-plane. In Figure 15b, those components of the drive mechanism are highlighted in red, which move in the plane of the bell crank and the linear actuator. The hinges with a fixed position are supposed to be connected to the support structure. All parts of the drive mechanism are modeled as rigid parts, whereas skin and stringers are deformable with the local material properties and ply stack sequence assigned. The actuator is represented by a “translator” feature, changing its length in order to move the MLE into the desired position. This movement drives the bell crank on the left hand side of Figure 15b. A connection rod between the two bell cranks makes them move synchronously. The sliding contact at the rear bottom end of the leading edge was modeled using frictionless perpendicular behaviour. Abaqus was used as FE software for this model. Skin and stringers were discretised using quadratic quadrilateral shell elements (S8). The bending stiffness of the shell elements representing the stringers are set equal to the bending stiffness of the stringer profiles described in Section 3.2. This allows us to model the behaviour of the stringers without modeling details of their geometry. Stringers and skin are modelled as one part in the FE model. This part has been meshed with an average element size of 10 mm × 10 mm. In total, the mesh contains 22,976 Elements and 66,320 nodes. A static general procedure was used with an implicit solver, considering geometric nonlinearity. In this FE model, the aerodynamic pressure loads for different flight conditions, which have been computed via CFD analyses (see Section 2.4), are applied to the skin surface. In Table4, these different load conditions are listed.

Table 4. Load conditions (steps) analysed in FE model.

No. 0 1 2 3 4 5 6 Geometry Manu. BL S4 S5 S5 S4 S4 Angle of Attack/◦ 0 0 0 0 14.41 14.41 1 True Air speed/(m/s) 0 0 0 0 130.6 130.6 184.4

• The initial condition “0” corresponds to the unloaded MLE in the stress-free manufacturing shape, which is depicted in Figure 15. • Step no. 1 corresponds to the installation of the MLE to the wing. While installing the MLE, the sliding surface for the gap closure is prestressed by pushing it up against the bottom wing contour (see Section6). For all subsequent steps, the rear bottom edge of the MLE slides on this contour. The prestressing of this sliding contact is maintained for all subsequent steps. • Steps no. 2 and 3 correspond to a test of the MLE on the ground without pressure loads. In step no. 2, the MLE is deformed into cruise aerofoil S4 without pressure loads. • In step no. 3, the MLE is deformed into the position for take-off and landing S5, still without pressure loads. • In step no. 4, pressure loads are applied assuming that now the plane takes off with MLE remaining in position S5 (Case S5 I in Table3). As shown in the results of CFD analyses (Section 2.4), the highest pressure loads at the MLE occur for the flight conditions with retracted flap. This may partly be due to the fact that a higher AoA is allowed with retracted flaps. In Section 2.4, it is also shown that the highest pressure loads at the MLE occur at the highest allowed AoA. Hence, the flight condition with the highest AoA, the highest true air speed at that AoA and with retracted flaps is used as the most critical case in position S5. • In step no. 5, the MLE is morphed to position S4, whereas the air speed and the AoA remain constant (Case S4 I in Table3). The results shown in Table3 show that the pressure load on the MLE is even higher for S4 than for S5 in that flight condition. Even though it is not intended to use aerofoil S4 at high AoA, this flight condition is included into the load cycle, because it is the most critical one. Not dimensioning for Appl. Sci. 2021, 11, 2752 17 of 28

it would mean adding constraints to the acceptable flight envelope depending on the MLE position. • Case S5 IV (aerofoil S5 with low AoA and high true air speed) in Table3 is not considered as a critical load case. The computed suction peak value is the highest one, but it occurs at the bottom, whereas the highest suction for all other cases occurs on the upper surface of the MLE. Suction at the bottom for aerofoil S5 reduces forces on the actuation system because the aerodynamic pressure loads act in the same direction as the deformation enforced by the actuation system. • In step no. 6, aerofoil S4 is maintained, but the pressure loads are modified in order to correspond to cruise condition with reduced AoA and increased air speed. • A load cycle corresponding to one flight consists of steps no. 1 to 6 and then repeating the load conditions 5, 4, and 3 in reversed order. Appl. Sci. 2021, 11, x FOR PEER REVIEW 18 of 29 For each of the last three steps, a pressure distribution was derived by CFD analyses presented in Section 2.4. In between the load cases, pressure loads have been interpolated linearly over time. Between each load case, the linear actuator moves with constant piston the actuation system because the aerodynamic pressure loads act in the same direc- velocity from one position to the next one. Hence, its stroke has also been interpolated tion as the deformation enforced by the actuation system. linearly over time. • In step no. 6, aerofoil S4 is maintained, but the pressure loads are modified in order The approximation of the target shapes is achieved with the proposed actuation to correspondsystem to cruise as computed condition with with the reduced FE model AoA shown and increased in Figure air 16 speed.. While the actuation system • A load cyclepositions corresponding the stringers to one exactly flight cons at theists desired of steps positions, no. 1 to 6 the and desired then repeating curvature of the aerofoil the load conditionscontour in5, 4, between and 3 in the reversed stringers order. can only be met approximately by tailoring the local For each ofstiffness the last three properties steps, ofa pressure the skin. distribution The effect ofwas the derived remaining by CFD deviation analyses on aerodynamic presented in Sectionperformance 2.4. In between is assessed the load using cases, 2D CFD pressure analysis. loads The have results been are interpolated depicted in Figure 17. linearly over time. BetweenFor aerofoil each S5load in highcase, liftthe configuration,linear actuator the moves lift coefficient with constant of the piston achieved approxima- velocity from onetion position is 0.2–0.8% to the higher next than one. the Hence, lift coefficient its stroke of has the also target been aerofoil, interpolated whereas the difference linearly over time.in drag is in the range between −1% and +0.4%. For aerofoil S4, the comparison is made The approximationfor the clean of the wing target configuration shapes is achieved (flaps retracted), with the becauseproposed it isactuation intended sys- for flight in cruise tem as computedcondition. with the Here,FE model larger shown relative in Figure deviations 16. While occur. the For actuation angles of system attack betweenpo- 0◦ and 6◦, sitions the stringerswhich exactly is typically at the the desired case for posi cruisetions, condition, the desired the curvature drag coefficient of the isaerofoil up to 1.3% lower than contour in betweenthe dragthe stringers coefficient can of only the targetbe met aerofoil, approximately whereas by the tailoring deviation the inlocal lift stiff- is in between −0.3% ness properties andof the +1.1% skin. inThe this effect interval. of the At remaining other angles deviation of attack, on aerodynamic the relative deviation perfor- is higher, but mance is assessedregarding using 2D typical CFD analysis. flight cycles, The theseresults conditions are depicted are ofin minorFigure relevance.17.

