Exploitation of a Multifunctional Twistable Wing Trailing-Edge for Performance Improvement of a Turboprop 90-Seats Regional Aircraft

aerospace

Article Exploitation of a Multifunctional Twistable Wing Trailing-Edge for Performance Improvement of a Turboprop 90-Seats Regional Aircraft

Francesco Rea 1,* , Francesco Amoroso 1, Rosario Pecora 1 and Frederic Moens 2 1 Department of Industrial Engineering (Aerospace Division), University of Naples “Federico II”, Via Claudio 21, 80125 Naples, Italy; [email protected] (F.A.); [email protected] (R.P.) 2 ONERA, The French Aerospace Lab, Aerodynamic Aeroelasticity and Acoustics Department, 92190 Meudon, France; [email protected] * Correspondence: [email protected]; Tel.: +39-081-768-3573

Received: 4 October 2018; Accepted: 9 November 2018; Published: 16 November 2018

Abstract: Modern transport aircraft wings have reached near-peak levels of energy-efficiency and there is still margin for further relevant improvements. A promising strategy for improving aircraft efficiency is to change the shape of the aircraft wing in flight in order to maximize its aerodynamic performance under all operative conditions. In the present work, this has been developed in the framework of the Clean Sky 2 (REG-IADP) European research project, where the authors focused on the design of a multifunctional twistable trailing-edge for a Natural Laminar Flow (NLF) wing. A multifunctional wing trailing-edge is used to improve aircraft performance during climb and off-design cruise conditions in response to variations in speed, altitude and other flight parameters. The investigation domain of the novel full-scale device covers 5.15 m along the wing span and the 10% of the local wing chord. Concerning the wing trailing-edge, the preliminary structural and kinematic design process of the actuation system is completely addressed: three rotary brushless motors (placed in root, central and tip sections) are required to activate the inner mechanisms enabling different trailing-edge morphing modes. The structural layout of the thin-walled closed-section composite trailing-edge represents a promising concept, meeting both the conflicting requirements of load-carrying capability and shape adaptivity. Actuation system performances and aeroelastic deformations, considering both operative aerodynamic and limit load conditions, prove the potential of the proposed structural concept to be energy efficient and lightweight for real aircraft implementation. Finally, the performance assessment of the outer natural laminar flow (NLF) wing retrofitted with the multifunctional trailing-edge is performed by high-fidelity aerodynamic analyses. For such an NLF wing, this device can improve airplane aerodynamic efficiency during high speed climb conditions.

Keywords: twistable trailing-edge; Natural Laminar Flow wing; actuator torque; instant centre analysis; regional aircraft

1. Introduction Worldwide air passenger traffic is predicted to grow at an average 4–5% per annum over the next few decades [1]. As the number of flights increases, environmental requirements, such as emissions and noise, will impose significant challenges for next generation transport aircraft development. According to Europe’s vision for aviation [1], technological breakthroughs are necessary to accomplish a major step towards the environmental goals of a 75% reduction in CO2 emissions per passenger/kilometre, a 90% cut in NOx emissions, a 65% reduction of perceived aircraft noise levels (all percentages referred to the transport aircraft performances measured in 2000).

Aerospace 2018, 5, 122; doi:10.3390/aerospace5040122 www.mdpi.com/journal/aerospace Aerospace 2018, 5, 122 2 of 23

Nowadays, modern transport aircraft wings have reached near-peak levels of energy-efficiency and further improvements seem extremely difficult to obtain. Indeed, aircraft wings are still designed with a fixed geometry fully optimized in only a few design points, which may not be so optimal for the entire flight mission. Therefore, whereas an aircraft operates in off-design conditions, sub-optimal performances lead to an increase of fuel burnt with impacts on air-pollution and aircraft operative costs. Morphing the shape of the aircraft wing during flight represents a very promising strategy to achieve some benefits throughout the entire aircraft mission [2]. During the very early days of aviation history, the possibility of changing the wing shape was considered a crucial design factor in order to generate lift and maintain lateral equilibrium. In 1903, the Wright Brothers achieved the first sustained, powered, heavier-than-air flight in a machine of their own design and construction [3]. After some experiments with kites and gliders, they developed a revolutionary wing design, enabling the lateral equilibrium of the aircraft: lateral control was indeed realized by twisting the rear of their fabric-and-wood wings in opposite directions [4]. Soon after, as aircraft became heavier and faster, engineers were forced to switch to stiff wings retrofitted with flaps and ailerons to satisfy the need for higher wing loading; morphing of these surfaces was proven to be impractical because of the higher structural stiffness required to withstand higher aerodynamic loads due to increased performances. Over the years, researchers and designers have conceived several promising concepts to enable shape-changing morphing devices. All wing morphing concepts can be categorized into three major types [5]: global plan form alteration (involving global aircraft characteristics such as span, chord and sweep changes), out-of-plane transformation (twist, dihedral, span-wise bending) and airfoil adjustment (camber and thickness). In each case, the design of smooth control-surface geometry variation must strike a balance between proper structural stiffness to withstand the external aerodynamic loads without appreciable deformations (or arising of aeroelastic instability issues) and sufficient flexibility to make the shape change possible with a reasonable amount of actuation power. In the mid 1980s, the Mission Adaptive Wing (MAW) research program demonstrated the concept of changing the wing in flight. The U.S. Air Force Research Laboratory and Boeing designed and tested adaptive wings that were installed on an F-111 aircraft [6]. Flight tests conducted on AFTI/F-111 aircraft confirmed the expected performance improvement: 20% range enhancement, 20% aerodynamic efficiency growth, 15% increase of wing air load at constant bending moment [6]. At the end of 1990s, the DaimlerChrysler Aerospace Airbus and DLR launched the research project ADIF (Adaptive Wing) and in this framework they proposed a concept to camber and twist the wing trailing edge [7]. The structural layout was conceived to be used for the replacement and enhancement of the flap trailing edge of the A340-300. Investigations of span-wise differential camber variation confirmed 12–15% reduction of the root bending moments (RBM) through the redistribution of the span-wise aerodynamic load. Within SARISTU—an industrial oriented research project in the frame of the 7th European framework Program (2011–2015)—different new concepts of morphing devices were developed and experimented in large wind tunnel tests (WT) at TsAGI facilities [8]. Different studies showed how controlled wing twisting can be an effective alternative for ailerons and that the amount of wing twist (washout) during the flight envelope can reduce he induced drag thus resulting in higher efficiency and less amount of fuel burn [9,10]. For this reason, different concepts were employed to induce wing twisting. Several works have been done on structural concepts for aerodynamic surfaces with adjustable torsional stiffness by means of rigid-body mechanisms. Indeed, the stiffness control for some selected components of the lifting surfaces was demonstrated to be a promising idea for the implementation of smart roll control [11] and adaptive lift-to-drag vertical tail ratio [12]. In such a case, the structural concept can change its own stiffness thanks to rotary wing spars with a controllably rigid attachment that permits aeroelastic amplification [12]. The Adaptive Torsion Wing (ATW) concept was based on the idea of a two-spar thin-walled closed section wing-box with all-movable spar webs in chordwise direction [13]. Aerospace 2018, 5, x FOR PEER REVIEW 3 of 24 Aerospace 2018, 5, 122 3 of 23 required to enable the predicted twisted shape [14]. In order to overcome the high-energy demand to control wing torsion and therefore to avoid the adoption of heavy actuators, a concept relying on These first concepts for inducing twist in a conventional wing structure were developed to reduce warping-induced deformations to an open-section airfoil was designed and tested [15]. Within the the torsion stiffness of the structure and, usually, high forces or moments were anyway required to EU FP7 CHANGE project, a similar concept was investigated for a 25 kg UAV [16]. A very interesting enable the predicted twisted shape [14]. In order to overcome the high-energy demand to control wing way to implement wing-twist morphing was finally addressed in Reference [17]. Here, the wing-twist torsion and therefore to avoid the adoption of heavy actuators, a concept relying on warping-induced was controlled by working on the bending-twist coupling induced by changes of shear centre deformations to an open-section airfoil was designed and tested [15]. Within the EU FP7 CHANGE location; the prototype, conceived for a glider of the FAI 15 m class, used smart material with project, a similar concept was investigated for a 25 kg UAV [16]. A very interesting way to implement controllable shear stiffness to adaptively modify the shear centre positions of wing cross sections. wing-twist morphing was finally addressed in Reference [17]. Here, the wing-twist was controlled by Most of the structural concepts conceived for wing twist are related to the trailing edge area working on the bending-twist coupling induced by changes of shear centre location; the prototype, where high benefits could be proved to be exploited on subsonic transport aircraft. Optimization of conceived for a glider of the FAI 15 m class, used smart material with controllable shear stiffness to wing trailing-edge shape could assure significant drag reduction within the flight envelope. With adaptively modify the shear centre positions of wing cross sections. respect to a civil transport aircraft configuration using conventional trailing-edge control surfaces, Most of the structural concepts conceived for wing twist are related to the trailing edge area where the benefits that might be brought by a morphing-camber system to the aircraft efficiency can high benefits could be proved to be exploited on subsonic transport aircraft. Optimization of wing approach more than 10 per cent, in off-design flight conditions and 1–3 per cent in cruise [18]. trailing-edge shape could assure significant drag reduction within the flight envelope. With respect to Estimated benefits for a reference transport aircraft (L-1011) prove the positive effects of variable a civil transport aircraft configuration using conventional trailing-edge control surfaces, the benefits camber system based on aileron-type trailing-edge surface deflections. that might be brought by a morphing-camber system to the aircraft efficiency can approach more than A competitive concept to enhance aircraft performances could be represented by a 10 per cent, in off-design flight conditions and 1–3 per cent in cruise [18]. Estimated benefits for a multifunctional trailing edge, retrofitting a Fowler flap, implementing wing camber-morphing reference transport aircraft (L-1011) prove the positive effects of variable camber system based on through rigid surface deflections (for lift-to-drag ratio improvement [18]) and continuous span-wise aileron-type trailing-edge surface deflections. twist control for root bending moments (RBM) alleviation (through the redistribution of the span- A competitive concept to enhance aircraft performances could be represented by a multifunctional wise aerodynamic load [19]). trailing edge, retrofitting a Fowler flap, implementing wing camber-morphing through rigid surface Reporting about the research activities developed in the framework of the Airgreen2 project deflections (for lift-to-drag ratio improvement [18]) and continuous span-wise twist control for (running within the “Clean Sky 2” Regional Integrated Development Platform), this paper is focused root bending moments (RBM) alleviation (through the redistribution of the span-wise aerodynamic on the preliminary design of a full-scale composite multifunctional and twistable trailing-edge load [19]). retrofitting the outboard morphing Fowler flap of a turboprop regional aircraft. The investigation Reporting about the research activities developed in the framework of the Airgreen2 project domain of the novel device (Figure 1) covers 5.15 m in wing span direction and the 10% of the wing (running within the “Clean Sky 2” Regional Integrated Development Platform), this paper is focused on local chord. The two functionalities of the flap tab (/trailing edge) device are activated when the the preliminary design of a full-scale composite multifunctional and twistable trailing-edge retrofitting fowler flap is stowed in the wing during cruise, climb and off-design flight conditions: the outboard morphing Fowler flap of a turboprop regional aircraft. The investigation domain of the novel Mode device A: Rigid (Figure deflections1) covers of 5.15 the m Fowler in wing flap span tip segment direction (from and the 90% 10% to of 100% the wing of the local local chord. wing Thechord) two functionalities within the angles of the range flap tab[+10°/ (/trailing−10°] (downwards/upwards), edge) device are activated when the fowler flap is stowed Mode in theB: “Continuous” wing during cruise,span-wise climb twist and with off-design a maximum flight conditions:differential twist-angle of 10° between the tip and root sections of the flap (up to ±5° at the root and tip sections respectively). • Mode A: Rigid deflections of the Fowler flap tip segment (from the 90% to 100% of the local wing Continuouschord) within span-wise the angles twist range is [+10achieved◦/−10 by◦] (downwards/upwards),elastic torsional deformation of the flap tab via distributed• Mode B:actuators. “Continuous” The span-wiseconceived twist structural with a maximumconcept differentialdeals withtwist-angle issues related of 10◦ betweento real implementationthe tip and on root large sections aircraft of theand flap is based (up to on± a5 ◦reasonableat the root number and tip sectionsof subcomponents respectively). integrating lightweight and energy efficient actuation systems.

