Composite corrugated structures for morphing wing skin applications

Abstract

Composite corrugated structures are known for their anisotropic properties. They exhibit relatively high stiffness parallel (longitudinal) to the corrugation direction and are relatively compliant in the direction perpendicular (transverse) to the corrugation. Thus, they offer a potential solution for morphing skin panels (MSPs) in the trailing edge region of a wing as a morphing control surface. In this paper, an overview of the work carried out by the present authors over the last few years on corrugated structures for morphing skin applications is first given. The second part of the paper presents recent work on the application of corrugated sandwich structures. Panels made from multiple unit cells of corrugated sandwich structures are used as MSPs in the trailing edge region of a scaled morphing aerofoil section. The aerofoil section features an internal actuation mechanism that allows chordwise length and camber change of the trailing edge region (aft 35% chord). Wind tunnel testing was carried out to demonstrate the MSP concept but also to explore its limitations. Suggestions for improvements arising from this study were deduced, one of which includes an investigation of a segmented skin. The overall results of this study show that the MSP concept exploiting corrugated sandwich structures offers a potential solution for local morphing wing skins for low speed and small air vehicles.

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Composite corrugated structures for morphing wing skin applications
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2010 Smart Mater. Struct. 19 124009
(http://iopscience.iop.org/0964-1726/19/12/124009)
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IOP PUBLISHING SMART MATERIALS AND STRUCTURES
Smart Mater. Struct. 19 (2010) 124009 (10pp) doi:10.1088/0964-1726/19/12/124009
Composite corrugated structures for
morphing wing skin applications
C Thill1,JAEtches
2,IPBond
2,3, K D Potter2and P M Weaver2
1Messier-Dowty, V´elizy, France
2Advanced Composites Centre for Innovation and Science (ACCIS), University of Bristol, UK
E-mail: i.p.bond@bristol.ac.uk
Received 22 June 2010, in final form 28 September 2010
Published 11 November 2010
Online at stacks.iop.org/SMS/19/124009
Abstract
Composite corrugated structures are known for their anisotropic properties. They exhibit
relatively high stiffness parallel (longitudinal) to the corrugation direction and are relatively
compliant in the direction perpendicular (transverse) to the corrugation. Thus, they offer a
potential solution for morphing skin panels (MSPs) in the trailing edge region of a wing as a
morphing control surface. In this paper, an overview of the work carried out by the present
authors over the last few years on corrugated structures for morphing skin applications is first
given. The second part of the paper presents recent work on the application of corrugated
sandwich structures. Panels made from multiple unit cells of corrugated sandwich structures are
used as MSPs in the trailing edge region of a scaled morphing aerofoil section. The aerofoil
section features an internal actuation mechanism that allows chordwise length and camber
change of the trailing edge region (aft 35% chord). Wind tunnel testing was carried out to
demonstrate the MSP concept but also to explore its limitations. Suggestions for improvements
arising from this study were deduced, one of which includes an investigation of a segmented
skin. The overall results of this study show that the MSP concept exploiting corrugated
sandwich structures offers a potential solution for local morphing wing skins for low speed and
small air vehicles.
(Some figures in this article are in colour only in the electronic version)
Nomenclature
caerofoil chord, m
Cddrag coefficient, non-dimensional
Cd0 zero-lift drag coefficient,
non-dimensional
Cllift coefficient, non-dimensional
Clmax maximum lift coefficient,
non-dimensional
Cppressure coefficient, non-dimensional
dthickness of MSP, m
dCl/dαlift curve slope, radians−1
(Exx)lam
equ longitudinal equivalent tensile elastic
modulus of corrugated laminate, Pa
(Exx)san
equ longitudinal equivalent tensile elastic
modulus of MSP, Pa
3Address for correspondence: Department of Aerospace Engineering,
University Walk, Queen’s Building, Bristol BS8 1TR, UK.