FigureFigure 16. 16. AchievedAchieved approximation approximation of of target target shapes shapes comp computeduted by by FE FE model model for for load load cases cases 4 4 and and 5. 5. Appl. Sci. 2021, 11, x FOR PEER REVIEW 19 of 29

Appl. Sci. 2021, 11, 2752 18 of 28 Appl. Sci. 2021, 11, x FOR PEER REVIEW 19 of 29

Figure 17. Effect of deviation between target aerofoil and achieved approximation on lift (a) and drag (b) coefficients.

For aerofoil S5 in high lift configuration, the lift coefficient of the achieved approximation is 0.2–0.8% higher than the lift coefficient of the target aerofoil, whereas the difference in drag is in the range between −1% and +0.4%. For aerofoil S4, the comparison is made for the clean wing configuration (flaps retracted), because it is intended for flight in cruise condition. Here, larger relative deviations occur. For angles of attack between 0° and 6°, which is typically the case for cruise condition, the drag coefficient is up to 1.3% lower than the drag coefficient of the target aerofoil, whereas the deviation in lift is in between −0.3% and +1.1% in this interval. At other angles of attack, the relative deviation Figure 17. Effect of deviation between target aerofoil and achieved approximation on lift (a) and Figure 17. Effect of deviationis higher, between but target regarding aerofoil andtypical achieved flight approximation cycles, these onconditions lift (a) and are drag of minor (b) coefﬁcients. relevance. drag (b) coefficients. The main purpose of the described FE model is the computation of the loads acting The main purpose of the described FE model is the computation of the loads acting on the components of the actuation system. The loads have been computed for all hinges For aerofoilon theS5 in components high lift con offiguration the actuation, the system.lift coefficient The loads of the have achieved been computed approxi- for all hinges shown in Figure 15. As an example, for the computed loads, the required actuation forces mation is 0.2–0.8%shown higher in Figure than 15the. Aslift ancoefficient example, of for the the target computed aerofoil loads,, whereas the required the differ- actuation forces for the different load conditions are depicted in Figure 18. ence in drag isfor in the the different range between load conditions −1% and are +0.4%. depicted For aerofoil in Figure S4 18, .the comparison is made for the clean wing configuration (flaps retracted), because it is intended for flight in cruise condition. Here, larger relative deviations occur. For angles of attack between 0° and 6°, which is typically the case for cruise condition, the drag coefficient is up to 1.3% lower than the drag coefficient of the target aerofoil, whereas the deviation in lift is in between −0.3% and +1.1% in this interval. At other angles of attack, the relative deviation is higher, but regarding typical flight cycles, these conditions are of minor relevance. The main purpose of the described FE model is the computation of the loads acting on the components of the actuation system. The loads have been computed for all hinges shown in Figure 15. As an example, for the computed loads, the required actuation forces for the different load conditions are depicted in Figure 18.

FigureFigure 18.18.Actuation Actuation forcesforces computedcomputed usingusing thethe describeddescribed FE FE model. model.

InIn thisthis figure, figure, it it is is shown shown that that the the highest highest actuation actuation force force (5.8 (5.8 kN) kN) is is required required in in step step 4 (S54 (S5 with with pressure pressure loads), loads even), even though though the the major major part part of of the the actuation actuation force force is is due due to to the the stiffnessstiffness ofof thethe MLEMLE (5.2(5.2 kNkN ininstep step3: 3: S5S5 withoutwithout pressurepressure loads)loads)and and notnot toto thethe pressurepressure loads.loads. TheThe requiredrequired powerpower ofof thethe actuatoractuator dependsdepends linearlylinearly onon thethe desireddesired actuationactuation speed.speed. Hence,Hence, thethe choicechoice ofof thethe velocityvelocity requirementrequirement hashas aa majormajor impactimpact onon thethe weightweight ofof thethe actuator.actuator. The The transition transition between between high high lift lift and and clean clean wing wing configuration configuration is typically is typically a slow a slow movement that may take several seconds. The actuation stroke for morphing from aerofoil S5 to S4 is 110 mm. Assuming an actuation speed of 22 mm/s (corresponding to a deployment time of 5 s) and an actuator efficiency of 90% the required power of the Figure 18. Actuation forces computed using the described FE model. electromechanical actuator is 142 W (without power reserve) for the MLE section of 1.3 m. The load cycles of all kinematic components have been computed by this FE model, and In this figure, it is shown that the highest actuation force (5.8 kN) is required in step the components have been dimensioned accordingly. 4 (S5 with pressure loads), even though the major part of the actuation force is due to the stiffness of the5. MLE Fatigue (5.2 TestskN in step 3: S5 without pressure loads) and not to the pressure loads. The required power of the actuator depends linearly on the desired actuation speed. The material selected for the MLE must withstand the bending stresses and deforma- Hence, the choice of the velocity requirement has a major impact on the weight of the tions during operation. A 4-point bending test was chosen to validate the material, since actuator. The transition between high lift and clean wing configuration is typically a slow Appl. Sci. 2021, 11, x FOR PEER REVIEW 20 of 29