(a) (b)

FigureFigure 1. 1. MultifunctionalMultifunctional Twistable Twistable Wing Wing TE: TE: ( (aa)) Outboard Outboard wing wing TE, TE, (/flap (/flap tab) tab) investigation investigation region region for for thethe structural structural concept; concept; ( (bb) )Explanation Explanation of of the the outboard outboard wing wing TE TE functionalities, functionalities, ( (upperupper)) rigid rigid deflection deflection ofof the the Fowler Fowler flap flap tip tip segment, segment, (lower (lower) continuous) continuous span-wise span-wise twist twist along along the the outboard outboard wing wing span. span.

Continuous span-wise twist is achieved by elastic torsional deformation of the ﬂap tab via distributed actuators. The conceived structural concept deals with issues related to real implementation

Aerospace 2018, 5, 122 4 of 23 on large aircraft and is based on a reasonable number of subcomponents integrating lightweight and energy efﬁcient actuation systems.

2. Aerodynamic Design of the Multifunctional Twistable Trailing-Edge The study presented in this section had the objective to define the optimal morphed shapes to be implemented for the wing trailing-edge (/flap tab) in order to improve the aircraft performance in high speed (climb) conditions. Retrofitting the segment of the AG2-NLF wing flap system (designed in cruise condition) with a multifunctional trailing-edge (/tab), the span load distribution could be optimized aiming at improve the global aircraft aerodynamic performance. The 3D computations of the aerodynamic flow around the airplane configurations have been carried out, referring to the ONERA elsA code [20]. This high fidelity CFD software solves the RANS equations on structured multi-block grids by a cell-centred finite volume technique. Spatial discretization uses the second order centred scheme of Jameson with 2nd and 4th order artificial dissipation. Convergence to steady flow solution is carried out thanks to a backward Euler technique with robust LU-SSOR implicit scheme method. The convergence is accelerated by the use of multigrid techniques for steady flows. Different turbulence models are available in elsA and in this work the Spalart-Allmaras turbulence model was used with the QCR modification [21]. In the RANS computations, the ONERA elsA software has the capability to compute laminar flow regions and to determine the transition location, by using the so-called AHD compressible criterion for Tollmien-Schlichting instabilities [22] and the so-called C1 criterion for crossflow instabilities [23], within the iterative convergence process. The generation of the wing shapes with morphed elements is done through the use of a grid deformation technique that has been used also in SARISTU project ([24]). The surface grid is firstly deformed according to the requested shape. Then, a displacement field of the grid nodes is derived for a volumetric transfinite deformation technic applied to the initial grid. The advantage of this method is that the same scripts can be used for the different computations, as the topological information is kept. The drawback is that it is based on the initial topology and some local grid inversion can be found if deformation is too large.

Trailing-Edge Aerodynamic Performance For the AG2-NLF regional airplane, multifunctional twistable trailing-edge could help to recover the laminar extent by an adaptation of pressure gradient in off-design condition. Considering the CL related to high speed climb condition, free transition computations show that laminar flow on the upper surface starts to be lost. Figure2 presents the different configurations considered for the multifunctional twistable trailing-edge: a rigid trailing-edge deflection (mode A) equal to 2.5◦ is presented in Figure2a, while a discretely increasing distribution of deflection angles along the span (mode B, angles 4◦/3◦/2◦) is sketched in Figure2b. Figure3 compares the computed Lift over Drag ratio (L/D) evolution versus CL in climb conditions for both the mentioned configurations. The L/D for the baseline configuration (no morphing) is also reported in Figure3. As expected, the efficiency of the trailing-edge morphing concept for the AG2-NLF airplane is visible for high CL, where an increase of about 0.4 in Lift over Drag ratio is found. More in detail, according to Figure3, the increase of Lift-over-Drag ratio (L/D) is up to +2% considering the CL evolution at climb condition (M∞ = 0.36 at 4372 m). This value resulted fully compliant with the industrial expectations in terms of benefits brought by the new technology; although marginal at aircraft level, the 2% increase of L/D was indeed considered very relevant at fleet level in force of the very positive impacts on large scale operations. Aerospace 2018, 5, x FOR PEER REVIEW 5 of 24 Aerospace 2018, 5, 122 5 of 23 Aerospace 2018, 5, x FOR PEER REVIEW 5 of 24

(a) (a)

(b)

FigureFigure 2. 2.ConfigurationsConﬁgurations Configurations for for thethe the multifunctionalmultifunctional multifunctionalTE: TE:( a((a))) Trailing-edge Trailing-edgeTrailing-edge rigidrigid rigid rotationrotation rotation (Mode(Mode (Mode A) A) equal equal to to2.5 to2.5°;◦;( 2.5°;b )(b Continuous )(b Continuous) Continuous increase increase increase in in deﬂection in deflection deflection angles anglesangles (Mode (Mode B) B)B) from fromfrom tip tip tip to to rootto root root (4 (4°/3°/2°).◦ (4°/3°/2°)./3◦/2◦).

Figure 3. Performance of the multifunctional trailing-edge. FigureFigure 3. PerformancePerformance of of the the multifunctional multifunctional trailing-edge.