(Eyy)lam
equ transverse equivalent tensile elastic mod-
ulus of corrugated laminate, Pa
(Eyy)san
equ transverse equivalent tensile elastic mod-
ulus of MSP, Pa
Lmax maximum lift force that the MSP
demonstrator sees during testing, N
lMSP length of MSP, m
Paxial force applied to MSP, N
Pactuation force, N
pzaerodynamic pressure applied
transversely to each MSP, Pa
saerofoil span, m
tmax aerofoil maximum thickness, m
Uair velocity, m s−1
(uzz)MSP,central central deflection of MSP due to aerody-
namic loads, m
wMSP width of MSP, m
x/cnon-dimensional chord, non-dimensional
0964-1726/10/124009+10$30.00 ©2010 IOP Publishing Ltd Printed in the UK & the USA1

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
αangle of attack, deg
Cd0 change in zero-lift drag coefficient, non-
dimensional
caerofoil chord length change, m
cmax maximum aerofoil chord length change,
m
δtrailing edge control surface deflection
angle, deg
(εxx)lam
equ longitudinal structural equivalent tensile
elastic strain of corrugated laminate, non-
dimensional
(εxx)san
equ longitudinal structural equivalent tensile
elastic strain of MSP, non-dimensional
(εyy)lam
equ transverse structural equivalent tensile
elastic strain of corrugated laminate, non-
dimensional
(εyy)san
equ transverse structural equivalent tensile
elastic strain of MSP, non-dimensional
θaerofoil trailing edge section angle
relative to the horizontal, deg
1. Introduction
Morphing technology applied to air vehicles allows for
aerodynamic efficiency improvement, expansion of flight
envelope and increased mission capabilities (Thill et al 2008b).
Over the last two decades research projects such as NASA’s
Morphing Project (Florance et al 2003, Bartley-Cho et al 2004,
Kudva 2004), DARPA’s Smart Wing programme (Florance
et al 2003, Bartley-Cho et al 2004, Kudva 2004) and Morphing
Aircraft Structures programme (Love et al 2004,Bowmanet al
2007, Bye and McClure 2007, Love et al 2007) developed and
applied morphing concepts in technology demonstrators for
wind tunnel and flight tests of small unmanned air vehicles.
A major challenge was the design of a skin material that
combines the two inherently different mechanical properties:
stiffness and compliance. A morphing skin needs to allow
large shape change while also being able to resist deformation
under aerodynamic loading. These requirements can be
treated separately when the compliant and stiff directions
differ. Such an application could be a morphing trailing edge
control surface of a wing. In such an application the skin
needs to allow camber and/or chord length change in the
chordwise direction and can be stiff in the spanwise direction
to resist the applied loading. Thus, a morphing skin material
or structure needs to be developed that exhibits orthotropic
stiffness properties.
Corrugated structures exhibit such properties: they are
stiff longitudinal to the corrugation direction and compliant
in the transverse direction (Thill et al 2007). Corrugated
structures in general are not new to the engineering or
natural world and have been used over many decades in civil,
naval, automotive and aerospace applications due to their
efficient performance, i.e. low mass to stiffness (longitudinal to
corrugations) ratio, buckling resistance and energy absorption
(Shanley 1947, Perel and Libove 1978, Wiggenraad et al 2001,
Sun and Spencer 2005). In Nature, numerous insects, such
as the damselfly, desert locust and dragonfly, have corrugated
wings for deployable mechanisms, stiffness requirements
and/or aerodynamic effects (Wootton 1981,1990,1992,
Herbert et al 2000, Wootton et al 2000). Yokozeki et al
(2006) were among the first to suggest corrugated laminates
made from carbon/epoxy composites for morphing wing skin
applications.
The present authors have investigated the corrugated
concept in more detail. An overview of this work is given in
the next section. The second part of this paper presents work
on the application of these corrugated structures when used
as morphing skin panels (MSPs) in the trailing edge region
of a scaled morphing aerofoil section. The aerofoil section
features an internal actuation mechanism that allows chordwise
length and camber change of the trailing edge region (aft 35%
of the chordwise section). Wind tunnel testing was carried
out to demonstrate the MSP concept but also to explore its
limitations. Suggestions for improvements arising from this
study were deduced, one of which, involving a secondary skin
on top of the MSP, was investigated.
