morphing composite structures for adaptive high lift devices · 2015-05-31 · l1t2), with high...

Proceedings of the 6th International Conference on Mechanics and Materials in Design,

Editors: J.F. Silva Gomes & S.A. Meguid, P.Delgada/Azores, 26-30 July 2015

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PAPER REF: 5555

MORPHING COMPOSITE STRUCTURES FOR ADAPTIVE HIGH

LIFT DEVICES

Alessandro Airoldi1(*)

, Paolo Panichelli1, Alessandro Gilardelli

1, Giuseppe Quaranta

1, Giuseppe Sala

1

1Department of Aerospace Science and Technologies, Politecnico di Milano, Milano, Italy

(*)Email: [email protected]

ABSTRACT

The work presents a design study for an aeronautical flap capable of smooth and progressive

shape variations. Optimal configurations for takeoff and landing are taken from literature to

define a reference design case and the corresponding loading conditions. An extraction

mechanism to set the flap in optimal position is identified. Then, a morphing composite

structure is devised to smoothly and progressively change the shape of the high lift device to

enhance the functions and the versatility of such aerodynamic surface. An internal structure

based on chiral auxetic cellular topology and a skin system supported by a corrugated

composite panel are proposed. The technological processes for manufacturing such

components are presented and discussed. A parameterized finite element model of the

morphing flap is developed and structural analyses are carried out to investigate the feasibility

of the morphing concept.

Keywords: morphing structures, chiral topologies, corrugated laminates, composite structures.

INTRODUCTION

In the last decades, several types of morphing structural concepts have been presented for

application in the aerospace field (Sofla, 2003, Barbarino 2011). In several cases, morphing

structures are proposed to substitute the motion of rigid surfaces to optimize the aerodynamic

shape in different mission segments. Although there is no general agreement of the meaning

of the term, a morphing structure is typically characterized by a smooth and progressive shape

change, which occurs without introducing gaps in the shape of the aerodynamic surface. High

lift devices are one of the most relevant example of wing shape adaptation to different flight

phases, since they combine both large chord and camber modifications to increase the lift

during takeoff and landing. However, optimal configuration in such two conditions can be

significantly different, as it can be proved by means of computational fluid dynamic

calculations (Bernini, 2011).

The application of morphing concepts to high lift devices have been studied to increase the

adaptability and the functions of the flaps located at trailing edges in (Monner, 2000). A

segmented rib covered by an elastomeric skin was proposed, to smoothly adjust the camber of

the airfoil during the flight, thus optimizing the shape for different flight conditions, and to

increase, with the same morphing surface, the lift during take-off and landing. Indeed, the

development of an effective morphing is an active research area in recent years, as it is

demonstrated by the effort carried out in important research programs at a European and

worldwide level.

Many researches are based on internal structures made of rigid segments, although an

alternative for the design of a morphing wing is represented by a completely deformable

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Composite and Advanced Materials

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structural architecture. Such solution, which is exemplified by the deformable rib concept

proposed in (Campanile, 2001), eliminates any type of internal mechanism and enhance the

adaptability of the morphing concepts.

The development of a completely deformable structure with the capability to adapt to

different shapes represents a difficult task for structural engineers. Indeed, morphing

capabilities should be obtained while avoiding the introduction of week points in the structure

and by using structural architectures that maintain adequate stiffness and strength in non-

morphing directions. In particular, a deformable morphing structure is required to change the

shape without losing the aerodynamic efficiency, so that a controlled shape change should not

significantly modify the thickness distribution in the profile. An appealing solution is

represented by honeycombs with a negative in-plane Poisson’s coefficient (auxetic) that can

be produced, among other alternatives, by adopting a chiral geometry, which is a special non

centre-symmetric topology, originally proposed in (Lakes,1999). Chiral topologies consists of

circular elements, called nodes, connected by straight ligaments. The ligaments are tangential

to the nodes and tend to wind or unwind about the nodes under the action of a compressive or

tensile in-plane force, respectively. Such deformation mechanism originates a negative

Poisson’s coefficient and has already been proposed by several authors to design morphing

structures that can smoothly change their shape under the actions of aerodynamic loads

(Bornengo, 2005, Spadoni, 2007, Bettini, 2010, Airoldi, 2012). Such works prove that

morphing structures based on chiral topologies can obtain shape variations still maintaining

structural integrity and aerodynamic efficiency. Moreover, such topologies are characterized

by a significant design flexibility and can be adapted to various morphing concepts and

actuation strategies.

Either if a completely deformable or a rigid internal structure is adopted, a flexible skin is a

fundamental ingredient for any morphing solution. Several solutions have been proposed to

address this issue (Thill, 2008), but a particularly promising concept is represented by

composite corrugated panels (Yokozeki, 2006). Thanks to their design flexibility and highly

anisotropic properties, corrugated laminates are optimal to allow large deformation in the

morphing direction retaining stiffness and strength in other directions. Although corrugation

can have a detrimental effect on aerodynamic performances (Xia, 2014), solutions have been

developed to integrate, at a moderate weight cost, an elastomeric cover to obtain an efficient

morphing skin system (Fournier, 2013).

The present work is aimed at presenting the technological routes to manufacture internal

composite chiral structures with morphing capabilities and aerodynamically efficient skins

based on corrugated composite laminates. Such basic components are then considered to

design a variable shape slotted flap, which can progressively change its shape to obtain

optimized configurations for takeoff and landing conditions.