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3. Concept Description Aerospace 2018, 5, x FOR PEER REVIEW 6 of 24 The concept of multifunctional twistable wing trailing-edge investigated in the present work is based3. Concept on the ideaDescription of retrofitting the tab of an outer board Fowler flap to enable new functionalities during cruiseThe concept and climb of multifunctional flight conditions. twistable The investigationwing trailing-edge domain investigated is presented in the in present Figure1 work and theis summarybased on of the main idea geometric of retrofitting data inthe Table tab of1. an outer board Fowler flap to enable new functionalities duringDuring cruise off-design and climb cruise flight flight conditions. conditions, The investigation the TE can rigidly domain rotate is presented around in its Figure main 1 hinge and the axis, duringsummary the climb of main phase, geometric continuous data in span-wise Table 1. twist can be enabled as explained in Figure1; in both cases theDuring goal isoff-design always to cruise enhance flight the conditions, aerodynamic the TE efficiency can rigidly of rotate the wing around and its get main consequent hinge axis, fuel savings.during The the climb multifunctional phase, continuous twistable span-wise trailing-edge twist can concept be enabled is a thin-walled as explained closed in Figure section 1; in whoseboth functionalitiescases the goal are is enabledalways to thanks enhance to the aerodynami actuation torquec efficiency provided of the by wing three and brushless get consequent rotary motor,fuel properlysavings. amplified The multifunctional by harmonic twistable drive geartrailing-edge units and concept inner mechanisms.is a thin-walled As closed rotary section actuators whose are activated,functionalities the inner are enabled mechanisms thanks can to the transfer actuation torque torque to theprovided structural by three concept brushless thus rotary providing motor, the requiredproperly performance. amplified by Upon harmonic the actuation drive gear of theunit actives and ribs,inner the mechanisms. Fowler flap As tab rotary is put actuators in movement are thusactivated, changing the the inner external mechanisms shape of can the transfer trailing torq edgeue (Figuresto the structural1 and4); ifconcept the shape thus changeproviding of eachthe ribrequired is prevented performance. by locking Upon the actuationthe actuation system, of the the acti compositeve ribs, the flap Fowler tab isflap elastically tab is put stable in movement under the actionthus of changing external the aerodynamic external shape loads. of the The trailing rigid edge rotation (Figures of the1 and wing 4); if TE the (Figure shape change1b, upper) of each can rib be is prevented by locking the actuation system, the composite flap tab is elastically stable under the obtained synchronizing the three actuators (R1, R2, R3). Conversely, “continuous” span-wise elastic action of external aerodynamic loads. The rigid rotation of the wing TE (Figure 1b, upper) can be twist of the flap trailing-edge (Figure1b, lower) can be activated by providing differential actuation obtained synchronizing the three actuators (R1, R2, R3). Conversely, “continuous” span-wise elastic control. For example, linear span-wise tab twist can be enabled providing a clockwise rotation to the twist of the flap trailing-edge (Figure 1b, lower) can be activated by providing differential actuation tip actuator (R3) and an anti-clockwise rotation to the root actuator (R1) while locking the central control. For example, linear span-wise tab twist can be enabled providing a clockwise rotation to the actuatortip actuator (R2).Fast (R3) and and reliable an anti-clockwise analytical androtation numerical to the methodsroot actuator in combination (R1) while locking with rational the central design criteria,actuator were (R2). implemented Fast and reliable to assess analytical the structural and numerical layout methods and actuation in combination system with with rational reference design to the mostcriteria, severe were load implemented condition expected to assess inthe service structural (limit layout load and condition); actuation AL2024-T5 system with alloy reference was used to the for themost great severe part of load the condition items of the expected inner mechanism,in service (limit while load 17-4PH condition); steel wasAL2024-T5 used for alloy the forkwas linkused of for the leveragethe great mechanism. part of the A items glass of fibre the prepreginner mechanism, with HexPly913 while 17-4PH from Hexcel steel was composites used for wasthe fork used link for of the skin,the “active” leverage ribs mechanism. and C-shape A glass spars fibre of prepreg the trailing withedge. HexPly913 from Hexcel composites was used for theThe skin, final “active” structural ribs and layout C-shape (Figure spars4) was of the analysed trailing by edge. means of an advanced finite element model which finallyThe final proved: structural layout (Figure 4) was analysed by means of an advanced finite element model which finally proved: • the capability of the actuation system to enable morphing through smooth rigid-body kinematic  of thethe innercapability mechanisms; of the actuation system to enable morphing through smooth rigid-body kinematic • theof absence the inner of mechanisms; any local plasticization and elastic instability at limit load condition for the items  madethe ofabsence aluminium of any andlocal steel plasticization alloy; and elastic instability at limit load condition for the items made of aluminium and steel alloy; • the strains for items made of GFRP to be lower than the maximum allowed strains at the limit  the strains for items made of GFRP to be lower than the maximum allowed strains at the limit load (both along the fibre direction and transversally with respect to the fibres); load (both along the fibre direction and transversally with respect to the fibres); • the absence of any failure up to the ultimate load condition (i.e., limit loads multiplied by a  the absence of any failure up to the ultimate load condition (i.e., limit loads multiplied by a contingencycontingency factor factor equal equal to to 1.5). 1.5).

FigureFigure 4. Multifunctional4. Multifunctional Twistable Twistable Wing Wing Trailing-edge Trailing-edgemounted mounted on the tip tip of of a a Fowler Fowler flap: ﬂap: 3D-CAD. 3D-CAD.

Aerospace 2018, 5, 122 7 of 23 Aerospace 2018, 5, x FOR PEER REVIEW 7 of 24 Aerodynamic Design Load Condition Aerodynamic Design Load Condition According to reference regulations ([25]), the structural design of any movable control surface on According to reference regulations ([25]), the structural design of any movable control surface a large airplane must comply with the following requirements: on a large airplane must comply with the following requirements: I. I. capabilitycapability to support to support limit limit loads loads without without permanent permanent detrimental detrimental deformation deformation and and deformation deformation levelslevels not notcompromising compromising safe safe operations operations (EASA (EASA CS CS25.305(a)); 25.305(a)); II. II. capabilitycapability to withstand to withstand ultimate ultimate loads loadswithout without failures failures in structural in structural components components or actuator or systems;actuator systems; III. III. clearanceclearance from from aeroelastic aeroelastic instability instability phenomena phenomena (EASA (EASA CS CS25.629). 25.629). Moreover, when when flying flying at at dive dive speed, speed, the the control control surface surface must must be able be ableto be to deflected be deflected by an byangle an equalangle to equal one tothird one of third the ofmaximum the maximum design design deflectio deflection.n. This condition This condition has been has considered been considered as limit as operativelimit operative configuration configuration for the forpreliminary the preliminary design of design the device, of the being, device, the being, highest the dynamic highest dynamicpressure occurringpressure occurring at the dive at the speed. dive speed.The limit The limitloads loads were were evaluated evaluated by bymeans means of ofan an in-house in-house code implementingimplementing a a 3D 3D Doublet Doublet Lattice Lattice Method Method (DLM); (DLM); the the adopted adopted aerodynamic aerodynamic model model is is depicted depicted in in Figure 55,, limitlimit resultantresultant loadsloads alongalong thethe outerouter wingwing trailingtrailing edge edge (/flap (/flap tab) havehave beenbeen summarizedsummarized inin Table1 1.. InIn thethe preliminarypreliminary designdesign phase,phase, thethe pressurepressure distributiondistribution waswas consideredconsidered asas uniformuniform onon flapflap tabtab upper upper and and lower lower external external surfaces surfaces (Figure (Figure 66a).a). Particular attentionattention waswas paidpaid toto the power required toto morph morph the the structure and to the consequent ac actuatorstuators size size and and weight; weight; in in summary, summary, the the entire entire preliminary design process of the system was driven by by the the need need of of simultaneously simultaneously meeting meeting different different requirementsrequirements (Figure 6 6b)b) inin orderorderto tocome come to toa asolution solutionof ofindustrial industrialrelevance. relevance.

(a) (b) Figure 5. Aerodynamic model used for limit load evaluation by means of DLM: (a) Aerodynamic Lattice Figure 5. 5. AerodynamicAerodynamic model model used used for for limit limit load load evaluation evaluation by by means means of of DLM: DLM: ( (aa)) Aerodynamic Aerodynamic Lattice Lattice of the reference regional aircraft TP90; (b) aerodynamic lattice of the outboard wing trailing-edge. of the the reference reference regional regional aircraft aircraft TP90; TP90; (b (b) )aerodynamic aerodynamic lattice lattice of of the the outboard outboard wing wing trailing-edge. trailing-edge.

(a) (b)

Figure 6. Design condition: ( (aa)) Pressure Pressure distribution distribution on on TE TE uppe upperr and and lower lower surfaces surfaces for for limit limit load condition;condition; ( (b)) Requirements Requirements simultaneously simultaneously fulfilled fulﬁlled to integrate a a multifunctional multifunctional TE TE at at aircraft aircraft level. level.

Aerospace 2018, 5, x FOR PEER REVIEW 8 of 24 Aerospace 2018, 5, 122 8 of 23 Table 1. Flap tab geometric data and Limit Load Condition considered for its design (loads obtained by means of DLM). Table 1. Flap tab geometric data and Limit Load Condition considered for its design (loads obtained by means of DLM). TE Root Chord 0.222 m TE root chord as % of local wing chord 10.57% TE tip chord TE Root Chord 0.222 0.167 m m TETE roottip chord asas % % of of local the winglocal chordwing chord 10.57% 7.770% TEOutboard tip chord Flap Span 0.167 5.015 m m TETE tiprigid chord deflection as % of the (downward) local wing chord 7.770% +5.00° Outboard Flap Span 5.015 m TETE rigidmean deflection geometric (downward) chord +5.00 0.236◦ m TEPressure mean geometric coefficient chord (upper), 𝐶, 0.2360.3247 m PressurePressure coefficient coefficient (upper), (lower),Cp, UP𝐶, 0.32470.5011 PressureDynamic coefficient pressure (lower), Cp, LOW 0.5011 12,005 Pa Dynamic pressure 12,005 Pa Dynamic pressure (upper), 𝑞∗𝐶, 3898 Pa Dynamic pressure (upper), q ∗ Cp, UP 3898 Pa 𝑞∗𝐶 DynamicDynamic pressure pressure (lower), (lower),q ∗ Cp, LOW, 60156015 Pa Pa ForceForce resultant resultant (upper), (upper),FUP 𝐹 23782378 N N Force resultant (lower), F 3668 N Force resultant (lower),LOW 𝐹 3668 N Total Hinge moment around TE hinge axis, MB3 475.8 N·m Total Hinge moment around TE hinge axis, 𝑀 475.8 N·m

4. Actuation System: Design Proc Processess and Estimated Performances The core element element of of an an adaptive adaptive structure structure is the is the actuation actuation system system including including its transmission its transmission line. line.Interactions Interactions between between the basic the elements basic elements of this ofmechanized this mechanized system systemand the andexternal the externalloads provide loads providefundamental fundamental insight into insight the behaviour into the behaviour of the over ofall the adaptive overall system. adaptive Power system. and Power weight and reductions weight reductionsare of paramount are of paramount importance importance to successfully to successfully integrate integrate adaptive adaptive systems systems in large in large airplanes airplanes for forimproving improving performances performances and and enlarge enlarge mission mission profiles. proﬁles. In In Figure Figure 7,7, the the flow-chart ﬂow-chart ofof thethe actuationactuation system design processprocess isis summarized.summarized.