2. Composite corrugated structures for morphing
skins: previous work
In a preliminary study (Thill et al 2007,2010b), tensile and
flexural experimental testing on various corrugated laminate
designs, both longitudinal and transverse to the corrugations,
was undertaken. The corrugated specimens were manufactured
using a variety of woven fibre/epoxy prepreg materials (aramid,
glass and carbon) and variables such as number of plies and
corrugation pitch were investigated. The aim was to investigate
the effects of constituent material and corrugation geometry on
the overall mechanical properties of the structure. The results
highlight the extreme anisotropic nature of these structures
and the primary dependence of the transverse tensile elastic
modulus on the laminate thickness. It was also found that the
use of pre-impregnated tape lay-up is impractical for consistent
corrugated laminate manufacturing quality. Resin transfer
moulding (RTM) processes were suggested as an alternative to
address some of the challenges and improve laminate quality
and reproducibility in a production environment. However,
developing such methods was considered beyond the scope of
this project and a manual manufacturing process was pursued.
A fundamental drawback of corrugated laminates is their
very low out-of-plane stiffness which is likely to be unable
to resist the aerodynamic loading at typical speeds required
for flight. Hence, a corrugated sandwich structure was
suggested wherein two corrugated laminates are bonded to a
foam core. In a second study, Thill et al (2008c)presented
experimental and analytical results for a sandwich structure
with trapezoidal corrugated skins. Furthermore, a parametric
study analysed the effects of corrugation geometry and form
(re-entrant, rectangular, round, trapezoidal, sinusoidal and
triangular) on the structural elastic strain and equivalent elastic
tensile modulus transverse to the corrugation direction. It
showed that for a rectangular corrugation, structural strains
of up to 50% can be realistically achieved if the stiffness is
not of primary concern. Conversely, corrugated laminates
with a transverse equivalent elastic modulus of about 4 GPa
2

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
Initial design
dimensions & materials
Analysis
aerodynamic, structural
& kinematic
MSP design
unit cell dimensions
Revised design
dimensions, materials &
manufacturing
Figure 1. Design iteration routine.
can be designed. It has been shown that for fixed projected
dimensions, re-entrant corrugations give the highest strain
while triangular corrugations provide the highest modulus. The
results give a good indication of the wide envelope within
which corrugated laminates can be designed to meet the
requirements for morphing skin applications.
While corrugated structures offer a potential structural
solution for morphing wing skin applications, it seems to be
an aerodynamic drawback to have non-smooth wing surfaces.
Hence, aerofoils with corrugated trailing edges have been
studied experimentally and computationally using various
corrugation geometries and forms (sinusoidal, trapezoidal and
triangular) at different Reynolds numbers (Thill et al 2010a).
The investigation showed that the aerodynamic performance
is highly dependent on corrugation amplitude, wavelength,
gradient (combination of amplitude and wavelength) and
Reynolds number. Evidence is given highlighting that
penalties for having a non-smooth trailing edge surface can be
eliminated for the lift curve slope and minimized for the zero-
lift drag coefficient by adequately designing the corrugation
geometry and choosing the operating Reynolds number range.
One such consideration is to recess the corrugations below the
aerofoil profile line. A segmented skin on top of the corrugated
profile was suggested for the suction surface to reduce the drag
penalties still further.
The results of these earlier studies led to an overall
conclusion that composite corrugated sandwich structures
could offer a potential solution for morphing wing skin panels
for slow speed, lightweight and small air vehicles. One
particular application could be as a continuous multi-purpose
Figure 2. Final design of MSP demonstrator front aerofoil section
without skin.
trailing edge control surface panel. This concept is to be
proven in this study through the design, manufacturing and
wind tunnel testing of an MSP demonstrator.
3. Methodology
3.1. Design and analysis of MSP demonstrator
Several design iterations of the demonstrator layout and
material selection (figure 1) were carried out before the final
design was determined (figures 2and 3). A NACA 0024
aerofoil profile was selected for the following reasons.