In the following section of the paper, the main issues regarding the design of flap devices are

presented. Thereafter, a design case is selected from literature so that an optimized shape of

the flap is available for the introduction of morphing capabilities. Moreover, the design case

will provide the evaluation of the pressures that act on the flap in operative conditions. The

design case is completely defined by identifying a mechanism that can set the rigid flap in the

optimal positions for the take-off and landing. In a subsequent section, the potential

advantages of a morphing flap are outlined and the technologies for the production of both

internal structure and the external skin are described. Such technologies are taken into

consideration in the final section, to prepare the tools for the development of a parameterized

model of the morphing flap and to apply them in order to select a configuration that can



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obtain a significant shape change under the action of an efficient actuation system. The

response of a representative configuration of the morphing flap is then analyzed. The main

findings are summarized in a conclusive section.

FLAP DESIGN CONSIDERATIONS AND SELECTION OF A REFERENCE CASE

A flap is a high lift devices installed at the trailing edge and is used to increase the wing

chamber and planform area, thus leading to higher values of the lift coefficient and to a

reduction of the stall velocity of the aircraft. During takeoff, the use of flaps reduces the

runaway distance and the climb rate required. In the landing phase, when the flaps are

typically fully extended, the additional drag generated contributes to a shorter landing. Flaps

also allow a better glide slope control and improve the pilot’s vision over the nose during

landing by reducing the aircraft attitude.

Historically, as cruise speed increased with the development of more powerful engines, wing

loading increased and a real need of high lift devices emerged to keep takeoff and landing

speeds within reasonable limits. Initially, the design objectives were mainly targeted to

maximum lift requirements, accepting an increase in the complexity of wing structure. Such

trends were also supported by research works, such as (Smith, 1975), where it was

demonstrated that a flap with N+1 elements is always more effective than one with N

elements. The complexity reached a peak on the solution adopted for Boeing 747 (Rudolf,

1996a and Rudolf 1996b) which consists of a triple slotted flap. Subsequently, aircraft

engineers tried to achieve high levels of lift with simpler devices, in order to reduce structural

complexity and maintenance costs. In fact, a high lift system can account in some cases for

about 10% of the production cost of a typical jet transport.

Issues in the design of high lift devices

A good trade-off between flap effectiveness and constructive simplicity is the single slotted

fowler flap (Rudolf, 1996a), which consists of a panel with a fully developed leading edge

that pivots below the wing surface. A linkage mechanism extracts the flap, which is partially

overlapped to the wing in retracted configuration, and increases the wing area. The motion

generates a gap between the wing body and the extracted flap and this allows a suction of

airflow from the lower surface boundary layer, which accelerates and stabilizes the upper one,

delaying the separation and reducing both drag and loss of lift increment (Di Matteo, 2011).

The mechanism that drives the flap to the deployed position should provide most of the

available fowler motion in the initial deployment, at low deflection angles. The increment of

wing area at low angle settings results in the best efficiency for takeoff. For such a reason the

combination of extraction motion with rotation is to be preferred with respect to the simple

rotation provided by a simple hinge. Moreover, a flap with a simple hinge mechanism (single

slotted flap) is difficultly adapted to fit with swept wings. In the single slotted fowler flap, a

four bar linkage or a link-track mechanism provides the combination of motions. In particular,

a link-track mechanism obtains a bigger initial fowler motion with a less frontal area than

other solutions, thus leading to lower aerodynamic drag (Van Dam, 1999).

Optimized flap configurations and shapes for a selected design case

The reference design case for the activities presented in this work is taken from (Benini,

2011). In such work a multi-objective optimization problem is presented, considering an

airfoil endowed with a single slotted fowler flap. The purpose of the optimization process was

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to enhance the aerodynamic performance both at takeoff and landing by searching for optimal

shape and setting parameters.

The selected baseline geometry is the case A-2 from AGARD-AR 303 (AGARD, 1993),

which is a two dimensional supercritical airfoil, referred to as NHLP 2-D (configuration

L1T2), with high lift devices consisting of a 12.5% chord leading edge slat and a 33% single

slotted fowler flap. The code used for fluid dynamics computations is the viscous/inviscid

interaction code MSES (Drela, 1987). Three design variables were used to define each of the

trailing edge device positions relative to the main wing element, namely the flap deployment

angle, the gap and the overlap length. The flap mechanism and the design variables are

graphically represented in Figure 1, which is based on the quantitative data published in

(Benini, 2011).

The optimization study was also aimed at a partial optimization of the shape. It should be

observed that the authors do not consider a morphing flap, so that the shape of the trailing

edge is not optimized, since it is considered a constraint for cruise conditions. Shape

optimization is limited to regions of the geometry that are hidden when the airfoil is in the

retracted configuration. A Bézier curve representation of the leading edge shape was adopted,

so that modifications could be carried on manipulating the control points’ position over the

curves.

Despite this limitation, the considered design modifications could change the flow behavior

and pressure distribution over the surface in the flap extracted conditions, potentially

increasing the overall aerodynamic performance of the multi-element airfoil.