Figure 7. Flow-chart of the actuation system design process.

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4.1. Kinematic Design of the Inner Mechanism The design engineer must ensure that the proposed mechanism will not fail under operating conditions. At the beginning of the design process, a tentative linkage has to be synthesized with the principal goal to provide the kinematic performances required by morphing operations; in a second stage, the obtained mechanism is properly investigated from the structural standpoint. The position of all the links or elements in the mechanism has to be evaluated for each increment of input motion and compared with the expected kinematic performance enabling the transition of the airfoil from its baseline configuration to the target morphed one. The tentative linkage was selected as the one that can be seen in Figure8a: a Watt’s Six-bar plus a Four-bar linkage. In such a way, the hollow-shaft rotary brushless motor can transfer rotation from the crank element (link 1 in Figure8a) to the output element (link 7 in Figure8a). As a first step, the lengths and positions of the links were defined as function of the input angle θ1 as the full linkage is a single degree of freedom (DOF) mechanism. Indeed, assuming trial link lengths, unknown link angles were evaluated and each link, represented as a position vector, was completely defined for each increment of input motion. The approach to linkage position analysis generates a vector loop (or several loops) around the linkage as first proposed by Raven [26]. In Figure8b, the links are represented as position vectors that form a vector loop. The lengths of the vectors are the link lengths, which are known. The choices of vector directions and sense, as indicated by their arrowheads, lead to this vector loop equation:

→ → → → r1 + rT0T1 − r2 − r3a = 0 → → → → r3b + rT1T2 − r5a − r4 = 0 (1) → → → → r5b + r6 − r7 − rT2H = 0

→ jθ where position vectors are deﬁned with complex number notation r i = ri·e i (with i = 1, . . . , 7). Each vector loop can be expressed as Freudenstein’s equation; if we solve for the angle θ3a, output of ﬁrst vector loop equation, we have:

KI · cos θ1 − KII · cos θ3a + KIII = cos(θ3a − θ1) (2)

2 2 2 2 rT0T1 rT0T1 r3a − r2 + r1 + rT0T1 KI = ; KII = ; KIII = ; (3) r3a r4 2·r3a·r1 Then, the ﬁrst vector loop, expressed as Freudenstein’s Equation (2), can be simpliﬁed as: √ ! −B ± B2 − 4·A·C θ = 2· tan−1 (4) 3a 2A where link lengths and known input angle θ1 terms have been collected as constants A, B and C:

A = KI · cos θ1 + KII + KIII + cos θ1; B = −2 cos θ1; C = KIII − KII − cos θ1 + KI · cos θ1 (5)

The full inner mechanism is made up of three four-bar linkages in series, as shown in Figure8b. These vector loop Equation (1) can be solved in succession with the results of the ﬁrst loop applied as input to the second loop. Note that there is a constant angular relationship between vectors r3a and r3b within ternary link 3. The solution for the four-bar linkage (4) is simply applied twice in the Watt’s Six-bar case: θ = f (r , r , r , r , θ ) 3a 1 2 3a T0T1 1 (6) θ5a = g(r3b, r4, r5a, rT1T2, θ3a) Aerospace 2018, 5, x FOR PEER REVIEW 9 of 24

𝜃 as the full linkage is a single degree of freedom (DOF) mechanism. Indeed, assuming trial link lengths, unknown link angles were evaluated and each link, represented as a position vector, was completely defined for each increment of input motion. The approach to linkage position analysis generates a vector loop (or several loops) around the linkage as first proposed by Raven [26]. In Figure 8b, the links are represented as position vectors that form a vector loop. The lengths of the vectors are the link lengths, which are known. The choices of vector directions and sense, as indicated by their arrowheads, lead to this vector loop equation:

𝑟⃗ 𝑟⃗ −𝑟⃗ −𝑟⃗ =0

𝑟⃗ 𝑟⃗ −𝑟⃗ −𝑟⃗ =0 (1)

𝑟⃗ 𝑟⃗ −𝑟⃗ −𝑟⃗ =0

where position vectors are defined with complex number notation 𝑟⃗ =𝑟 ∙𝑒 (with 𝑖=1,…,7). Each vector loop can be expressed as Freudenstein’s equation; if we solve for the angle 𝜃, output of first vector loop equation, we have:

𝐾 ∙ cos 𝜃 −𝐾 ∙ cos 𝜃 𝐾 = cos𝜃 −𝜃 (2)

𝐾 = ; 𝐾 = ; 𝐾 = ; (3) ∙∙ Then, the first vector loop, expressed as Freudenstein’s Equation (2), can be simplified as: Aerospace 2018, 5, 122 −𝐵 √𝐵 −4∙𝐴 ∙𝐶 10 of 23 𝜃 =2∙tan (4) 2𝐴 and onewhere more link time lengths for the and last known four-bar input linkageangle 𝜃 in terms Figure have6 beena: collected as constants A, B and C:

𝐴 =𝐾 ∙ cos 𝜃 𝐾 𝐾 cos 𝜃; 𝐵 = −2 cos 𝜃 ; 𝐶=𝐾 −𝐾 − cos 𝜃 𝐾 ∙ cos 𝜃 (5) The full inner mechanism is madeθ7 = upf (ofr5 threeb, r6, four-barr7, rT2H linkages, θ5a) in series, as shown in Figure 8b. (7)

Aerospace 2018, 5, x FOR PEER REVIEW 10 of 24

These vector loop Equation (1) can be solved in succession with the results of the first loop applied as input to the second loop. Note that there is a constant angular relationship between vectors 𝑟 and 𝑟 within ternary(a )link 3. The solution for the four-bar linkage (4) is simply(b) applied twice in the Watt’s Six-bar case: Figure 8. Kinematic model of the inner mechanism: (a) full leverage can be seen as six-bar linkage Figure 8. Kinematic model of the inner mechanism: (a) full leverage can be seen as six-bar linkage plus plus a four-bar linkage; (b) Vector loop𝜃 equations.= 𝑓𝑟,𝑟 ,𝑟,𝑟,𝜃 a four-bar linkage; (b) Vector loop equations. (6) 𝜃 =𝑔𝑟,𝑟,𝑟,𝑟,𝜃

The independentand one more time variable for the islastθ1 four-barwhich linkage will be in controlledFigure 6a: with the brushless motor. In such a way, each angle link was expressed as function𝜃 = of𝑓 the𝑟,𝑟 crank,𝑟,𝑟 angle,𝜃 θ1 (Figure9a), once the(7) link lengths were defined withinThe independent the minimum variable availableis 𝜃 which design will be spacecontrolled of thewithtip the trailing-edgebrushless motor. section In such (Figurea 9b). When theway, hollow-shaft each angle brushlesslink was expressed rotary as motor function is activated,of the crank theangle input 𝜃 (Figure rotation 9a), once is transferred the link to the crank elementlengths (link were 1).defined Leverage’s within the output minimum element available (link design 7) space must of providethe tip trailing-edge adequate section control action (Figure 9b). When the hollow-shaft brushless rotary motor is activated, the input rotation is during trailing-edge evolution from baseline position to the target shape. As shown in Figure9a, transferred to the crank element (link 1). Leverage’s output element (link 7) must provide adequate the final outputcontrol angleaction during curve trailing-edgeθ7(θ1) was evolution able to fulfill from base performanceline position angleto the target requirements shape. As shown within in the range ◦ ◦ [+10 /−10Figure] with 9a, thethe followingfinal output additionalangle curve criteria𝜃𝜃 was observed able to fulfill to assure performance effective angle trailing-edge requirements transition during morphingwithin the operations: range [+10°/−10°] with the following additional criteria observed to assure effective trailing-edge transition duringd morphingθ operations: 7 6= ∀ ∈ + ◦ − ◦ 0 θ7(θ1) [ 10 ; 10 ], (8) dθ 0 ∀ 𝜃𝜃 ∈ 10°; −10°, (8) 1 When conditionWhen condition (8) is verified,(8) is verified, inversion inversion points points (i.e., (i.e., toggle toggle positions positions of the oflinkage) the linkage) are avoided are avoided and smoothnessand smoothness transition transition from from the baselinethe baseline position position to to the the target target shape shape (and (andvice versa) vice can versa) occur. can occur.

(a)

(b)

Figure 9. KinematicFigure 9. Kinematic design design of the of innerthe inner mechanism: mechanism: (a (a) )Link Link Angles Angles as functi as functionon of the crank of the angle crank angle input rotation 𝜃; (b) Link lengths within the minimum available design space of the tip TE section. input rotation θ1;(b) Link lengths within the minimum available design space of the tip TE section.