•A simple and relatively cheap actuation system needs to be
fitted into the aerofoil section at about 65% chord hence at
that position adequate aerofoil thickness of about 100 mm
is required.
•The larger the corrugations of the morphing skin panel
are the easier it is to make using in-house hand lay-
up processes. This means about 5–10 mm in height;
furthermore, the height of one corrugated cell should
be around 1% chord or less for acceptable aerodynamic
performance and realistic scale (Thill et al 2008a). Hence,
thechordneedstobelessthan1m.
•The demonstrator has to fit in a wind tunnel working
section of 1.52 m ×1.52 m ×2.13 m (5 ft ×7ft×7ft)
by keeping blockage effects to a minimum, hence keeping
the overall size of the demonstrator as small as possible.
Since the chord is limited to less than 1 m the span cannot
be more than 1.52 m (5 ft) to avoid excessive wind tunnel
blockage.
The above meant that a relatively thick aerofoil profile had
to be selected for which aerodynamic data were available. The
NACA 0024 profile fits the requirements. The final overall
dimensions of the demonstrator wing section are a span, s=
1.520 m, a chord, c=0.800 m, and a maximum thickness at
30% chord, tmax =0.192 m.
The angle of attack of the aerofoil section, the morphing
trailing edge control surface deflection angle (camber) and
3

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
Figure 3. Final design of MSP demonstrator front aerofoil section with skin (left) and actuation system in between MSPs (right).
chord length can all be altered. The actuation system (figure 3)
of the trailing edge control surface consists of two cross-bar
elements that can be extended or retracted by turning a lead
screw manually using a torque wrench. This allows the chord
to be changed, c, to any length within the elastic limit of the
MSP. The camber can be changed by removing a locator pin
and adjusting the trailing edge deflection angle δaround the
65% chord position. Seven fixed positions are possible (0◦,
±6◦,±12◦,±18◦). Furthermore, 30 static pressure tappings
on one surface ahead of the MSP are incorporated in the
demonstrator to assess the flow behaviour ahead of the MSP
i.e. from 0% to 65% chord. The pressure tappings consist
of brass tubes with an outer diameter of 1.59 mm and a wall
thickness of 0.36 mm.
In order to find the maximum aerodynamic forces applied
to the MSP demonstrator, XFOIL (open source code), a
two-dimensional computational fluid dynamics (CFD) panel
method with viscous effects (Drela 1989), is used to carry out
an aerodynamic analysis at a Reynolds number ∼2×106and
Mach number ∼0.1 over a range of angles of attack for the
conventional NACA 0024. XFOIL predicts that about 2% of
the total lift is generated aft of 65% chord i.e. over the trailing
edge region. For the following analysis it is assumed that
5% of the total lift acts aft of 65% chord. This assumption
is conservative and simplified; it is used in the analysis to
dimension the structure and actuation mechanism.
The structural analysis is carried out only for some critical
parts of two load cases. It uses linear-elastic stress–strain
relationships; bending, torsion and shear analysisof closed thin
sections; bending and buckling analysis of thin plates; bending
and buckling analysis of beams; lug and thread analysis.
The relevant dimensions from the initial design are input
into a spreadsheet from where they are analysed by Matlab
(version 7.2.0.232; R2006a) to provide outputs of minimum
reserve factors for each analysed part and load case in a new
spreadsheet. A safety factor of 1.5 was used throughout. The
kinematic analysis of the actuation system yielded the forces
required to deform the MSP as a function of the cross-bar angle
and the resulting in-plane deflections.
During the design of the MSP the aim was to optimize
the geometry and form of the corrugated sandwich structure in
Figure 4. Schematic of MSP location relative to horizontal (x-axis).