Fig. 1 - Flap mechanism and design varibles for aerodynamic optimization of the flap:

The optimization work was aimed at improving the aerodynamic characteristics of the multi-

element airfoil in both takeoff and landing conditions. The definition of the objective function

in takeoff was based on the reduction of the time-to-climb (which is directly related to fuel

consumption). On the other hand, the reduction of fuel consumption and noise emission is one

of primary requirements for high lift configurations in landing conditions and this objective

overlap

gap

δ



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can be achieved by minimazing the drag coefficent. It is worth noting, however, that,

especially in a 2-D design study, still achieving high lift coefficients is essential for landing.

This is the reason why an inequality constraint has been considered for the lift coefficent in

this condition. A second family of constraints was adopted considering the pitching moment

coefficent about the quarter-chord point, to avoid an excessive increase of the airfoil negative

pitching moment: the pitching moment coefficient of new airfoils were allowed to be grater

than or equal to 110% of the original values.

The multi-objective optimization study resulted in a Pareto’s front and leads to the selection

of two solutions named “OPT1” and “OPT2”, which quantify the improvements in one

objective while the other is kept fixed. OPT1 is the case with the best time-to-climb (15% of

reduction with respect to initial configuration) while keeping at landing the same drag

coefficient of initial configuration. Configuration OPT2 assures an 8% reduction in CD at

landing, mantaining the same time-to-climb at takeoff. The shape of the Pareto front

demonstrates that a conflict does exist between the two objectives, thus suggesting that

different optimal airfoil shape having different settings can be found, which maximize airfoil

performance in the two operating conditions.

Generally, optimization leads to configurations with thicker airfoils in the leading edge region

and a more pronounced curvature on the pressure side with respect to the initial guess.

Moreover, a lower value for the rotation angle was identified. Such choices assured higher lift

and lower drag at takeoff, as well as a much lower drag in landing.

Objective of the presented work

The two optimal shapes identified in the work conducted in (Benini, 2011) are very similar,

because only the leading edge region of the flap was modified. The main differences in the

optimal configuration lies in the nose shape and the rotation angle.

Accordingly, the first part of the research activity was aimed to define a mechanism that could

obtain both the optimal flap configuration for takeoff and landing, defined in the reference

design case, for a rigid flap. A tool to solve the inverse kinematics of the mechanism has been

implemented, by selecting as design variables the geometric parameters of the mechanism.

The objective functions were represented by the two final positions of the flap, which must be

reached within a given tolerance.

The second part of the paper introduces morphing capabilities in the defined configuration. A

structural architecture based on selective deformable structures is applied. The morphing

structure is aimed at increasing the aerodynamic performance and versatility through the

adoption of a lightweight and highly deformable structure with integrated actuators. The

capability of an aerodynamic surface to achieve a continuous shape change upon a specific

actuation allows the wing to adapt its shape by smooth and progressive variations.

Such morphing capabilities will make possible to adapt the flap shape in order to achieve

near-optimal lift and drag profiles for all the possible takeoff and landing conditions, in the

presence of different external conditions, for all possible weight configurations and even

taking into account different versions of the same aircraft, without the need of an expensive

re-configuration. Moreover, morphing capability will be exploited even in retracted

conditions, throughout all the different phases of the flight to reduce fuel consumption, to

increase aerodynamic efficiency, to improve control of the aircraft and airload distribution

over the wing. All these benefits allow the aircraft to operate over a wide range of flight

conditions, whereas a conventional rigid wing would be in a non-optimal configuration during

a large part of the mission.

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MECHANISM IDENTIFICATION FOR A RIGID FLAP

A tool has been implemented in Matlab to find a mechanism whose kinematics is able to

meet, within a given tolerance, both the optimum takeoff and landing shown in [2]. The

design variables used to define the kinematics are shown in Figure 2, where XC and XPini are

vectors representing the in-plane position of the crank hinge and of the flap pivot,

respectively, whereas R, L, x, α, β, and γ are scalars that describe the mechanism

configuration.

Fig. 2 - Design variables of the mechanism.

The kinematic equation can be written as:

+ cos ∝ + cos = + cos

+ sin + sin = + sin Eq. 1

Taking the x form the first equation, and replacing it in the second one, the kinematic equation

in the only variable β can be obtained, as α is an input data and the other variables are

determined by the kinematic:

Φ = sin − tan cos + − − tan ∙ − + cos + sin = 0 Eq. 2

The flap deflection δ can be easily calculated as it is defined as β minus a rigid rotation.

The kinematic parameters of the mechanism are collected inside the column vector ωcin:

= ∝ Eq. 3

where αini is the crank angle when the flap is in retracted position.

Variable γ is not directly defined but it is calculated through the position of the end point of

the track XPfin:

= tan!" #$%&'(!$%'(')%&'(!)%'('* Eq. 4

XPfin is not an independent variable, because it is the final position of the flap pivot and can be

determined once its initial position is defined, through a rigid roto-translation dictated by the

optimum landing configuration in (Benini, 2011). Hence the independent parameters defining

the kinematics are:



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+ = ∆ ∆ ∝ Eq. 5

where the initial position of the pivot is considered in the flap reference system.

The inputs to the function are the kinematic parameters Ωcin, the α angle and the parameters of

gap, overlap and deflection shown in Fig.1, defining the flap in landing. The outputs are the

trajectory followed by the pivot (xp, yp) and the deflection δ of the flap for a given angle α.

Optimal solutions have been found by applying a Monte-Carlo method, considering a space

around an initial guess ΩcinM. A symmetric range of variability around the initial guess was

chosen and represented by the vector ±∆Ωcin.