Aerospace 2018, 5, 122 11 of 23

4.2. Inner Mechanism Mechanical Advantage Modern adaptive systems, designed for large demonstrators, are commonly based on the seamless integration of actuators, mechanisms and structures with the purpose of reshaping the external surface on demand. The reduction of power required to morph the structure and of the actuation system weight are of paramount importance to successfully integrate adaptive systems in large airplanes. For this reason, an energy-efficient approach must be adopted since the preliminary design phase of the actuation system. Parameters capable to express interactions between the basic elements of this mechanized system and the external loads have to be defined to provide insight into the behaviour of the overall adaptive architecture. The mechanical advantage of a mechanism could be defined as the ratio between the output and the input torque [27]. In order to cut down the capacities of electro- mechanical actuators, the mechanical advantage of the inner mechanism within the TE angle working range should be as high as possible. For the inner mechanism (IM) in Figure9b, the main output is the moment transferred around the trailing-edge hinge axis (MOUT) and the input refers to the torque applied to the crank (MIN). Assuming that the friction and inertia are neglectable, according to the principle of the virtual works, the following relationship can be found for the inner mechanism:

PIN = MIN · ωIN = MOUT · ωOUT = POUT (9) where ωIN is the crank angular speed and ωOUT is the angular speed around the trailing-edge hinge axis. According to the deﬁnition of instant center of rotation, at a given instant of time, a linkage mechanism undergoing planar movement has a point showing the same speed for both the input and output parts, thus the following relationship can be deﬁned:

ωIN·IC12 − IC2B3 = ωOUT·IC1B3 − IC2B3 (10) where IC12 is the instant centre between frame and input part (crank), IC1B3 is the instant centre between frame and output part, IC2B3 is the instant centre between input part and output part; IC12 − IC2B3 is the distance of instant centers IC12 and IC2B3, IC1B3 − IC2B3 is the distance of instant centres IC1B3 and IC2B3. Let’s now recall the Aronhold-Kennedy’s theorem which deals with the three instant centres between three links of a system of rigid members [28]:

Aronhold-Kennedy’s Theorem: The three instantaneous centres of three bodies moving relative to one another must lie along a straight line.

By returning to the inner mechanism obtained at the end of the kinematic design process (Figure9b) and applying this theorem, we can further simplify Equation (10) as follows:

ω IC − IC ω IC − IC IN · 26 12 = IN · 6B3 1B3 (11) ω6 IC26 − IC16 ωOUT IC6B3 − IC16 where IC6B3 is a first order instant centre and IC26 is a second order instant center. Therefore, the mechanical advantage of the inner mechanism can be defined as a function of particular first order and second order Instant Centres (ICs) of the linkage:

M ω IC − IC IC − IC M.A. = OUT = IN = 26 16 · 6B3 1B3 (12) MIN ωOUT IC26 − IC12 IC6B3 − IC16 where IC12, IC16 and IC1B3 are the primary instant centers which respectively coincide with leverage’s ﬁxed hinges (1, 2), (1, 6) and (1, B3). Construction of required instant centers for the inner mechanism, as shown in Figure 10a, is based on the intersection of proper Aronhold-Kennedy (AK) lines. All Aerospace 2018, 5, x FOR PEER REVIEW 12 of 24

Figure 10b) were evaluated by intersection of the respective Aronhold-Kennedy lines. In the second iteration, the second order ICs (2, 6) was finally evaluated and the mechanical advantage, for this specific linkage position, was obtained. When the crank is activated, the rib rotates from the baseline position to the maximum downward (upward) deflection equal to +10° (−10°). During the transition from the baseline position Aerospace 2018, 5, 122 12 of 23 to the target shape, for each intermediate linkage position, all required instant centres were estimated to completely obtain the mechanical advantage curve as function of the input crank rotation. requiredIn Figure ICs for 11a, the estimationthe evolution of the of mechanical each instant advantage centre position are summarized is reported. in the The IC matrixmechanical of the advantagelinkage (Figure for the 10 workingb). output angle range [+10°; −10°] is within the range [5.18; 8.97] (Figure 11b).

(a)

(b)

FigureFigure 10. 10. InnerInner mechanism mechanism instant instant centres centres analysis: analysis: ( (aa)) Construction Construction of of the the instant instant centres centres by by means means ofof AK lines; ( b)) IC IC matrix of of the linkage: main main hinges hinges (lab (labelledelled in in green), green), first ﬁrst order order ICs (labelled in in blue),blue), second second order IC (labelled in orange) and required ICs for MA estimationestimation (marked in red).

If the main hinges (labelled in green in Figure 10a) of the inner mechanism are defined as the output of the kinematic design process, in the first iteration the first order ICs (labelled in blue in Figure 10b) were evaluated by intersection of the respective Aronhold-Kennedy lines. In the second iteration, the second order ICs (2, 6) was finally evaluated and the mechanical advantage, for this specific linkage position, was obtained. When the crank is activated, the rib rotates from the baseline position to the maximum downward (upward) deflection equal to +10◦ (−10◦). During the transition from the baseline position to the target shape, for each intermediate linkage position, all required instant centres were estimated to completely obtain the mechanical advantage curve as function of the input crank rotation. In Figure 11a, the evolution of each instant centre position is reported. The mechanical advantage for the working output angle range [+10◦; −10◦] is within the range [5.18; 8.97] (Figure 11b). Aerospace 2018, 5, 122 13 of 23 Aerospace 2018, 5, x FOR PEER REVIEW 13 of 24

(a)

(b)

FigureFigure 11. 11.Inner Inner mechanism mechanism instant instant centrescentres analysis:analysis: (a (a) )Evolution Evolution of ofICs ICs positions positions for forseveral several IM IM conﬁgurations;configurations; (b) ( Mechanicalb) Mechanical Advantage Advantage and and outputoutput angle angle curve curve as as function function of crank of crank rotation rotation angle. angle.

4.3.4.3. Structural Structural Assessment Assessment of of the the Inner Inner Mechanism Mechanism To enableTo enable the transitionthe transition of the of Multifunctional the Multifunctio Twistablenal Twistable trailing-edge trailing-edge concept concept from from the reference the (baseline)reference shape (baseline) to the targetshape onesto the (Mode target Aones and (M B),ode three A and “active B), three ribs” “active were definedribs” were along defined the span-wisealong direction:the span-wise root, central direction: and root, tip sections. central and tip sections. Each active rib has the same inner mechanism, which was synthesized to be placed within the Each active rib has the same inner mechanism, which was synthesized to be placed within minimum design space available in the rib block 2 (B2) of the Fowler flap. Indeed, each single degree- the minimum design space available in the rib block 2 (B2) of the Fowler flap. Indeed, each single of-freedom (DOF) leverage is activated by a single brushless rotary motor. Therefore, the rigid degree-of-freedomrotation of the Fowler (DOF) flap leverage tab (Mode is activated A) can be by obtained a single brushlessby synchronizing rotary motor.the three Therefore, actuators. the On rigid rotationthe other of the hand, Fowler “continuous” flap tab (Mode span-wise A) can trailing be obtained edge twist by synchronizing (Mode B) can thebe threeactivated actuators. providing On the otherdifferent hand, “continuous”control actions. span-wise For example, trailing linear edge sp twistan-wise (Mode flap B)tab can twist be can activated be enabled providing providing different controlclockwise actions. rotation For example,to the tip actuator linear span-wiseand anti- clockwise flap tab rotation twist can to the be root enabled actuator. providing clockwise rotationFor to the this tip reason, actuator the andgeometric anti- clockwiseparameters rotation of the leverage to the root were actuator. synthesized in order to obtain almostFor this linear reason, relationship the geometric without oscillations parameters betw of theeen leveragethe crank wereangle synthesizedand the TE angle in order(Figure to 9a), obtain almostas well linear as relationshipa high mechanical without advantage. oscillations If the between first loop the design crank was angle mainly and the driven TE angleby kinematic (Figure 9a), as wellperformance as a high according mechanical to the advantage. minimum available If the first room loop within design the Fowler was mainly flap structure, driven the by second kinematic performanceloop was mainly according driven to the by minimumthe structural available sizing of room the withinlinks and the by Fowler the mechanical flap structure, arrangement the second loop was mainly driven by the structural sizing of the links and by the mechanical arrangement definition in compliance with limit aerodynamic loads. As the actuation system was the core of the Aerospace 2018, 5, 122 14 of 23

Aerospace 2018, 5, x FOR PEER REVIEW 14 of 24 adaptive structure, reliable and accurate finite-element model were defined to simulate its kinematics, verifydefinition the preliminary in compliance design with toollimit based aerodynamic on instant loads. centres As the (ICs) actuation as well system as to was prove the core its structural of the adaptive structure, reliable and accurate finite-element model were defined to simulate its integrity upon limit loads. A three-dimensional finite-element model was generated; it consisted kinematics, verify the preliminary design tool based on instant centres (ICs) as well as to prove its of six-faced solid elements (CHEXA [29]) for the links of the inner mechanisms and trailing edge structural integrity upon limit loads. A three-dimensional finite-element model was generated; it hinge fitting and beam elements (CBEAM [29]) coupled to rigid body elements (RBE2 [29]) for the consisted of six-faced solid elements (CHEXA [29]) for the links of the inner mechanisms and trailing cylindricaledge hinge hinges fitting and and related beam elements pins. For (CBEAM structural [29]) analysiscoupled to purpose, rigid body the elements implicit (RBE2 nonlinear [29]) for solver ® of MSC-NASTRANthe cylindrical hinges(SOL and 400)related was pins. used For to structural account analysis for nonlinear purpose, effects the implicit and large nonlinear displacements solver (rotations).of MSC-NASTRAN The capability® (SOL of 400) the was actuation used to system account to for enable nonlinear morphing effects through and large smooth displacements rigid-body kinematic(rotations). of the The inner capability mechanism of the actuation was verified system by to applying enable morphing enforced displacementsthrough smooth(SPCD rigid-body [29]) to thekinematic crank and of resistant the inner torque mechanism (equal was to verified 160 N·m) by alongapplying the enforc hingeed axis displacements of the actuator (SPCD hinge [29]) fitting. to Thethe magnitude crank and of resistant the enforced torque rotation(equal to was160 N·m) defined along on the the hinge base axis of a of preliminary the actuator iterative hinge fitting. analysis ◦ finalizedThe magnitude to get a specific of the enforced rotation rotation angle of was the trailingdefined edge.on the Abase crank of a rotation preliminary (θ1) equaliterative to −analysis35.3 was finalized to get a specific rotation angle of the trailing edge. A crank rotation◦ (𝜃 ) equal to −35.3° was found to be required to obtain a trailing-edge rotation (θ7) equal to +5 . foundVon Misesto be required stress distribution to obtain a trailing-edge (Figure 12b) rotation over the (𝜃 actuation) equal to mechanism +5°. confirms the absence of any localVon plasticization Mises stress at distribution limit load condition:(Figure 12b) maximum over the actuation peak stress mechanism is equal confirms to 484 MPa the andabsence located of any local plasticization at limit load condition: maximum peak stress is equal to 484 MPa and in the “fork” link made of 17-4PH steel which in turn has a Yield Strength higher than 700 MPa. All located in the “fork” link made of 17-4PH steel which in turn has a Yield Strength higher than 700 the other leverage’s items made of aluminium, show lower stress than the Yield Strength of Al2024-T5 MPa. All the other leverage’s items made of aluminium, show lower stress than the Yield Strength of alloy. Finally, the actuation torque required to hold the leverage in the final position resulted equal to Al2024-T5 alloy. Finally, the actuation torque required to hold the leverage in the final position 23.25resulted N·m: equal the mechanical to 23.25 N·m: advantage the mechanical resulted advantage equal to resulted 6.88 in equal accordance to 6.88 within accordance the outcomes with the of the instantoutcomes centre of tool the usedinstant in centre the preliminary tool used in design the preliminary stage. design stage.