movement that may take several seconds. The actuation stroke for morphing from aerofoil S5 to S4 is 110 mm. Assuming an actuation speed of 22 mm/s (corresponding to a deployment time of 5 s) and an actuator efficiency of 90% the required power of the electromechanical actuator is 142 W (without power reserve) for the MLE section of 1.3 m. The load cycles of all kinematic components have been computed by this FE model, and the components have been dimensioned accordingly.

5. Fatigue Tests Appl. Sci. 2021, 11, 2752 5. Fatigue Tests 19 of 28 The material selected for the MLE must withstand the bending stresses and defor- mations during operation. A 4-point bending test was chosen to validate the material, since the skin of the MLE is mainly subjected to bending. The 4-point bending test was thepreferred skin of theover MLE the is3-point mainly bending subjected test to becaus bending.e a Thegreater 4-point portion bending of the test specimen was preferred is sub- overjected the to 3-pointthe maximum bending bending test because stresses a this greater way. portion The most of severe the specimen IRF according is subjected to Puck’s to thefailure maximum criterion bending [30] occurring stresses thisin the way. MLE The was most determined severe IRF via according a FE model to Puck’s similar failure to the criterion [30] occurring in the MLE was determined via a FE model similar to the one one described in Section 4. The displacement of the upper supports necessary to repro- described in Section4. The displacement of the upper supports necessary to reproduce this duce this IRF in the fatigue test was then determined via FEA of the specimens including IRF in the fatigue test was then determined via FEA of the specimens including test setup. test setup. The number of load cycles was set to 102,000, corresponding to approx. 56,000 The number of load cycles was set to 102,000, corresponding to approx. 56,000 ﬂights. flights. The bending amplitude used in the 4-point bending tests corresponds to deformation The bending amplitude used in the 4-point bending tests corresponds to deformation from S4 to S5. However, due to the test setup, a pulsating load was realised, which is from S4 to S5. However, due to the test setup, a pulsating load was realised, which is conservative in comparison to the alternating loads in the MLE. The IRF reached in these conservative in comparison to the alternating loads in the MLE. The IRF reached in these bending tests are 70% higher than the ones expected for deformation from the intermediate bending tests are 70% higher than the ones expected for deformation from the intermedi- baselineate baseline shape shape into eitherinto either S4 or S4 S5 or so S5 that so athat sufﬁcient a sufficient safety safety margin margin is provided. is provided. ate baseline shape into either S4 or S5 so that a sufficient safety margin× is provided. SpecimensSpecimens werewere manufactured manufactured from from plates plates measuring measuring 50 50 cm cm ×50 50 cm, cm, which which were were laminatedSpecimens by hand were lay-up manufactured using HexPly from® platesIM7/8552 measuring prepreg 50 material cm × 50 by cm, Hexcel which® ,were see laminated by hand lay-up using HexPly® IM7/8552 prepreg material by Hexcel®, see Fig- Figurelaminated 19. by hand lay-up using HexPly IM7/8552 prepreg material by Hexcel , see Fig- ure 19.

FigureFigure 19. 19.Hand Hand lay-up lay-up of of the the laminate; laminate; after after curing curing in in a a hot-press, hot-press, specimens specimens are are cut cut from from the the plate plate by abrasive water-jet. byplate abrasive by abrasive water-jet. water-jet.

AfterAfter hand hand lay-up, lay-up, the the plates plates were were cured cured in in a a hot hot press press reaching reaching a a nominal nominal thickness thickness ofof about about 1 1 mm. mm. TheThe coupon coupon specimens specimens werewere then then cut cut from from the the plates plates using using an an abrasive abrasive water-jetwater-jet cutter. cutter. The The dimensions dimensions of of the the specimens specimens are are shown shown in in Figure Figure 20 20..

FigureFigure 20. 20.Specimen Specimen for for four four point point bending bending tests tests (all (all dimensions dimensions in in mm). mm).

Appl.Appl. Sci.Sci. 20212021,, 1111,, xx FORFOR PEERPEER REVIEWREVIEW 2121 ofof 2929

To perform the 4-point-bending fatigue tests with constant amplitudes, two different setups were used. The test setups differ in the distance between the lower supports, see Figures 21 and 22 to realise two different load cases, which generate slightly different IRFs ToTo performperform thethe 4-point-bending4-point-bending fatiguefatigue testtestss withwith constantconstant amplamplitudes,itudes, twotwo differentdifferent in the specimen. setupssetups werewere used.used. TheThe testtest sesetupstups differdiffer inin thethe distancedistance betweenbetween thethe lowerlower supports,supports, seesee Appl. Sci. 2021, 11, 2752 FiguresFigures 2121 andand 2222 toto realiserealise twotwo differentdifferent loadload cases,cases, whichwhich generategenerate slightlyslightly differentdifferent20 ofIRFsIRFs 28 inin thethe specimen.specimen.

Figure 21. Four point bending setup for test scenario 1 (all dimensions in mm).