terms of stiffnesses in the chord- and spanwise directions. The
maximum structural strain expected from the MSP is given by
(εyy)san
equ cmax
cos θwMSP
.(1)
The angle θcomes from the fact that the MSP is at an
angle relative to the y-direction which is dependent on the
aerofoil profile (figure 4; for NACA 0024 θ=16.5◦) while
cmax is equal to the elastic limit of the MSP. The maximum
force Papplied to each MSP depends on the actuation force
P(figure 4) and hence determines the maximum allowed
equivalent transverse tensile elastic modulus of the MSP:
(Eyy)san
equ P
dlMSP
1
(εyy)san
equ
.(2)
A two-dimensional beam bending analytical analysis and
finite element analysis (FEA) using ABAQUS/CAE (version
6.6-1) are carried out to determine the maximum central
deflection. The MSP is assumed to withstand the aerodynamic
pressure forces (distributed transverse load: lift force over
aerofoil section lifting area). Note that in reality the upper
panels take more load than the lower panels, however, for
simplicity both upper and lower panels are treated similarly.
The pressure is calculated to be
pz=0.05Lmax
cs .(3)
4

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
Figure 5. Schematic of MSP corrugated unit cell.
Table 1. MSP geometry.
s
(mm)
t
(mm)
b
(mm)
h
(mm)

(deg)
b1
(mm)
b3
(mm)
12.0 0.3 6.0 4.9 0.0 2.0 4.0
e
(mm)
ncor
(—)
d
(mm)
LUC
(mm)
nUC
(—)
wMSP
(mm)
lMSP
(mm)
1.5 2 10.0 15.0 14 210 760
The maximum central deflection of the MSP based on simple
beam bending theory (clamped–clamped boundary conditions)
is (Young 1989)
(uzz)MSP,central =12 pzw4
MSP
384(Eyy)san
equd3.(4)
Since the equivalent elastic strain and modulus have been
fixed, the detailed dimensions of the MSP unit cell can be
back-calculated. The MSP is a corrugated sandwich panel
as described in Thill et al (2008c). Its equivalent moduli
in the transverse (yy) and longitudinal (xx) directions are
calculated using the analytical models for the corrugated
sandwich structure (Thill et al 2008c). The corrugation form
and geometry, and sandwich geometry must be manipulated
until the transverse equivalent structural elastic strain and
modulus of the MSP equal the allowed values from (1)and(2),
respectively. This could notbe matched exactly due to practical
limitations of the dimensions of the MSP. Note that from Thill
et al (2008c) it is best to have a re-entrant corrugation form for
a low modulus/high strain MSP, however, these are difficult to
manufacture hence the next best option is chosen: rectangular
corrugations (figure 5). Due to the in-house manufacturing
limitations, two panels were used for each surface (top and
bottom) i.e. four panels in total. The final dimensions and
properties of each MSP are given tables 1and 2, respectively.
3.2. Manufacturing
The corrugated laminates were manufactured from pre-
impregnated E-glass/977-2 epoxy tape (Cytec, UK) that was
Figure 6. Picture of laying-up process of prepreg onto tool.
Figure 7. Picture of rectangular corrugated sandwich structure.
Table 2. Mechanical properties of MSP.
(Exx)lam
equ
(GPa)
(εxx)lam
equ
(%)
(Exx)san
equ
(GPa)
(εxx)san
equ
(%)
61.36 1 3.83 1
(Eyy)lam
equ
(MPa)
(εyy)lam
equ
(%)
(Eyy)san
equ
(MPa)
(εyy)san
equ
(%)
13.08 21 0.82 17
hand-laid onto an aluminium tool (figure 6). The prepreg,
sandwiched between layers of release film, was formed
corrugation by corrugation with the help of a heat gun to ease
forming of the prepreg onto the tool which has rectangular
grooves. Rectangular aluminium bars were located on top
and weights were used to hold the complete assembly in place
until the vacuum bagging process. After applying vacuum to
the entire tool, the assembly was cured at the recommended
prepreg cure cycle (180 ◦C for 3 h at 689.5 kPa) in the autoclave
(Quicklock Thermoclave, Leeds & Bradford Boiler Company
Limited). The properties of the prepreg material used are given
in table 3.
Two types of corrugated sandwich panels were manufac-
tured: one without and one with an external discontinuous
segmented skin.