Ω . = 0.652 0.0352 0.252 02 ∝ 0.132

Δ+ . = 30 30 30/750 30/750 Δ ∝ 30 Eq. 6

The angle αini this has been evaluated between 180 and 270 degrees, with a constraint on the

minimum value to avoid the interference of the mechanism with the rear spar, whose position

(xspar, yspar) was assumed at 60% of the chord.

For each vector of parameters inside the range defined above, the kinematic has been

calculated for α going from αini to 0 degrees. Among all such configurations, the closest

positions to optimum takeoff (∆LETO, ∆TETO) and landing (∆LELDG, ∆TELDG) have been

identified and stored, together with the maximum frontal area of the mechanism, which is

given by:

9:;) = − − <=;> Eq. 7

Among all the kinematics, a solution that leads to a very low value of objective function f,

given in Eq. 8, was finally chosen:

? = ∆@ AB + ∆C@ AB + ∆@DEFB + ∆C@DEFB Eq. 8

The optimal solution is characterized by the following parameters:

Ω A = 1456.7HH 48.8HH 0.292 −0.042 242.2° 310.5HH Eq. 9

Such solution obtains in the following discrepancies with respect to the optimal positions:

∆@ A = 1.78HH, ∆C@ A = 0.62HH, ∆@DEF = 2.97HH and ∆C@DEF = 1.46HH

The resulting flap is shown in Fig. 3-a, whereas Fig. 3-b show the comparison between the

solution found and the ideal optimal solutions that have been deducted by the data published

in (Benini, 2011). Since the OPT1 and OPT2 solutions identified in (Benini, 2011) are very

similar, only the comparison with OPT1 is presented.

Results reported in Fig. 3 indicates that the adopted procedure leads to the identification of a

mechanism that obtains positions very close to the ones identified in (Benini, 2011) for

optimal performances in takeoff (minimum time-to-climb) and landing (minimum drag).

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Fig. 3 - Progression of flap deployment according to the identified mechanisms (a) and comparison with

optimized solutions (b)

A TECHNOLOGICAL SOLUTION FOR THE DEVELOPMENT OF A

MORPHING FLAP

Potential advantages of a morphing flap

The scope of high lift devices is to modify in a profitable way the shape of the wing in the

different segments of the aircraft mission. However, it is hard, with only rigid movements of

airfoil’s parts, to achieve the optimum airframe configuration for any single flight condition,

as it varies and depends on several parameters, such as altitude velocity and mass. Moreover,

the attempt to obtain optimal shapes for more and more segments of the flight mission

brought, in the past, to an increase in the complexity of the mechanisms, which means an

increase in weight, noise and maintenance cost.

The use of morphing technologies instead of conventional high lift devices would offer

several advantages. The possibility to modify smoothly and seamlessly the trailing edge shape

would enable the aircraft to fly in the optimal configuration during all the phases of the flight,

achieving higher aerodynamic efficiency and generating less off-design effects. The lower

number of components and gaps would allow a more laminar airflow over the wing surface

reducing drag, and, consequently, fluid consumption and noise.

Moreover, in cases where torsional stiffness of the wing is not enough to avoid the reversal of

outboard ailerons, such devices have to be located in more inboard sections, with several

penalties regarding aerodynamic performance and maximum allowable flap span. The

adoption of morphing flaps would make available such surfaces as high-speed ailerons or

control tab (flaperon).

A single slotted fowler flap with morphing capabilities could provide the aforementioned

advantages in the retracted positions, since it would represent a morphing trailing edge for the

wing, and would be able to perform the high lift device functions with enhanced efficiency.

Considering takeoff and landing performances, a single rigid slotted fowler flap poses two

potential major obstacles in landing, since the maximum lift coefficient produced is lower

than in the case of multiple slotted flap and this can lead to very high angles of attack. For

such type of flaps, deflection angles up to 40 degrees are possible without any additional

devices, but higher angles can be obtained with additional features, such as vortex generators

on the upper surface of the flap leading edge or Gourney flaps on the underside of the spoiler.

(a) (b)



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The adoption of a morphing flap with variable camber would provide a great aid in all these

situations and may allow avoiding such solutions, which increase complexity and often leads

to undesired effects.

Additional considerations regarding the advantages of morphing flaps derives from the need

of increasing the lift coefficient to extend the mission envelope or in case of airplane re-

configurations. Such need arises frequently, since the mission envelope of an airplane is often

extended with respect to initial design requirements. The reconfiguration of high lift system to

accomplish new conditions, for instance a landing with a higher maximum weight, typically

penalizes operations with less demanding requirements but higher recurrence. Therefore, the

airplane is no more able obtain optimal performances during a large part of the mission.

Moreover, the airplane plans for growth, in weight or fuselage dimensions, can generally be

accommodated by changes in trailing edge flaps. This issue is generally handled by selecting

a device configuration that lies in the middle of the desired range, without the capability of

reaching an optimum in any of the cases. Alternatively, different high lift systems for aircrafts

of the same family are developed, although it could be very expensive.

The idea of using morphing capabilities can relax some design requirements on the high lift

system and allow a more profitable and efficient device in terms of design, production and

maintenance cost. Considering the cost for the development of a high lift device, a morphing

flap would represent an appealing solution for the next generation of aircraft.