Figure 21. Four point bending setup for test scenario 1 (all dimensions in mm). FigureFigure 21.21. FourFour pointpoint bendingbending setupsetup forfor testtest scenarioscenario 11 (all(all dimensionsdimensions inin mm).mm).

Figure 22. Four point bending setup for test scenario 2 (all dimensions in mm).

All tests were generally performed with 1 Hz test frequency. FigureFigure 22.22. FourFour pointpoint bendingbending setupsetup forfor testtest scenarioscenario 22 (all(all dimensionsdimensions inin mm).mm). FigureThe 22. Fouraim pointof the bending studies setup was forto testdemonstrate scenario 2 (allthat dimensions the samples in mm). survive the respective bending load at least 102,000 times. AllAll teststests werewere generallygenerally performperformeded withwith 11 HzHz testtest frequency.frequency. ForFor test test scenario scenario 1, 1, six six specimens specimens were were tested tested up up to to the the load load cycle cycle limit. limit. The The maximum maximum downwardTheThe aimaim displacement ofof thethe studiesstudies of the waswas upper toto demonstratedemonstrate supports was thatthat 54 thethe mm. samplessamples In Figure survivesurvive 23, the thethe test respectiverespective set up is downward displacement of the upper supports was 54 mm. In Figure 23, the test set up is shownbendingbending at loadload maximum atat leastleast displacement. 102,000102,000 times.times. shown at maximum displacement. ForFor testtest scenarioscenario 1,1, sixsix specimensspecimens werewere tetestedsted upup toto thethe loadload cyclecycle limit.limit. TheThe maximummaximum downwarddownward displacementdisplacement ofof thethe upperupper supportssupports waswas 5454 mm.mm. InIn FigureFigure 23,23, thethe testtest setset upup isis shownshown atat maximummaximum displacement.displacement.

FigureFigure 23. 23. SpecimenSpecimen in four point bending bending test test setup setup with with maximum maximum displacement displacement of ofthe the upper upper supports [28]. supports [28]. FigureFigure 23.23. SpecimenSpecimen inin fourfour pointpoint bendingbending testtest setupsetup withwith maximummaximum displacementdisplacement ofof thethe upperupper supportssupportsFor test[28].[28]. scenario 2, another six specimens were tested up to the load cycle limit. The maximum downward displacement of the upper supports was 52 mm (and thus 2 mm less than in scenario 1 because the distance between the supports is also different, see Figures 20 and 21). In summary, a service life of 102,000 cycles was demonstrated for both scenarios. Appl. Sci. 2021, 11, x FOR PEER REVIEW 22 of 29

For test scenario 2, another six specimens were tested up to the load cycle limit. The maximum downward displacement of the upper supports was 52 mm (and thus 2 mm Appl. Sci. 2021, 11, 2752 less than in scenario 1 because the distance between the supports is also different, see 21 of 28 Figures 20 and 21). In summary, a service life of 102,000 cycles was demonstrated for both scenarios.

6. Gap Closure6. Gap Closure The threeThe most three prominent most prominent configurations configurations of the MLE of the are MLE shown are shownin Figure in 24. Figure After 24 . After assemblyassembly of the skin of the and skin actuation and actuation system systemand mounting and mounting on the front on the spar front of sparthe wing, of the wing, the MLEthe is MLEin the isintermediate in the intermediate baseline baselineshape sh shapeown in shown grey. For in grey.cruise For flight, cruise the flight, skin is the skin tilted upwardsis tilted upwardsby the actuation by the actuation system to system assume to the assume contour the contourshown in shown blue in(S4). blue For (S4). For take-offtake-off and landing, and landing, the skin the is skintilted is downward tilted downward into the into high-lift the high-lift configuration configuration shown shown in in orangeorange (S5). (S5).

Figure 24.Figure Schematic 24. Schematic view of cruise-configuration view of cruise-configuration (S4, blue), (S4,intermediate blue), intermediate baseline shape baseline (grey), shape (grey), and high-liftand high-liftconfiguration configuration (S5, orange) (S5, of orange) the MLE; ofthe CFRP-stringers MLE; CFRP-stringers are each stiffened are each by stiffened a pair of by a pair of steel profilessteel with profiles box- with or C-shaped box- or C-shaped cross-sect cross-sections,ions, respectively—the respectively—the figure shows figure showstheir cross- their cross-section section includingincluding the the fittings fittings for for the the kinematic system (see(see FigureFigure 1111 forfor moremore detail); detail); the the rearward rearward extension extension of the MLE’s skin slides along the lower surface of the front spar. of the MLE’s skin slides along the lower surface of the front spar.