Two laminates manufactured from one ply of pre-
impregnated tape E-glass/977-2 epoxy (table 3) were made per
corrugated sandwich panel. The laminates were then bonded
to a foam core (Rohacell®31) using a two part epoxy paste
adhesive (Redux 810, Hexcel Composites). To accelerate the
cure of the adhesive the assembled sandwich structure was
put in an oven for 1 h at 70◦C. Figure 7shows a finished
rectangular corrugated sandwich structure.
5

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
Figure 8. Schematic (left) and picture (right) of segmented skin bonded to corrugated laminate.
Table 3. Properties of prepregs.
Va
fρaE11 =Eb
22 Gc
12 νc
12
Fibre type Weave Resin (%) (kg m−3) (GPa) CoV (GPa) (—)
Glass Five harness satin Cytec 977-2 42.7 2080 23.3 2% 5 0.2
aData from manufacturer. bData from experiment. cEstimated.
Figure 9. Picture of morphing skin panel (MSP).
For the corrugated sandwich panel with an external skin,
first the panel was manufactured as described above. Then
20 mm wide strips (segments) made from one ply of pre-
impregnated tape E-glass/977-2 epoxy (table 3) were bonded
on to one surface of the panel using a two part epoxy paste
adhesive (Redux 810, Hexcel Composites). To accelerate the
cure of the adhesive the assembled sandwich structure was put
in an oven for 1 h at 70 ◦C. The strips are bonded as shown in
figure 8to form a segmented outer skin.
The individual components of the demonstrator were
either procured commercially or manufactured in-house. The
MSPs were manufactured as described in the previous two
sections. On each end an aluminium connector was bonded
using a two part epoxy paste adhesive (Redux 810, Hexcel
Composites) to the laminates to allow installation of the MSPs
on the demonstrator. A completed MSP is shown in figure 9.
Note that the corrugations are recessed below the aerofoil
profile line as recommended by Thill et al (2008a). It has
nominal projected dimensions of 760 mm ×258 mm×10 mm.
The MSPs were fastened through their connectors to the
secondary and tertiary spars of the demonstrator using M5
socket head countersunk screws.
3.3. Experimental testing
Two types of experimental tests were undertaken: in and out
of the wind tunnel. The out of wind tunnel testing validated
the structural and kinematic integrity of the MSP demonstrator.
This comprised taking the actuation system through its entire
range of movement of trailing edge control surface chord
length and camber change to guarantee kinematic reliability
in the wind tunnel.
The wind tunnel testing was carried out in the University
of Bristol large low speed wind tunnel which has an octagonal
test section with dimensions 2.13 m (7 ft) by 1.52 m (5 ft) and
a maximum speed of 60 m s−1. The effects of lift, drag and
pitching moment coefficients on the aerofoil with extended and
deflected morphing trailing edge control surface at different
angles of attack and wind speeds were investigated. The actual
forces and moments of the MSP demonstrator were measured
using an OR6-7 2000 series force platform (AMTI) which uses
strain gauges to measure three force and three moment compo-
nents. It is interfaced via a six-channel strain gauge amplifier
(MSA-6 MINIAMP, AMTI) to a dSpace (DS1103/CLP1103)
unit using six BNC connectors. The air velocity is recorded
via a FCO510 Micromanometer (Furness Controls Limited)
connected to the dSpace unit via a RS232 connector. The
dSpace unit is connected to a PC data logger via ‘ControlDesk’
software with data acquired at a rate of 100 Hz for 10 s.
The static surface pressures on one surface of the aerofoil
section ahead of the MSP were measured electronically using
piezoresistive pressure transducers. The measurement and data
acquisition system (Scanivalve Corporation) uses an electronic
pressure scanner (ZOC22B32PxX2) connected via tubes to the
pressure tappings, an analogue to digital converter (RAD3200)
connected to a PC via USB and a power supply (RPM 1000).
Data are acquired via ‘Radlink’ software at a rate of 100 Hz for
10 s.