Composite deformable ribs based on chiral topologies

A specific structural architecture is considered in this work for the introduction of morphing

capabilities in the selected reference design case. The proposed solution makes use of two

basic structural concepts, which are applied to design the internal structure and the skin of the

morphing flap, respectively.

Fig. 4 - Main geometrical parameters of hexa-chiral topologies (a) and deformed shape of a variable camber

airfoil hosting a hexa-chiral core

The internal rib of the flap is designed according to a cellular auxetic topology that is

manufactured by using composite materials. In particular the hexa-chiral geometry, which

was already proposed in (Bornengo, 2005, Spadoni, 2007, Bettini, 2010 and Airoldi 2012) for

(a) (b)

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morphing application is considered. The unit shown in Fig. 4-a presents the main geometrical

parameters of such cellular topology, which are represented by the ligament length, L, the

distance between the centre of the nodes, R, the node radius, r, and the ligament thickness, t.

The angle ϑ is fixed to 30°, whereas the angle ϕ can be expressed as a function of the other

parameters. The three parameters R, L and r can be related by an additional equation.

The hexa-chiral topology shares, with other types of chiral network (Alderon, 2010) peculiar

properties, which are particularly appealing for morphing applications.

- auxetic behavior (that is negative Poisson’s ratio) promotes the diffusion of the

deformation due to a locally applied force throughout the structure (Bettini, 2010);

- negative Poisson’s ratio involves high in-plane shear stiffness (G=E/2(1+v)) and this

opposes to local deformation;

- the two aforementioned considerations indicates that an hexa-chiral tessellation is a

deformable structures that inherently avoids weak points in the structure, for any type

of actuation, and can be adapted to assume different shapes; therefore it is not a

flexible mechanics that is optimized for a single morphing performance;

- the hexa-chiral topology provides a large design flexibility, since two parameters

among R, L, r and thickness t can be varied to obtain a large range of stiffness

responses.

- adoption of composite lamiantes to manufacture the ligaments can enhance design

flexibility, by properly selecting the material and the lay-up sequence; moreover, it

represents a convenient method to functionally grade the properties of the cellular

structure.

- chiral networks are suited to design passive morphing concepts, where the forces that

induce morphing are the aerodynamic pressures, as well as active morphing concepts;

in particular application of torques to the chiral nodes is able to induce an overall

change of shape for the entire airfoil.

A feasible technological route to produce composite chiral honeycombs was developed in

previous works (Bettini, 2010, Airoldi 2012b) and the availability of this technology

motivated the study of morphing applications such as the chiral sail concept presented in

(Airoldi, 2012a). In such work, an airfoil hosting an hexa-chiral composite core, with the

capability of increasing the camber at increasing angle of attack was designed by means of an

optimization process that took into account aerodynamic, aeroleastic and structural strength

issues. A deformed shape for such morphing concept is shown in Fig. 4-b.

The technology for composite chiral honeycombs, described in (Bettini, 2010 and Airoldi,

2012b), was based on the production of curved ligaments that were subsequently superposed

to form the chiral nodes. The assembly of ligaments was bonded in an autoclave process by

using a special assembly mould where elastomeric inserts were introduced to apply a

uniformly distribute the pressure on the ligament surfaces and to optimize bonding quality.

In this work, an enhancement of such technology is taken into account, which moves from the

considerations that hexa-chiral networks can be produced by adjoining triangular elements, as

it is shown in Fig. 5.


Editors: J.F. Silva Gomes & S.A. Meguid, P.Delgada/Azores, 26

Fig. 5 - Generation of an hexa

Such configuration originates a hexa

to behave as the original topology, since nodes are simply required to be the most rigid as

possible in order to obtain the desired auxetic response.

The preliminary technology for the development of suc

honeycomb is presented in Fig. 6. It consist of four steps:

1. the production of composite tubes by using lamination technology and vacuum bag

curing process; tubes are then subsequently cut to obtain triangular elements;

2. the production of a positioning tool for the assembly, which is conveniently carried

out by applying Fusion Deposition Molding, a Rapid Prototyping technique;

3. the assemblage in a special mold, endowed with silicon rubber inserts that are

separately produced by u

adhesive paste is set on the adjacent sides of the triangular elements;

4. bonding, which is carried out at room temperature under a press, and subsequent

filling of with liquid resin, to increase th

reduce the weight of the honeycomb.

Fig. 6 - Technology for the production of hexa

Production of composite

triangular tubes (Lamination

+ Vacuum Bag Technology)

Production of glass



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Generation of an hexa-chiral tessellation by superposition of triangular elements

Such configuration originates a hexa-chiral topology with polygonal nodes, which is expected


possible in order to obtain the desired auxetic response.

The preliminary technology for the development of such new type of composite chiral

honeycomb is presented in Fig. 6. It consist of four steps:

the production of composite tubes by using lamination technology and vacuum bag


oduction of a positioning tool for the assembly, which is conveniently carried


the assemblage in a special mold, endowed with silicon rubber inserts that are

separately produced by using a liquid siliconic rubber; in the assembly process, an


bonding, which is carried out at room temperature under a press, and subsequent

filling of with liquid resin, to increase their stiffness; nodes are subsequently drilled to

reduce the weight of the honeycomb.