(a)

(b)

FigureFigure 12. 12.Finite Finite Element Element model model of ofthe the InnerInner Mechanism: (a ()a Trailing-edge) Trailing-edge rotation rotation equal equal to 5 to degrees, 5 degrees, DisplacementDisplacement distribution distribution (mm); (mm); ( b(b)) Von Von Mises Mises stressstress distribution (MPa). (MPa).

Aerospace 2018, 5, x FOR PEER REVIEW 15 of 24

4.4. Definition of the Mechanical Arrangement The aeronautical needs for compactness and lightness guided the choice of the mechanical components necessary for the actuation system design. The mechanical system, shown as exploded view in Figure 13b, consists of a crank, two ternary links and three binary links. The items are Aerospaceconnected2018, 5, 122by cylindrical hinges. The crank and the ternary links are doubly supported on the rib15 of 23 block 2 (B2) plates. In such a way, out-of- plane rotations of the actuation system are strongly reduced and the actuator moment is effectively transferred along the tab hinge axis. Commercial Off-the-Shelf 4.4. Deﬁnition(COTS) actuators of the Mechanical and gearbox Arrangement were considered for the finalized design. The Harmonic-drive® strain wave gear unit was selected because of its high-power density (gear ratioThe equal aeronautical to 120, for needs CPL-17A for [30]), compactness overall dimensions and lightness and repeatability. guided the In choice such a ofway, the the mechanical torque componentsprovided necessaryby a brushless for therotary actuation motor can system be amplifi design.ed and The transferred mechanical by system, the gear shown unit to asthe exploded inner viewmechanism. in Figure 13 Eachb, consists rotary actuator of a crank, (R1, two R2 ternaryand R3) is links connected and three to a binarysegmented links. shaft The which items can are transfer connected by cylindricalthe torque hinges.to the harmonic The crank drive and gear the unit ternary of each links rib. areEach doubly gear is supportedproperly joined on the to the rib rib block block 2 (B2) plates.plate In and such it can a way, transfer out-of- the planetorque rotationsto the crank of of the each actuation inner mechanism system are (Figure strongly 13c). reduced Each crank and is the actuatordoubly moment supported is effectively by the rib transferred plates and ball along bearings the tab were hinge used axis. to reduce Commercial friction Off-the-Shelf between moving (COTS) actuatorsparts during and gearbox operations. were considered for the ﬁnalized design.

(a)

(b)

(c)

FigureFigure 13. Mechanical13. Mechanical Arrangement Arrangement of of the the Inner InnerMechanism: Mechanism: ( (aa)) Leverage Leverage within within the the available available design design spacespace in the in wingthe wing tip tip section, sectio 3D-CADn, 3D-CAD (side-view); (side-view); ((bb)) 3D-CAD3D-CAD exploded exploded view; view; (c () csection) section view. view.

The Harmonic-drive® strain wave gear unit was selected because of its high-power density (gear ratio equal to 120, for CPL-17A [30]), overall dimensions and repeatability. In such a way, the torque provided by a brushless rotary motor can be ampliﬁed and transferred by the gear unit to the inner mechanism. Each rotary actuator (R1, R2 and R3) is connected to a segmented shaft which can transfer the torque to the harmonic drive gear unit of each rib. Each gear is properly joined to the rib block plate and it can transfer the torque to the crank of each inner mechanism (Figure 13c). Each crank is doubly supported by the rib plates and ball bearings were used to reduce friction between moving parts during operations. Aerospace 2018, 5, 122 16 of 23

Aerospace 2018, 5, x FOR PEER REVIEW 16 of 24 5. Multifunctional Twistable Trailing-Edge 5. Multifunctional Twistable Trailing-Edge The Multifunctional Twistable TE concept is based on a thin-walled closed section beam layout The Multifunctional Twistable TE concept is based on a thin-walled closed section beam layout (Figure 14). Three rotary actuators (placed in root, central and tip regions) transfer the torque to three (Figure 14). Three rotary actuators (placed in root, central and tip regions) transfer the torque to three independentindependent actuation actuation systems systems consisting consisting of of harmonic harmonic drive gear gear unit unit with with a six-bar a six-bar linkage. linkage.

FigureFigure 14. 14.Multifunctional MultifunctionalTwistable Twistable Trailing-edge structural structural layout. layout.

ControlControl action action of the of the TE TE shape shape can can be be provided provided by means means of of active active ribs ribs connected connected to the to actuation the actuation systemssystems with fittingwith fitting and bondedand bonded to the to skin. the Twoskin. “c-shape”Two “c-shape” spars spars are bonded are bonded to the to skin the and skin connected and to theconnected Fowler flap to the by Fowler means flap of hingeby means fittings. of hinge fittings. The energyThe energy required required to theto the actuation actuation systems systems to twist the the trailing trailing edge edge is strictly is strictly related related to the to the geometricgeometric layout layout and and material material data. data. The The twist twist ofof aa homogeneous closed closed thin-walled thin-walled section section beam beam loaded by a torque 𝑀𝑡 and that is free to warp can be evaluated according to the Bredt equation [31]: loaded by a torque Mt and that is free to warp can be evaluated according to the Bredt equation [31]: 𝑑𝜗 𝑀 𝑑𝑠 𝑀 𝜓 = = = (13) d𝑑𝑥ϑ 4𝐴M 𝐺𝑡I ds 𝐺𝐽 M = = t = t ψ 2 (13) dx 4Am Gt GJt where 𝐴 is the area enclosed by the mid-line of the profile’s wall and ϑ is the twist angle about the X-axis, normal to the beam section. For a homogeneous cross section, the twist in case of restrained where Am is the area enclosed by the mid-line of the profile’s wall and ϑ is the twist angle about the warping can be expressed by the following differential equation [32], X-axis, normal to the beam section. For a homogeneous cross section, the twist in case of restrained 𝑑𝜗 𝑑𝜗 warping can be expressed by the following𝐸𝐶 differential− 𝐺𝐽 𝑀 equation = 0 [32], (14) 𝑑𝑥 𝑑𝑥 where 𝐽 = 𝑀 /𝐺𝜓 is the torsional constantd3ϑ of the dcrossϑ section and Cω is the sectorial moment of EC − GJ + M = 0 (14) inertia of the cross section, expressedω asdx follows:3 t dx t where Jt = Mt/Gψ is the torsional constant𝐶 = of the 𝜔 𝑡 cross 𝑑𝑠 section and Cω is the sectorial moment(15) of inertia of the cross section, expressed as follows: where 𝜔 is the sectorial area and 𝑡 is the section thickness. Z s In case of tip section free to twist but warping-constrained2 and central section fixed (no twist nor Cω = ω t ds (15) warp), the solution of Equation (14) is given by: 0

𝑀 𝛽𝑙 where ω is the sectorialϑ area = and t issinh the 𝛽𝑥 section thickness.𝛽𝑙 − 𝑥 −tanh 1 cosh 𝛽𝑥 (16) 𝐸𝐶 𝛽 2 In case of tip section free to twist but warping-constrained and central section ﬁxed (no twist nor warp), theThe solution twist angle of Equation will be maximum (14) is given at 𝑥=0 by: (tab tip section),

𝑀 𝛽𝑙 Mt 𝜗 = 𝛽𝑙 − 2tanh β l (17) ϑ = [sinhβ𝐸𝐶x + 𝛽β(l − x) − tanh2 (1 + cosh βx)] (16) EC 3 ω β / 2 where the parameter 𝛽 = 𝐺𝐽/𝐸𝐶 is the ratio between the Saint-Venant torsion rigidity 𝐺𝐽 Theand twistthe warping angle willrigidity be maximum𝐸𝐶 which atin xturn= 0depends (tab tip only section), on the cross-section geometry.