The deformationThe deformation from the from cruise- the to cruise- the high-lift to the high-lift configuration configuration not only not increases only increases the the wing profile’swing profile’s camber, camber, but also but its alsochord-length. its chord-length. Thus, and Thus, as mentioned and as mentioned in Section in Section3.1, 3.1, the contourthe contourof the high-lift of the configuration high-lift configuration is significantly is significantly longer than longer that of than the cruise that of con- the cruise figuration.configuration. To avoid excessive To avoid membrane excessive strain membranes in the skin, strains the in lower the skin, edge theof the lower leading- edge of the edge skinleading-edge is allowed to skin move is allowed forward to and move aft. forwardThe formation and aft. of Thea gap formation between ofit and a gap the between front spar,it and however, the front needs spar, to be however, avoided needs to prevent to be the avoided ingress to of prevent dirt and the foreign ingress objects, of dirt and and alsoforeign because objects, the airflow and also interacting because with the airflow it could interacting cause significant with it couldnoise causeemission. significant Therefore,noise the emission. CFRP-skin Therefore, is extended the CFRP-skinaft to overlap is extended the lower aft surface to overlap of the the front lower spar surface of (and thethe parts front of the spar wing-skin (and the attached parts of theto it wing-skin in this region). attached The laminate to it in this lay-up region). and Theshape laminate of the extensionlay-up and are shapeshape- of and the stiffness-tailored extension are shape- so that and installation stiffness-tailored of the skin so on that the installation front of spar resultsthe skin in a on bending the front deformation spar results and in a th bendingus a tensioning deformation of the and overlapping thus a tensioning part of the against overlappingthe contact surface. part against To reduce the contact wear, surface.a thin profile To reduce running wear, in athe thin span-wise profile running direc- in the tion andspan-wise made of an direction appropriate and madelow friction of an appropriatematerial is attached low friction to the material overlapping is attached part to the of the skin.overlapping When the part MLE of thetransitions skin. When from the on MLEe shape transitions to another, from this one slides shape along to another, the this lower surfaceslides alongof the thefront lower spar. surface of the front spar. At this point,At this the point, exact the amount exact amount of pre-tens of pre-tensionion necessary necessary to avoid to the avoid lossthe of contact loss of contact betweenbetween the skin-extension the skin-extension and the andfront the spar front in all spar situations, in all situations, e.g., also e.g., in the also presence in the presence of of local turbulence,local turbulence, has not hasyet been not yet analysed been analysed in detail. in For detail. the cruise For the and cruise high-lift and high-lift conditions conditions investigatedinvestigated so far, Figure so far, 25 Figure shows 25 theshows contact the contactforce per force metre per span metre at spanthe sliding at the contact sliding contact for the loadfor the cases load given cases in given Table in 4 Tablecomputed4 computed with the with FE themodel FE modeldescribed described in Section in Section 4.2. 4.2. This figure shows that the aerodynamic pressure distribution does not cause opening of the contact. As shown in Table4, the load cases 4, 5, and 6 are with different pressure fields, whereas the load cases 1, 2, and 3 represent cases without aerodynamic loads. The pre-tension reached with the current design is therefore sufficient. Whether this holds true for all possible flight conditions—e.g., also for negative angles of attack, which might cause suction on the wing’s lower surface—needs to be investigated in the future. Appl. Sci. 2021, 11, x FOR PEER REVIEW 23 of 29 Appl. Sci. 2021, 11, x FOR PEER REVIEW 23 of 29

Appl. Sci. 2021, 11, x FOR PEER REVIEW 23 of 29 Appl. Sci. 2021, 11, 2752 22 of 28 Contact force (N/m)force Contact Contact force (N/m)force Contact

Figure 25. Contact force per meter span at the sliding contact for the load cases given in Table 4 computedFigure 25. withContact the forceFE model per meter described span in at Sectionthe sliding 4.2. contact for the load cases given in Table 4 computed with the FE model described in Section 4.2. This figure shows that the aerodynamic pressure distribution does not cause opening of theThis contact. figure As shows shown that in theTable aerodynamic 4, the load prescasessure 4, 5, distribution and 6 are with does differentnot cause pressure opening fields,of the whereascontact. Asthe shown load cases in Table 1, 2, and4, the 3 representload cases cases 4, 5, andwithout 6 are aerodynamic with different loads. pressure The pre-tensionfields, whereas reached the load with cases the current 1, 2, and design 3 represent is therefore cases sufficient. without aerodynamicWhether this holdsloads. true The FigureFigureforpre-tension all 25. 25. possible ContactContact reached forceflight force withper perconditions—e.g., meter the meter current span span at atthedesignthe alsosliding sliding is for therefore contact negative contact for sufficient. forthe angles theload load casesof Whether casesattack, given given thisinwhich Table inholds Table might4 true4 computed with the FE model described in Section 4.2. computedcausefor all suction possible with theon flight FEthe model wing’s conditions—e.g., described lower surface— in Section alsoneeds 4.2 for. negative to be investigated angles of inattack, the future. which might causeThe suction local onbending the wing’s deformation lower surface— necessaryneeds to achieveto be investigated the pretension in the does future. not cause undueThisTheThe stresses figure local local bending showsbendingin the material,that deformation deformation the aerodynamic as shown necessary necessary in pressureFigures to to achieve achieve 26 distribution and the 27.the pretension pretension does not doescause does not openingnot cause cause ofundueundue the contact. stresses stresses As in in shown the the material, material, in Table as as shown4 shown, the load in in Figures Figurescases 4, 26 265 and, and 27 627.. are with different pressure fields, whereas the load cases 1, 2, and 3 represent cases without aerodynamic loads. The pre-tension reached with the current design is therefore sufficient. Whether this holds true for all possible flight conditions—e.g., also for negative angles of attack, which might cause suction on the wing’s lower surface—needs to be investigated in the future. The local bending deformation necessary to achieve the pretension does not cause undue stresses in the material, as shown in Figures 26 and 27.

FigureFigure 26.26. InverseInverse reserve reserve factors factors in in the the skin-extension skin-extension according according to Puck’s to Puck’s failure failure criterion criterion in in cruise-configurationFigure 26. Inverse reserve (S4); IRFs factors are invery the low. skin-extension according to Puck’s failure criterion in cruise-conﬁguration (S4); IRFs are very low. cruise-configuration (S4); IRFs are very low.

Figure 26. Inverse reserve factors in the skin-extension according to Puck’s failure criterion in cruise-configuration (S4); IRFs are very low.