The MSP demonstrator is mounted vertically in the wind
tunnel to minimize blockage effects and to make interfacing
with the force platform easier. The force platform is attached
to the demonstrator via an aluminium interface to one end of
the secondary spar. The force platform and the interface are
located underneath the wind tunnel test section floor. In order
6

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
Figure 10. Pictures of out of wind tunnel tests.
to ensure two-dimensional flow over the majority of the MSP
demonstrator, the gap between each of its ends and the side
walls of the wind tunnel was restricted to 2.5 mm (less than
0.2% of the full height of the tunnel). Note that the forces
and moments are usually measured at the quarter chord point,
however, for stability reasons in this case, the measurements
were taken at the secondary spar location (about 65% chord)
which coincided with the centre of gravity. Note that the
calibration factors of the force platform and the amplifier were
taken into account. The wind tunnel data were also corrected
for free stream conditions using three theoretical correction
factors to account for streamline curvature, solid and wake
blockage. These were applied using methods suggested by
Barlow et al (1999).
Five tests (runs) were carried out for the fully retracted
configuration, while a single run was carried out for the other
configurations. Data points given in this article are averaged
over the 10 s data acquisition interval for each run; the
corresponding coefficient of variation is calculated.
4. Results and discussion
The out of wind tunnel testing proved the integrity of the
actuation system and the MSPs under deformation. Figure 10
shows the morphing trailing edge section of the demonstrator
(aft 35% chord) with the upper and lower MSPs as it is
stretched and deflected. In these figures the tip is displaced
up to 4% chord and deflected up to 12◦. Note that the
lower surface bulges inwards when it is not under tension.
If this lower surface is the pressure surface, no significant
aerodynamic penalty is expected, although more detailed
studies would have to be undertaken for verification.
The repeatability of the five wind tunnel tests for the
fully retracted configuration is shown in figures 11 and 12 for
free stream velocities of 20 and 30 m s−1. Generally, good
repeatability is achieved for both cases, however for the case
at the higher velocity the drag data are inconsistent. This
is attributed to interference of the MSP demonstrator with
the wind tunnel walls that contract during operation. Hence,
the aerodynamic data for the higher velocity case need to be
treated with care as absolute values cannot be taken as being
representative but relative changes are comparable.
Since no published experimental data for an aerofoil like
the MSP demonstrator are publicly available it is difficult to
Figure 11. Polar plot of lift and drag coefficients for MSP
demonstrator at U=20 m s−1.
Figure 12. Polar plot of lift and drag coefficients for MSP
demonstrator at U=30 m s−1.
validate the present experimental results. As a reference the
data for a conventional NACA 0024 aerofoil can be used for a
preliminary validation. These are taken from published data
(Sarpkaya 1973) at a Reynolds number of 0.82 ×106and
from XFOIL including viscous effects at a Reynolds number
of 1.1×106and Mach number of 0.06 over a range of angles
of attack. Figure 13 shows the lift coefficient against angle of
attack curves for the XFOIL data for the conventional NACA
0024 and the experimental data for the MSP demonstrator. The
data for three cases are presented when the morphing trailing
edge control surface is retracted (δ=0◦;c=0), extended
(δ=0◦;c=2.5%c), and deflected (δ=12◦;c=0). For
7

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
Figure 13. Validation of experimental results: lift coefficients against
angle of attack for NACA 0024 from XFOIL and for MSP
demonstrator from experimental results at U=20 m s−1.
Figure 14. Effect of morphing trailing edge control surface chord
length on lift coefficients against angle of attack at U=20 m s−1.
the same cases table 4compares lift curve slopes, maximum
lift coefficients and zero-lift drag coefficients for the XFOIL,
published and experimental data.
From these results it can be concluded that the lift curve
slope does not change significantly as the control surface
is extended or deflected and the maximum lift coefficient
increases as the surface is deflected. The zero-lift drag
coefficient increases as the surface is deflected as predicted by
the XFOIL results. These results agree well with published
data for conventional trailing edge control surface deflections
(Kermode 1996). In terms of validation of the experimental
data with the XFOIL and published data, the lift curve slope
and maximum lift coefficient agree well, however, the zero-
lift drag coefficients have been measured to a lesser degree
of agreement. The possible reasons are the wind tunnel wall
interference as was discussed above and the corrugated trailing
edge region as was shown in Thill et al (2008a).