Technology for the production of hexa-chiral network with polygonal nodes

3D printed positioning tool

Production of glass- and carbon –fiber reinforced

units

Assembly mould for bonding

(elastomeric tooling with

silicon rubber inserts)

chiral tessellation by superposition of triangular elements

al topology with polygonal nodes, which is expected


h new type of composite chiral

the production of composite tubes by using lamination technology and vacuum bag


oduction of a positioning tool for the assembly, which is conveniently carried


the assemblage in a special mold, endowed with silicon rubber inserts that are

sing a liquid siliconic rubber; in the assembly process, an


bonding, which is carried out at room temperature under a press, and subsequent

eir stiffness; nodes are subsequently drilled to

network with polygonal nodes

3D printed positioning tool

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Preliminary tests proved that the new technology is feasible and can obtain hexa-chiral

networks with good strength characteristics. Therefore, such process can be applied to

produce the ribs or a core that will constitute the internal deformable structure of the

morphing flap that will be proposed in the subsequent section.

Composite corrugated panels for morphing skin systems

A fundamental component of any type of morphing structure is constituted by the skin, which

must collect the aerodynamic loads and transfer them to the main structure. Moreover, an

efficient morphing skin should also perform additional structural roles. Such considerations

derives from the role of the skin in conventional aircraft stressed-skin constructions. Since

morphing applications typically require shape changes only in some directions, a deformable

skin with strongly anisotropic characteristic would be able to perform traditional structural

roles in non-morphing directions, thus limiting the weight costs derived from the introduction

of morphing capabilities in the structure.

Corrugated laminates can be considered promising solutions for supporting a morphing

skin, since they are characterized by inherent anisotropy (Yokozeki, 2005, Thill, 2010,

Fournier, 2013). However, they do not provide a smooth and continuous cover and this

lead to detrimental effects in the aerodynamic performances, as it has been proved by

recent works (Xia, 2014).

The solution proposed in (Fournier, 2013) can be considered particularly effective to

integrate an elastomeric layer, bonded to the external horizontal sides of the corrugated

support. In the “valleys” of the corrugation, the skin is sustained by aramid honeycomb

inserts, as it is shown in Fig. 7-a

In the work presented in (Fournier, 2013), a composite corrugated laminate was first

manufactured with an autoclave, by using a mould and an elastomeric counter-mould

shown in Fig. 7-b.

Fig. 7 - Skin system based on corrugated support and elastomeric supported cover (a) and mold for the

production of the corrugated composite panel (b)

The technological process is summarized in Fig. 8. First, the composite corrugated panel

is produced by using a vacuum bag technology. Then honeycomb inserts are cut and

positioned in the valleys of the corrugate, which is supported by the original mold. The

mold is then capsized and introduced in a vulcanization mold that contains a liquid

silicon rubber. During vulcanization, the rubber penetrates inside the cells of the

honeycomb by capillarity and, at the same time, adhesion is obtained between the rubber

layer and the horizontal sides of the corrugated support, thanks to the application of a

special primer.

(a) (b)



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Fig. 8 - Production and testing of a skin system

The produced skin system can be integrated in the morphing structure. The test shown in Fig.

8 indicates that, during elongation, some waviness is obtained, but the irregularities are one

order of magnitude lower than the ones inherent to the corrugation. In such conditions, small

detrimental effects on the aerodynamic performances are expected (Thill, 2010, Xia, 2014).

This type of corrugated skin is suitable to be adopted in the design of a morphing flap and will

be considered in the following section to study the performance of such innovative high lift

device.

IDENTIFICATION AND VERIFICATION OF AN ACTIVE MORPHING FLAP

SOLUTION

Generation of morphing flap configurations

The general configuration of the morphing flap is shown in Fig. 8. The basic idea is to

introduce a beam that support the chiral structure of the flap, between the slider linked to the

extraction mechanism and the flap structure. Such beam pivots around the sliding support,

which moves along the track fixed to the wing body. The pivoting beam carries the flap

through two hinges, where actuators can be connected to the chiral core. Accordingly,

application of torques to the chiral core induces the morphing of entire flap, without

interfering with the extraction mechanism.

In this section, the structural configuration is identified with more details and a finite element

model of the morphing flap is developed.

The design case previously described is adopted as reference case. Flap chord is 896 mm. The

position of the two actuation points is fixed at 19.5% and 42% of the chord. The load

conditions are taken from the results presented in (Benini, 2011). Therefore, they are

Production of composite

corrugated panel

Positioning of honeycomb inserts

Vulcanization of elastomeric layer

Skin system ready for

testing

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rigorously valid only for the original shape of flap, but they can reasonably used for a

preliminary investigation of the concept.

Fig. 9 - Morphing flap solution

For such analyses a flap span of 500 mm is considered. Indeed, two different structural

layouts can be based on the scheme represented in Fig. 8. The first solution considers that the

internal chiral core is continuous along the entire flap span, whereas in the second solution the

chiral supporting structure is lumped at two discrete ribs.

In both cases, preliminary numerical analyses are conducted by modelling a strip of the chiral

profile, endowed with the deformable skin. However, for the first solution, a strip width of 20

mm is adopted and the applied aerodynamic forces corresponds to ones exerted on such span,

whereas the preliminary analyses for the solutions with ribs are carried out considering an 80

mm-wide rib, which are subjected to the aerodynamic loads collected over a span of 500 mm.