M βl = t ( − ) ϑmax 3 βl 2tanh (17) ECω β 2 Aerospace 2018, 5, 122 17 of 23

1/2 where the parameter β = (GJt/ECω) is the ratio between the Saint-Venant torsion rigidity GJt and the warping rigidity ECω which in turn depends only on the cross-section geometry. The parameter χ can be deﬁned to give indication to whether the Saint-Venant torsion or warping torsion predominates [33], l2 GJ χ2 = t = l2 β2 (18) Aerospace 2018, 5, x FOR PEER REVIEW ECω 17 of 24

For small valuesThe parameter of χ, only χ can warping be defined torsion to give needs indication to be to considered, whether the howeverSaint-Venant there torsion will or be a certain region for χwarpingwhere torsion neither predominates Saint-Venant [33], torsion nor warping torsion may be neglected and thus the 2 structural system must be analysed for mixed torsion.𝑙 𝐺𝐽 If the numerator in the expression for χ is large 𝜒 = = 𝑙 𝛽 (18) as compared to the denominator, one may expect𝐸𝐶 that Saint-Venant torsion is predominant. Indeed, the twist angleFor distribution small values will of χ be, only described warping torsion by the needs following to be considered, equation: however there will be a certain region for χ where neither Saint-Venant torsion nor warping torsion may be neglected and thus the structural system must be analysed for mixed torsion. If the numerator in the expression for Mt x 𝜒 is large as compared to the denominator,ϑ = one may expect that Saint-Venant torsion is (19) SV GJ predominant. Indeed, the twist angle distribution will bet described by the following equation:

= 𝑀 𝑥 and for x l, the maximum Saint-Venant twist𝜗 angle= will be: (19) 𝐺𝐽

and for 𝑥=𝑙, the maximum Saint-Venant twist angle willMt be:l ϑSV, max = (20) 𝑀 𝑙GJt 𝜗, = (20) 𝐺𝐽 In the aboveIn the equations, above equations, the elastic the elastic moduli moduli are are related related to the the generic generic section section (Figure (Figure 15) and 15 can) and can be obtained asbe integration obtained as integration along the along section the section proﬁle profile contour: contour: 𝐸 𝑑𝑠 R s 𝐸 = E ds (21) 0𝑑𝑠 E = (21) R s ds 0 𝐺 𝑑𝑠 ̅ 𝐺 = s (22) R 𝑑𝑠 0 G ds G = s (22) 𝑠 R where is a coordinate following the profile’s contour.0 ds The actuation torque to be provided by the actuation system can be preliminary estimated with where s is aEquation coordinate (17) assuming following the properties the proﬁle’s of the outer contour. tab mean geometric section, summarized in Table 2.

FigureFigure 15. Outer 15. Outer Trailing-edge Trailing-edge mean geometric geometric section. section.

The actuation torque toTable be provided2. Data of the byOuter the Trailing-edge actuation mean system geometric can section. be preliminary estimated with Equation (17) assuming the properties of theChord outer tab mean geometric190.8 section, mm summarized in Table2. Area 542.7 mm2 Table 2. DataShear of Centre the Outer (fromTrailing-edge Centroid), SCy mean geometric−1.995 mm section. Shear Centre (from Centroid), SCz 0.904 mm TorsionalChord Constant, Jt 2.918 × 10 190.85 mm mm4 Warping Constant, Cω 7.567 × 107 mm6 2 Characteristic LengthArea (Torsional-Bending length) 31.18542.7 mm mm Shear Centre (fromTorsion Centroid), Parameter, SC χ y 6428−1.995 mm Shear Centre (from Centroid), SCz 0.904 mm 5 4 Torsional Constant, Jt 2.918 × 10 mm 7 6 Warping Constant, Cω 7.567 × 10 mm Characteristic Length (Torsional-Bending length) 31.18 mm Torsion Parameter, χ 6428 Aerospace 2018, 5, 122 18 of 23

5.1. Assessment of the Structural Layout Aerospace 2018, 5, x FOR PEER REVIEW 18 of 24 The analytical model represented an effective tool to estimate main figure-of-merit and parameters during5.1. the Assessment preliminary of the Structural design phase.Layout To verify the mechanical behaviour of the multifunctional twistableThe trailing-edge analytical inmodel operative represented conditions, an effective finite elementtool to estimate (FE) simulations main figure-of-merit were instead and taken in account.parameters during the preliminary design phase. To verify the mechanical behaviour of the Amultifunctional glass fibre prepreg twistable with trailing-edge HexPly913 in fromoperative Hexcel conditions, composites finite [34 element] was chosen(FE) simulations as material were for the upperinstead and lower taken in skins, account. the active ribs and the “C-shape” spars. This material is widely used for the manufacturingA glass offibre rotor prepreg blades. with Indeed, HexPly913 the materialfrom Hexcel assures composites good [34] compromise was chosen between as material robustness for and capabilitiesthe upper and to lower accommodate skins, the theactive large ribs strainsand the arising “C-shape” while spars. twisting This material the trailing-edge. is widely used for the manufacturing of rotor blades. Indeed, the material assures good compromise between A 3D finite-elements model was generated by using quadrilateral plane elements (CQUAD4 [29]) robustness and capabilities to accommodate the large strains arising while twisting the trailing-edge. for the skin,A 3D C-shape finite-elements spars and model active was ribs, generated while six-facedby using solidquadrilateral elements plane (CHEXA elements [29 ])(CQUAD4 for the foam, actuators[29]) andfor the hinge skin, fittings. C-shape The spars numerical and active simulations ribs, while six-faced were performed solid elements using (CHEXA the linear [29]) static for solverthe of ® MSC-Nastranfoam, actuators[29]. and A preliminary hinge fittings. analysis The numerical was performed simulations on were the composite performed tabusing without the linear aerodynamic static loadssolver to assess of MSC-Nastran the maximum® [29]. twisting A preliminary torque requiredanalysis was by theperformed actuation on systemthe composite to implement tab without Mode B as wellaerodynamic as the structural loads to strength assess the of maximum the twisted twisting skin. Thetorque central required “active” by the rib actuation was constrained system to in all degreesimplement of freedom; Modeenforced B as well as rotations the structural (SPCD strength [29]) were of the applied twisted atskin. the The tip central active “active” rib (+5 degrees,rib was “+” standingconstrained for downward) in all degrees and of at freedom; the root enforced active rib rota (−tions5 degrees, (SPCD “ [29])−“standing were applied for upward).at the tip active As shown in Figurerib (+5 16 degrees,b, a maximum “+” standing strain for of downward) about 1.34% and is at reached the root near active the rib cut-out (−5 degrees, of the “ central−“ standing section. for For upward). As shown in Figure 16b, a maximum strain of about 1.34% is reached near the cut-out of the most part of the remaining structure the strain is lower than 0.8%. According to [35], the maximum the central section. For the most part of the remaining structure the strain is lower than 0.8%. allowedAccording strain toto 2.5%[35], the along maximum fibres directionallowed strain is equal to 2.5% and along to 0.5% fibres along direction the direction is equal perpendicularand to 0.5% to the fibres.along Thethe direction torque required perpendicular to enable to the linear fibres. twist The to isrque equal required to 64.05 to Nenable·m for linear the tiptwist actuation is equal systemto and 103.664.05 NN·m·m for for the the tip root actuation one. system and 103.6 N·m for the root one.

(a)

(b)

FigureFigure 16. Finite 16. Finite Element Element analysis analysis with with the the maximummaximum linea linearr twist twist for for Mode Mode B from B from tip to tip root to rootsection section (−5◦/0(−5°/0°/+5°):◦/+5◦): ( a(a)) DisplacementDisplacement (mm); (mm); (b ()b Strain) Strain distribution, distribution, Max Max Principal Principal [MicroStrain]. [MicroStrain].

Aerospace 2018, 5, 122 19 of 23

Aerospace 2018, 5, x FOR PEER REVIEW 19 of 24 A second analysis was performed to prove the structural integrity of the composite structure A second analysis was performed to prove the structural integrity of the composite structure upon the limit load condition (depicted in Figure 17) pertaining to Mode A (TE rigid rotation equal to upon the limit load condition (depicted in Figure 17) pertaining to Mode A (TE rigid rotation equal +5 degrees downward). In such a case, all actuation systems must withstand the total aerodynamic to +5 degrees downward). In such a case, all actuation systems must withstand the total aerodynamic hinge-momenthinge-moment pertaining pertaining to theto the limit limit load load of of Table Table2 .2.

FigureFigure 17. 17.Strain Strain distribution distribution [MicroStrain] [MicroStrain] under under limit limit load load condition. condition.

As shownAs shown in Figurein Figure 17 17,, a a maximum maximum strain strain of ofabout about 1.42% 1.42% is reached is reached near the near central the section. central The section. The maximummaximum displacement is isequal equal to to7.39 7.39 mm mm at atthe the midpoint midpoint of the of theinboard inboard tab tabregion. region. In the In the remaining part of the structure, the overall displacement is below 3 mm which demonstrates remaining part of the structure, the overall displacement is below 3 mm which demonstrates sufficient sufficient stiffness of the system; in compliance with regulations (EASA CS 25.305(a)) the deformation stiffnesslevels of do the not system; interfere in compliancewith the safe withoperation regulations at limit load (EASA condition. CS 25.305(a)) the deformation levels do not interfere with the safe operation at limit load condition. 5.2. Estimation of Trailing-Edge Performance in Operative Conditions 5.2. Estimation of Trailing-Edge Performance in Operative Conditions According to airworthiness requirements [25], aircraft structures must withstand the limit load Accordingconditions in to order airworthiness to prove its requirements structural integrity [25], across aircraft the structures overall flight must envelope. withstand the limit load conditionsHowever, in order during to prove the its regular structural airplane integrity flight acrossmission the profile, overall load-bearing flight envelope. aircraft structures However,have to withstand during aerodynamic the regular airplaneloads lower flight than mission the ones profile, prescribed load-bearing in the limit aircraft condition. structures have to withstandIn such aerodynamic cases, the multifunctional loads lower than twistable the ones trailing-edge prescribed used in on the the limit AG2-NLF condition. wing is expected Into suchbe activated cases, the in multifunctional high speed climb twistable flight trailing-edgeconditions. To used enable on thethe AG2-NLF transition wingfrom isbaseline expected to configuration to the rigid trailing-edge deflection (mode A) equal to 2.5°, all rotary brushless motors be activated in high speed climb flight conditions. To enable the transition from baseline configuration have to be activated (R1, R2 and R3) transferring to their respective “active” ribs the same angle. In to the rigid trailing-edge deflection (mode A) equal to 2.5◦, all rotary brushless motors have to be this case, the actuating torques required to enable this transition are reported in Figure 18. activated (R1, R2 and R3) transferring to their respective “active” ribs the same angle. In this case, the Aerospace 2018Continuous, 5, x FOR PEER span-wise REVIEW increase in deflection angles (mode B) from tip to root can be enabled20 of 24 actuatingwhit torquesthree independent required toactuation enable systems this transition transferring are reported angles equal in Figure to (4°/3°/2°) 18. to their respective