FigureFigure 27. 27.Inverse Inverse reserve reserve factors factors in in the the skin-extension skin-extension according according to Puck’sto Puck’s failure failure criterion criterion in high-lift in conﬁgurationhigh-liftFigure 27. configuration Inverse (S5); IRFs reserve (S5); are higherfactors IRFs are thanin higherthe in skin-extension cruise-conﬁguration than in cruise-configuration according but to still Puck’s acceptable. but failurestill acceptable. criterion in high-lift configuration (S5); IRFs are higher than in cruise-configuration but still acceptable. 7. Mountability Concept

7.1.System Boundaries For the development of a mounting concept of the morphing leading edge, certain system boundaries need to be taken into account. One important requirement is the Figureconsideration 27. Inverse of reserve the cable factors routes in the which skin are-extension attached according to the front to Puck’s spar. failure Certain criterion repositioning in highof the-lift cable configuration routes can (S5); be IRFs considered, are higher butthan inin general,cruise-configuration the best accessible but still acceptable areas for. a joint between MLE assembly and the front spar are in the top and bottom regions of the front spar. Furthermore, the inaccessible areas for mounting the MLE need to be taken into account, see Figure 28. The backside of the front spar is not accessible due to the tank positioned in this area. The sides of the MLE are not accessible due to the neighbouring systems. As a conclusion, the assembly between skin structure and actuation system, Appl.Appl. Sci. Sci. 2021 2021, 11, 11, x, xFOR FOR PEER PEER REVIEW REVIEW 2424 of of 29 29

7.7. Mountability Mountability Concept Concept 7.1.7.1. System System Boundaries Boundaries ForFor the the development development of of a amounting mounting concep concept tof of the the morphing morphing leading leading edge, edge, certain certain systemsystem boundaries boundaries need need to to be be taken taken into into a ccount.account. One One important important requirement requirement is is the the con- con- siderationsideration of of the the cable cable routes routes which which are are attach attacheded to to the the front front spar. spar. Certain Certain repositioning repositioning ofof the the cable cable routes routes can can be be considered, considered, but but in in general, general, the the best best accessible accessible areas areas for for a ajoint joint betweenbetween MLE MLE assembly assembly and and the the front front spar spar are are in in the the top top and and bottom bottom regions regions of of the the front front spar.spar. Furthermore, Furthermore, the the inaccessible inaccessible areas areas for for mounting mounting the the MLE MLE need need to to be be taken taken into into ac- ac- Appl. Sci. 2021, 11, 2752 count,count, see see Figure Figure 28. 28. The The backside backside of of the the fron front spart spar is is not not accessible accessible due due to to the the tank tank23 posi- ofposi- 28 tionedtioned in in this this area. area. The The sides sides of of the the MLE MLE are are not not accessible accessible due due to to the the neighbouring neighbouring sys- sys- tems.tems. As As a aconclusion, conclusion, the the assembly assembly between between skin skin structure structure and and actuation actuation system, system, which which requiresrequires accessibility accessibility from from the the sides, sides, are are best best realised realised in in a apreassembly preassembly stage stage before before which requires accessibility from the sides, are best realised in a preassembly stage before mountingmounting to to the the front front spar. spar. The The preassembled preassembled system system can can then then be be attached attached as as a asingle single mounting to the front spar. The preassembled system can then be attached as a single item. item.item.

FigureFigure 28. 28. Space Space inaccessible inaccessible for for assembly. assembly.

7.2.7.2. Mounting Mounting Concept Concept AfterAfter the the review review ofof of thethe the systemsystem system boundaries,boundaries, boundaries, it it isit is clearis clear clear that that that the the the MLE MLE MLE is is bestis best best realized realized realized as as aas preassembly,a apreassembly, preassembly, which which which is is thenis then then attached attached attached to to theto the the front fr frontont spar spar spar in in in one one one piece.piece. piece. TheThe The areasareas areas mostmost most easily easily accessibleaccessible are are at at the the top top and and bottom bottom of of thethe the frontfront front spar.spar. spar. Therefore,Therefore, Therefore, thethe the interfacesinterfaces interfaces betweenbetween between MLEMLE and and front front spar spar are are best best positioned positioned there, there, see see Figure Figure 2929. 29.. Apart Apart from from accessibility, accessibility, thethe the proposedproposed design design with with its its two-point two-point support support offersoffe offersrs advantages, advantages, such such as as reduction reduction of of force force andand moment moment reaction reaction in in the the joint joint compared compared to to a a single-point single-point support, support, as as well well as as sufﬁcientsufficient sufficient spacespace for for cable cable routes routes between between the the two two joints. joints.

Figure 29. Positioning of interfaces between MLE and front spar.

In order to attach the MLE, joining elements attached to the front spar at the top and bottom are necessary, see Figure 30. These are best positioned in those areas, where ribs on the backside support the front spar and therefore sufﬁcient stiffness for load introduction into the front spar is provided. Structural analyses based on estimated loads show that eight M8 bolts are sufﬁcient to transmit the loads safely. After full preassembly, the complete MLE can be positioned on the front spar. A cut-out area in the skin on the top and bottom side allows access to the interfaces, see Figure 31. Once they are tightened, the bolts sit below the level of the skin surface. This prevents interference with the sliding movement of the skin extension (see Section6) on the bottom surface. The cut-out area can then be covered by a lid construction. Appl. Sci. 2021, 11, x FOR PEER REVIEW 25 of 29

Figure 29. Positioning of interfaces between MLE and front spar.

In order to attach the MLE, joining elements attached to the front spar at the top and Appl. Sci. 2021, 11, x FOR PEER REVIEW 25 of 29 bottom are necessary, see Figure 30. These are best positioned in those areas, where ribs on the backside support the front spar and therefore sufficient stiffness for load introduction into the front spar is provided. Structural analyses based on estimated loads show Figurethat eight 29. Positioning M8 bolts are of interfacessufficient between to transmit MLE theand loads front spar.safely.