The identification of the most promising solutions was performed by using a parametric

model of the morphing flap. The model is generated through a Matlab® tool that automates

the mesh generation procedure for both the internal rib and the external skin and that produces

a NASTRAN input file, ready for finite element analysis.

Fig. 10 - Geometrical parameters of the chiral core (a) and the corrugated skin (b)

The auxetic internal structure is modeled through composite shell elements with a symmetric

lamination sequence of 0/90 fabric plies, with the possibility to choose between carbon/epoxy

(a) (b)

Actuation torque

Track fixed to the wing body

Extraction mechanism

Pivoting beam hinged to the flap @ actuation locations



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(CFRP) or a glass/epoxy (GFRP) pre-pregs. A reference thickness is given in input, and the

tool will calculate how many plies to put, depending on the ply thickness. It should be

remarked that, in the preliminary study presented, the entire chiral network has the same

thickness.

Fig. 11 - Flowchart for automatic model generation of morphing flap configurations

The chiral nodes are partially filled in the model, analogously to the chiral units that are

shown in Fig. 6. Filling is obtained by using solid elements with material properties that can

be considered representative for an epoxy resin.

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The automatic tool is able to fill the internal space within the airfoil with a uniform hexa-

chiral tessellation with polygonal nodes. The geometrical parameters that can be varied are

included in Fig. 10-a.

The geometry of the corrugated panel is approximated by considering a series of circular

segments, as shown in Fig. 10-b. The corrugations is represented in details, without adopting

an homogenized plate model, by using shell elements with the properties of the composite

laminate. The elastomeric cover and the supporting honeycomb, which are included in the

technological solution presented in Fig. 8, are represented by a shell element with isotropic

properties and by an orthotropic tridimensional element, respectively.

The operations performed by the Matlab tool for automatic mesh generation are presented in

Fig. 11 and the properties of the materials involved in the generation of the model are

reported in Table 1.

In this preliminary study, several configurations have been investigated to analyses their

response under the application of pressure loads given in (Benini, 2011). In particular, two

type of analyses will be performed:

- an evaluation of the passive response, which regards the shape modifications induced

by nominal aerodynamic loads is the flap is left free to morph without actuation

moments, in different constraint conditions;

- an evaluation of the active response, which is focused on the camber variation that can

be obtained by applying actuation moments without exceeding the structural strength

of the ligaments.

Table 1 - Mechanical properties of materials used in morphing flap model

Materials E11

(Mpa)

E22

(Mpa)

E33

(Mpa)

G12 G13 G23 V12 V13 V23

CFRP 56550 56550 10000 4500 3000 3000 0.05 0.3 0.3

GFRP 22000 22000 8000 5500 4000 4000 0.11 0.3 0.3

Resin (chiral

node filler) 3000 3000 3000 1290 1290 1130 0.33 0.33 0.33

Elastomeric skin 50 50 50 18 18 18 0.4 0.4 0.4

Honeycomb in

skin system 0.01 0.01 138 0.0038 40 25 0.3 1e-04 1e-04

Several configurations have been investigated and a representative one, with the properties

reported in Table 2, has been selected to show the potential of the proposed morphing

solution. The selection of configurations has been based on the possibility to achieve a

smooth and aerodynamically efficient shape under actuation and on the maximum camber that

can be achieved without endangering structural integrity.

The representative solution adopt carbon/epoxy plies for both the chiral core and the skin. The

thickness of the chiral ligaments is 0.8 mm, whereas the thickness of the corrugated laminate

is 0.6 mm.



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Table 2 - Geometrical parameters of the selected representative configuration

Parameter Description Value r Chiral node radius 5 mm

rf Triangle fillet radius 0.4 mm

pci Chiral forward end, in chord percentage 0%

Pcf Chiral rearward end, in chord percentage 70%

rred Chiral exagon apothem, percentage of node radius 60%

Rcup Corrugated upper radius 6 mm

Rcdown Corrugated lower radius 6 mm

Lc Half- Length of corrugated vertical flange 1 mm

Ofst Offset between corrugated mean plane and reference plane 7 mm

ϑ Slope between chiral nodes and triangles face 25°

R Distance between two chiral nodes 24 mm

L Triangle side 22 mm

l Length of superposed triangle’s faces 15 mm

Passive and active responses of a representative configuration

The analysis of the passive response is aimed at verifying that the morphing flap, with an

auxetic internal structure and a corrugated skin, can withstand the aerodynamic pressures,

without problems of structural safety and without detrimental effects on the aerodynamic

profile shape.

Both the configuration with continuous chiral core and discrete ribs have been considered,

with the aerodynamic pressure taken from the landing condition of (Benini, 2011), which is

the most severe. Two constraints conditions have been adopted by fixing and allowing the

rotations at the hinges between the pivoting beam and the internal chiral structure,

respectively.

Figure 12-a shows the passive response of the case with continuous chiral core clamped to the

pivoting beam. The solution with the continuous core is significantly stiffer than the one with

discrete ribs, as it was expected by simple structural considerations. Therefore, the

configuration that shown in Fig. 12-a is the one with the minimal shape variations among the

considered cases. On the contrary, Fig. 12-b, which is referred to the case with discrete ribs

hinged to the pivoting wing, shows the case that experiences the highest deflections.

Quantitative results, regarding the displacements of trailing edge (TE) and leading edge (LE)

and maximum stress in chiral structure, are reported in Table 3.