“active” ribs. Considering the aerodynamic load pertaining to the climb condition (𝑀 =0.36 at 4572 m) with TE morphing in Mode B (4°/3°/2°), the structural layout shows the maximum strain equal to 0.28 per cent and the required actuating torques (per each active rib) are summarized in Figure 19. During morphing TE operations, each actuation system can provide the respective active rib with an output torque equal to

𝓜𝑶𝑼𝑻 =𝑴𝒕 ∙ 𝑴. 𝑨. ∙𝑭𝑯𝑫 ∙𝜼𝑯𝑫 ∙𝜼𝑰𝑴 (23)

where 𝑴𝒕 is the torque provided by each rotary hollow-shaft brushless motor, 𝑭𝑯𝑫 is the gear ratio of the Harmonic Drive (CPL 17-2A) [30], 𝜂 is the efficiency of the Harmonic Drive and 𝜂 the efficiency of the inner mechanism. The output torque at each active rib can be expressed as a function of the crank position (Figure 20).

Figure 18. Actuating torques required by the actuation systems for enabling morphing TE modes Figure 18. Actuating torques required by the actuation systems for enabling morphing TE modes during high speed climb ﬂight conditions (M∞ = 0.36 at 4572 m). during high speed climb flight conditions (𝑀 =0.36 at 4572 m).

(a)

(b)

Figure 19. TE finite element analysis with continuous increase deflection angles (Mode B) from tip to root section (+4°/+3°/+2°): (a) Displacement (mm); (b) Strain distribution, Max Principal [MicroStrain].

Aerospace 2018, 5, x FOR PEER REVIEW 20 of 24

Aerospace 2018, 5, 122 20 of 23

Continuous span-wise increase in deﬂection angles (mode B) from tip to root can be enabled whit three independent actuation systems transferring angles equal to (4◦/3◦/2◦) to their respective “active” ribs. Considering the aerodynamic load pertaining to the climb condition (M∞ = 0.36 at 4572 m) with ◦ ◦ ◦ TE morphingFigure 18. inActuating Mode B torques (4 /3 /2 required), the structuralby the actuation layout systems shows thefor maximumenabling morphing strain equal TE modes to 0.28 per

centduring and the high required speed climb actuating flight torquesconditions (per (𝑀 each =0.36 active at 4572 rib) arem). summarized in Figure 19.

(a)

(b)

FigureFigure 19. 19. TETE finite ﬁnite element element analysis analysis with with continuous continuous increase increase deflection deﬂection angles angles (Mode (Mode B) from B) tip from to roottip tosection root section(+4°/+3°/+2°): (+4◦/+3 (a) ◦Displacement/+2◦): (a) Displacement (mm); (b) Strain (mm); distribution, (b) Strain Max distribution, Principal Max[MicroStrain]. Principal [MicroStrain].

During morphing TE operations, each actuation system can provide the respective active rib with an output torque equal to = ·( )· · · MOUT Mt M.A. FHD ηHD ηIM (23)

where Mt is the torque provided by each rotary hollow-shaft brushless motor, FHD is the gear ratio of the Harmonic Drive (CPL 17-2A) [30], ηHD is the efﬁciency of the Harmonic Drive and ηIM the efﬁciency of the inner mechanism. The output torque at each active rib can be expressed as a function of the crank position (Figure 20). Aerospace 2018, 5, 122 21 of 23 Aerospace 2018, 5, x FOR PEER REVIEW 21 of 24

FigureFigure 20. 20. ActuatingActuating torque torque provided provided by each actuation systems to its active rib during operations.

6. Conclusions Conclusions The use of of a a morphing morphing system system using using a a multifunct multifunctionalional twistable twistable trailing-edge trailing-edge has has been been evaluated evaluated with reference to the NLF wing ofof thethe CLeanSky2CLeanSky2 regionalregional aircraft.aircraft. For For such such subsonic subsonic aircraft, aircraft, performance improvements cancan bebe expectedexpected only only by by recovering recovering the the loss loss of of laminar laminar flow flow that that occurs occurs at low (on the lower surface) or high (on the upper surface) C values. at low (on the lower surface) or high (on the upper surface)L 𝐶 values. As concerns the wing trailing-edge, the structur structuralal and kinematic design process of the actuation actuation system were completely addressed: three three rotary rotary brushless brushless motors motors (placed (placed in in root, root, central and tip sections) were required to activate the inner mechanismsmechanisms enabling differen differentt trailing-edgetrailing-edge morphing modes. The The structural layout of the thin-walled cl closed-sectionosed-section composite trailing-edge represents a promising conceptconcept toto balance balance the the conflicting conflicting requirements requirements between between load-carrying load-carrying capability capability and shape and shapeadaptivity. adaptivity. Actuation Actuation system system performances performances and aeroelastic and aeroelastic deformations, deformations, considering considering both operative both operativeaerodynamic aerodynamic and limit loadand conditions,limit load conditions, prove the potential prove the of thepotential proposed of the structural proposed concept structural to be conceptenergy efficientto be energy and lightweight efficient and for lightwe real aircraftight for implementation. real aircraft implementation. Final weight of the multi-functional twistable trailing-edge is summarized in Table 3 3.. OverallOverall system implications havehave toto bebe made made with with reference reference to to the the weight weight of of a conventionala conventional outboard outboard flap flap tab tabused used for regionalfor regional aircraft aircraft (Table (Table4). 4). RetrofittingRetrofitting a regional aircraft with such device, a 3.52% weight weight increase of the the outboard outboard flap flap tip tip segment willwill bebe produced. produced. At At aircraft aircraft level, level, the th Maxe Max Zero Zero Fuel Fuel Weight Weight (MZFW) (MZFW) will increase will increase of 0.012% of 0.012%only. Finally, only. theFinally, mechanical the mechanical power required power required to enable to load enable control load (LC) control functionalities (LC) functionalities can be proved can beto beproved extremely to be affordable.extremely affordable. Assuming theAssuming maximum the resisting maximum torque resisting (on crank torque link) (on equal crank to link) 23.35 equal N·m toand 23.35 a crank N·m speedand a crank rotation speed of20 rotation deg/s, of the 20 totaldeg/s, mechanical the total mechanical power will power be equal will tobe 24.34equal Wattto 24.34 for ◦ ◦ ◦ Wattelastic for twist elastic mode twist (4 mode/3 /2 (4°/3°/2°).).

TableTable 3. Multifunctional twistable wing trailing-edge: weight breakdown.

SkeletonSkeleton (Kg) (Kg) SkinSkin 4.582 4.582 C-Spars 1.381 C-Spars 1.381 Active ribs 0.371 Active ribs 0.371 Foam 1.506

Hinge ﬁttingsFoam 1.506 0.103 ActuatorHinge ﬁttings fittings 0.103 0.115 TotalActuator weight fittings 0.115 8.058 ActuationTotal system weight 8.058 (Kg) Inner mechanisms,Actuation joints, system bearings (Kg) 3.13 BrushlessInner mechanisms, motors, gear-boxes, joints, shafts bearings 3.13 6.45 BrushlessTotal motors, weight gear-boxes, shafts 6.45 9.58 Total weight 9.58

Aerospace 2018, 5, 122 22 of 23

Table 4. Regional Turbo-Prop aircraft (90 passengers).

Maximum Take-Off Weight (MTOW) 33,200 Maximum Zero Fuel Weight (MZFW) 31,200 Operative Empty Weight (OEW) 19,360 Outboard Fowler Flap (conventional) 104.65 Flap tip segment (all-metallic) 13.95 Multifunctional twistable wing trailing-edge (Kg) Structural skeleton (composites materials) 8.06 Actuation system 9.58 Total weight 17.64 Outboard Fowler ﬂap weight increase 3.688

Author Contributions: F.R.: Conceptualization, Methodology, Software, Validation, Formal Analysis, Investigation, Resources, Data Curation, Writing-Original Draft Preparation. F.A.: Conceptualization, Methodology, Investigation, Resources, Supervision, Project Administration. R.P.: Investigation, Resources, Conceptualization, Writing-Review & Editing, Visualization, Supervision, Project Administration, Funding Acquisition. F.M.: Methodology, Software, Validation, Investigation, Visualization, Data Curation, Writing-Original Draft Preparation. Funding: Part of the research described in this paper has been carried out in the framework of AIRGREEN2 Project, which gratefully received funding from the Clean Sky 2 Joint Undertaking, under the European’s Union Horizon 2020 research and innovation Program, Grant Agreement No. 807089—REG GAM 2018—H2020-IBA-CS2-GAMS-2017. Acknowledgments: The authors would like to thank Leonardo Aircraft, for having provided the industrial guidelines and the necessary support to the research work addressed by this paper. Conﬂicts of Interest: The authors declare no conﬂict of interest.

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