In order to attach the MLE, joining elements attached to the front spar at the top and bottom are necessary, see Figure 30. These are best positioned in those areas, where ribs Appl. Sci. 2021, 11, 2752 on the backside support the front spar and therefore sufficient stiffness for load introduc-24 of 28 tion into the front spar is provided. Structural analyses based on estimated loads show that eight M8 bolts are sufficient to transmit the loads safely.

Figure 30. Joining elements for MLE attached to the front spar (left) and supporting stiffening elements on opposite side (right).

After full preassembly, the complete MLE can be positioned on the front spar. A cut- out area in the skin on the top and bottom side allows access to the interfaces, see Figure 31. Once they are tightened, the bolts sit below the level of the skin surface. This prevents FigureFigureinterference 30.30. JoiningJoining with elements elementsthe sliding for for MLEmovement MLE attached attached of to the tothe theskin front front extension spar spar (left (left) (see and) and Sectionsupporting supporting 6) on stiffening the stiffening bottom elements on opposite side (right). elementssurface. onThe opposite cut-out side area (right can). then be covered by a lid construction.

FigureFigure 31. 31.Bolted Bolted jointjoint betweenbetween MLEMLE and front spar (left), cut-out areas areas at at top top and and bottom bottom of of skin skin elementelement ( right(right)[) 28[28].].

8.8. BirdBird StrikeStrike KeepingKeeping an an eye eye on on potentialpotential futurefuture certiﬁcation certification of of a a morphing morphing leading leading edgeedge concept,concept,

oneone aspectaspect toto considerconsider is the proof of of bird bird strike strike resistance resistance [36]. [36]. In In particular, particular, it itmust must be beFigureshown shown 31. that Bolted that the the birdjoint bird doesbetween does not MLE notdamage damageand front the front thespar front( sparleft), sparcut-outof the of wing areas the wing andat top the andand fuel bottom the tank, fuel of as tank,skin this element (right) [28]. asmight this mightresult resultin catastrophic in catastrophic failure failureof the fuel of the system fuel system and subsequent and subsequent loss of lossthe aircraft of the aircraft[37–39].[ 37For–39 this]. Forwork, this the work, proof theof the proof bird of strike the birdresistance strike of resistance the morphing of the leading morphing edge 8. Bird Strike leadingwas demonstrated edge was demonstrated by means of bysimulation. means of For simulation. this purpose, For this simulati purpose,ons simulationsof a bird (mass of a2 bird kg)Keeping impacting (mass 2 an kg) theeye impacting MLE on potential at a the velocity MLE future atof certific a180 velocity m/sation were of of 180 conducteda morphing m/s were using le conductedading the edgeexplicit using concept, finite the explicitoneelement aspect finite software to elementconsider LS-DYNA. software is the Aproof LS-DYNA.summary of bird of Astrike the summary simulation resistance of the approach[36]. simulation In particular, and approachresults it mustis given and be resultsshownbelow. is thatFor given furtherthe below. bird details, does For not further the damage reader details, is the referred thefront reader spar to [40]. isof referred the wing to and [40]. the fuel tank, as this mightFollowing result in thecatastrophic industry standard,failure of the the fuel bird system was represented and subsequent by a cylinder loss of the with aircraft two semi-spherical[37–39]. For this caps work, and the modelled proof of using the bird smooth strike particle resistance hydrodynamics. of the morphing The leading bird model edge hadwas previouslydemonstrated been by validated means of against simulation. experimental For this purpose, data [41 ].simulati The compositeons of a bird skin (mass was modelled2 kg) impacting following the the MLE established at a velocity approach of 180 fromm/s were [42], consideringconducted using strength-based the explicit failure finite criteriaelement and software non-linear LS-DYNA. shear. A The summary internal of architecture the simulation of the approach MLE (actuators, and results etc.) is given was simplified.below. For Thefurther simulation details, resultsthe reader clearly is referred indicate to that [40]. the relatively thin composite skin is not capable of stopping the bird from penetrating into the wing. For some impact locations, the internal architecture is capable of stopping the bird before impacting the front spar. However, for the majority of cases, the bird was sliced by the internal architecture, and large portions of the bird subsequently impacted the front spar. Therefore, a bird strike protection concept had to be sought. Following recommendations from the literature [27,42], a bird splitter/deflector concept was integrated into the FE model consisting of thin aluminium plates. The simulation results indicated that a thickness of 2.5 mm was sufficient to prevent the bird from impacting the front spar. Figure 32 shows a typical simulation result for the Appl. Sci. 2021, 11, x FOR PEER REVIEW 26 of 29

Following the industry standard, the bird was represented by a cylinder with two semi-spherical caps and modelled using smooth particle hydrodynamics. The bird model had previously been validated against experimental data [41]. The composite skin was modelled following the established approach from [42], considering strength-based failure criteria and non-linear shear. The internal architecture of the MLE (actuators, etc.) was simplified. The simulation results clearly indicate that the relatively thin composite skin is not capable of stopping the bird from penetrating into the wing. For some impact locations, the internal architecture is capable of stopping the bird before impacting the front spar. However, for the majority of cases, the bird was sliced by the internal architecture, and large portions of the bird subsequently impacted the front spar. Therefore, a bird Appl. Sci. 2021, 11, 2752 strike protection concept had to be sought. Following recommendations from the 25litera- of 28 ture [27,42], a bird splitter/deflector concept was integrated into the FE model consisting of thin aluminium plates. The simulation results indicated that a thickness of 2.5 mm was sufficient to prevent the bird from impacting the front spar. Figure 32 shows a typical simulationcase including result abird for the strike case protection including shield. a bird strike For illustrative protection purposes, shield. For the illustrative skin has beenpur- poses,hidden the in skin order has to been increase hidden the visibilityin order to of in thecrease interaction the visibility of the of bird the withinteraction the internal of the birdstructure. with the internal structure.