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Fig. 12 - Passive response for the solution with continuous chiral core clamped to the pivoting beam (a) and with

discrete chiral ribs hinged to the pivoting beam (b)

Table 3 - Passive response of a representative morphing flap

Case Constraints TE displacement

(mm)

LE displacement

(mm)

Max ply stress

(Mpa)

Continuos chiral

core

Clamped 7.51 0.41 57.2

Continuos chiral

core

Hinged 9.51 1.03 50.1

Discrete ribs Clamped 13.36 1.30 221.0

Discrete ribs Hinged 16.76 1.50 155.8

It can be observed that the displacements obtained in the different cases are not negligible.

This fact underline that a proper design must be performed considering aeroelastic

interactions, even if the morphing structure is not free to rotate at the hinges.

(a)

(b)



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However, thank to the properties of the chiral core, a smooth shape change is obtained under

the action of the aerodynamic force. Moreover, stress levels are below the maximum stress

that can be applied to carbon/epoxy composite plies.

The active response of the morphing flap has been investigated by applying torques in

correspondence of the hinges that connect the chiral structure to the pivoting beam.

This torques can restore the initial shape of the flap under the action of aerodynamic loads.

Calculations show that by applying 120 Nm on the rear hinge and no torque on the forward

hinge, the displacements induced by the aerodynamic forces are almost eliminated in the

solution with continuous core. The following results are obtained: 0.01 mm and 0.274 mm of

displacement at LE and TE, respectively, and a maximum stress of 168.7 MPa in the chiral

core. For the solution with the discrete 80 mm-wide rib, the required torques are 197 Nm on

the rear hinge and no torque on the forward hinge. This leads to 1.3 mm on the TE and – 0.03

mm on the LE, with a maximum stress of 598.7 MPa.

If the torques are increased, the camber of the airfoil is increased with respect to the original

unreformed configurations. However, it is important to consider that the torques work against

the aerodynamic forces, so that the reference configuration to evaluate the maximum

achievable camber should be evaluated considering the deformed configuration that was

previously described, in particular the one obtained in the free hinge conditions. Table 4

reports the results obtained by increasing the rear torque up to 450 Nm and to 300 Nm for the

case of continuous core and discrete ribs, respectively. The forward hinge is left free to rotate.

Figure 13 shows the deformed shapes and the contour of stress in the chiral ligaments for the

two considered solutions.

Table 4 - Active response of a representative morphing flap

Case Torque

applied to

rear hinge

(Nm)

TE displ

(mm) ∆∆∆∆ TE (mm)

w.r.t. free

hinge

Max ply stress

(Mpa)

Continuos

chiral core

450 -27.55 -35.06 720.7

Discrete ribs 300 -8.7 -25.46 959.1

It can be observed that the maximum stress are beyond the safety limits for carbon/epoxy

fabric plies. However, Fig. 13 shows that such maximum stresses are obtained only in few

localized spots. The possibility of varying selectively the thickness of the ligaments and to

balance in the best way the stress experienced by the core and the skin of the morphing flap is

likely to improve the state of stress obtained in this conditions, without affecting too much the

overall performance.

Moreover, it has to be remarked that the structural architecture adopted allowed obtaining

very smooth shape changes also for the case of the actuated solutions, in the presence of

concentrated torque applied to the structure.

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Fig. 13 - Active response for the solution with continuous chiral core (a) and with discrete chiral ribs (b)

CONCLUSIONS

The activities presented in this work define a possible solution to endow a single slotted flap

with morphing capabilities. The studies moved from two optimal configurations, which were

selected in literature for takeoff and landing conditions. Indeed, they do not differ very much

for the shape, but this is consequence of the optimization process carried out, which did not

take into account the possibility to change the shape of the trailing edge, since this would

have had a detrimental effect on the performance in cruise conditions.

The adoption of a morphing technology allows eliminating such constraint and the analyses

conducted throughout this work indicate that the flap overall camber can be varied with

several advantages regarding the possibility to optimize the aircraft in cruise conditions

during the entire flight. Moreover, a morphing flap can optimize the shape for a large variety

of conditions, including the ones arising from the change of configuration and the extension

of mission envelope for the aircraft.

The results obtained are based on technological processes that have been developed and

assessed. In particular a novel technological process is suggested for the development of thin-

(b)

(a)



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walled composite networks with a chiral topology In parallel, a composite skin concept has

been proposed, which is based on the integration of an elastomeric cover, supported by

honeycomb inserts, in a composite corrugated laminate.

These two structural concept proved to be suited to design a morphing flap that is actuated by

a single torque applied to a rear hinge. The results obtained show the chiral core is able to

control shape variations so that a smooth and aerodynamic efficient profile can be obtained in

various conditions. Moreover, the skin system proved to be able to sustain the aerodynamic

pressures without undergoing excessive deformations. The investigation of a representative

configurations indicate that the morphing flap can obtain a change of camber with trailing

edge displacement of the order 20 mm ÷ 30 mm.

However, margins of improvement of the proposed solutions have to be considered, since the

presented results are referred to a representative and non-optimized configuration, with a

fixed position of actuators and a uniformly thick chiral core.

Since the numerical approach developed is inherently suited to be used for automatic process,

the real potential of the proposed solution could be investigated by properly set up an

optimization algorithm, taking into account aeroelastic interactions and introducing the

possibility of grading the characteristics of the skin and of the chiral core.

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