Figure 32. Bird strike simulation. Left:Left: Impact Impact configuratio conﬁguration;n; right:right: Interaction Interaction of bird with internal structure.

9. Conclusions The morphing leading edge is an appropriateappropriate solutionsolution for increasingincreasing maximummaximum liftlift and stall stall angle angle of of an an aircraft. aircraft. In In contrast contrast to toclassical classical high high lift liftdevices devices like like Krueger Krueger flaps, flaps, the MLEthe MLE avoids avoids adding adding sources sources of ofturbulence turbulence on on the the upper upper wing wing surface. surface. Additionally, Additionally, it avoids gaps, which are typically sources of noise radiation. avoids gaps, which are typically sources of noise radiation. Because of high aerodynamic pressure loads on the leading edge, a structure is re- Because of high aerodynamic pressure loads on the leading edge, a structure is required, which is both stiff and deformable. These conflicting requirements are addressed quired, which is both stiff and deformable. These conflicting requirements are addressed by a CFRP composite structure with a ply stacking sequence optimised for bending, local by a CFRP composite structure with a ply stacking sequence optimised for bending, local stiffness variation by tailoring of the composite (Section3), and a pre-stressed sliding stiffness variation by tailoring of the composite (Section 3), and a pre-stressed sliding con- contact (Section6) at the bottom of the wing. Additionally, using an intermediate baseline tact (Section 6) at the bottom of the wing. Additionally, using an intermediate baseline shape as initial geometry reduces the required strainability of the material. Cyclic material shape as initial geometry reduces the required strainability of the material. Cyclic material tests and numerical analysis show that the proposed skin structure is able to withstand the tests and numerical analysis show that the proposed skin structure is able to withstand occurring strains for a full service life of an aircraft. the occurring strains for a full service life of an aircraft. The proposed stringer design and positioning combine a high stringer stiffness in the The proposed stringer design and positioning combine a high stringer stiffness in the spanwise direction with a minimised impact on deformation of the skin in the peripheral spanwisedirection. direction Furthermore, with the a minimised material composition impact on deformation of the stringers of the reduces skin in thermal the peripheral stresses, direction.as explained Furthermore, in Section 3.2the. material composition of the stringers reduces thermal stresses, as explainedThe designed in Section actuation 3.2. system for the MLE fits completely inside the leading-edge, leavingThe space designed for other actuation systems system and componentsfor the MLE like fits cablescompletely und tubes, inside which the leading-edge, are typically leavinglocated inspace that for region other of systems a wing. and The components required actuation like cables forces und allow tubes, for which an electromechani- are typically locatedcal actuation in that of region the MLE of a system.wing. The Together required with actuation the proposed forces allow mounting for an concept,electromechan- which icaldoes actuation not require of the any MLE ingress system. from Together difficult with to access the proposed areas behind mounting or at theconcept, sides which of the doesleading not edge, require the MLEany ingress system from is designed difficult for to minimizing access areas the behind impact or on at the the aircraft sides designof the leadingin order edge, to allow the integrationMLE system in is future designed versions for minimizing of existing the aircraft impact types. on the aircraft design in orderAs birdto allow strike integration is an important in future issue versions forthe of existing leading aircraft edges oftypes. any aircraft wing, a protection solution against bird strike has been developed. Numerical simulations show its ability to stop birds from hitting the front spar of the wing. Beyond the design progress presented in this paper, further work on design and qualification of the MLE system needs to be done, before such a system can be used in commercial flight operations. This includes lightweight optimization of the system, exhaustive analysis of functional and structural safety aspects, prototype testing, and certification. Furthermore, a feasible ice-protection system has to be integrated into the skin structure. A possible solution for such a system in a morphing CFRP structure is described in [6]. Further improvement of aerodynamic performance for morphing systems may be achieved using an integrated aerodynamic optimisation approach, which considers material limitations like the one reported in [43] together with structural optimisation algorithms, e.g., [44]. Appl. Sci. 2021, 11, 2752 26 of 28

Author Contributions: Following the individual contribution of the authors is provided: Conceptu- alization, C.C.A., V.L., and D.L. Methodology, V.L., D.L., and J.-D.W. Formal analysis, S.A. and M.M. Investigation, C.C.A. and V.L. Writing—original draft preparation, C.C.A. V.L., D.L., S.A., M.M., J.-D.W., and J.D. Writing—review and editing, C.C.A., V.L., D.L., and M.M. Visualization, C.C.A. and V.L. Supervision, C.C.A. Project administration, C.C.A. and M.M. All authors have read and agreed to the published version of the manuscript. Funding: This research was funded by the EC in the frame of JTI Clean Sky 2, grant number CS2-AlR-GAM-2016-2017 and CS2-AlR-GAM-2018-2019. Institutional Review Board Statement: Not applicable. Informed Consent Statement: Not applicable. Data Availability Statement: Restrictions apply to the availability of these data. Data was partly obtained from Airbus Defence & Space and are available from the authors with the permission of Airbus Defence & Space. Acknowledgments: The task performed by Fraunhofer was supported by the EC in the frame of JTI Clean Sky 2, which is gratefully acknowledged. We would also like to thank Airbus Defence & Space for the good cooperation within this project. Conﬂicts of Interest: The authors declare no conﬂict of interest.

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