Figures 14 and 15 reiterate the same findings by showing
the experimentally determined lift coefficients of the MSP
demonstrator with a morphing trailing edge control surface
with respect to changing chord length and surface deflection.
No changes are observed when changing the chord length
while the maximum lift coefficient increases with control
surface deflection.
From figures 14 and 15, a further observation can be
made: the stall angle decreases when the morphing trailing
Figure 15. Effect of morphing trailing edge control surface
deflection on lift coefficients against angle of attack at
U=20 m s−1.
Figure 16. Coefficients of pressure against non-dimensional chord of
suction surface of MSP demonstrator with morphing corrugated
trailing edge control surface fully retracted.
edge control surface is deflected or extended. This agrees with
the pressure plots for the fully retracted and fully deflected and
extended morphing trailing edge control surface (figures 16
and 17). For the fully deflected and extended case (δ=12◦;
c=2.5%c; figure 17) the pressures on the suction surface
aft of 10% chord at an angle of attack of 15◦fall below the
values of the pressures at lower angles of attack. This is an
indication of separated flow over the suction surface which
decreases the total lift produced. Furthermore, for the fully
retracted case (δ=0◦;c=0%c; figure 16) the flow seems
to remain attached up to 65% chord up to 15◦angle of attack.
This suggests that the flow is still attached when it reaches the
MSP and thus changing its geometry will change the overall
aerodynamic performance of the aerofoil.
The results presented herein and the previous study by
Thill et al (2008a) highlight a major aerodynamic drawback
between a morphing aerofoil with a corrugated and a smooth
trailing edge region: increased generation of drag. One way of
mitigating this effect is a suggestion to introduce an external
discontinuous layer on the MSP suction surface which has
no effect on structural performance. Table 5compares the
lift curve slopes, maximum lift coefficients and zero-lift drag
coefficients of the previously tested corrugated arrangement
against the configuration with a discontinuous segmented skin
for the cases where the morphing trailing edge control surface
8

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
Table 4. Validation of experimental results: lift curve slope, maximum lift coefficient and zero-lift drag coefficient for NACA 0024 from
XFOIL, from published experimental data (Sarpkaya 1973) and for MSP demonstrator from experimental results (coefficient of variation
given in brackets) at U=20 m s−1.
XFOIL data Published experimental data Experimental data
dCl/dα
(rad−1)
Clmax
(—)
Cd0
(—)
dCl/dα
(rad−1)
Clmax
(—)
Cd0
(—)
dCl/dα
(rad−1)
Clmax
(—)
Cd0
(—)
δ=0◦;c=0 4.6 1.2 0.009 5.3 1.0 0.020 5.1 1.1 (5%) 0.023 (10%)
δ=0◦;c=2.5%c5.0 1.3 0.009 5.0 1.1 (5%) 0.018 (7%)
δ=12◦;c=0 4.2 1.6 0.016 5.5 1.4 (4%) 0.024 (6%)
Table 5. Comparison of aerodynamic experimental data of MSP demonstrator without (corrugated) and with (segmented) discontinuous skin
(coefficient of variation given in brackets).
U=20 m s−1U=30 m s−1
dCl/dα
(rad−1)
Clmax
(—)
Cd0
(—)
Cd0
(%)
dCl/dα
(rad−1)
Clmax
(—)
Cd0
(—)
Cd0
(%)
δ=0◦;c=0 Corrugated 5.1 1.1 (5%) 0.0232 (10%) 4.6 1.0 (4%) 0.0241 (10%)
Segmented 4.6 1.1 (5%) 0.0179 (14%) −23 4.3 1.0 (2%) 0.0214 (7%) −11
δ=12◦;c=2.5%cCorrugated 4.4 1.3 (4%) 0.0220 (8%) 5.0 1.2 (5%) 0.0185 (7%)
Segmented 4.7 1.4 (8%) 0.0155 (7%) −30 5.0 1.3 (2%) 0.0101 (17%) −45
Figure 17. Coefficients of pressure against non-dimensional chord of
suction surface of MSP demonstrator with morphing corrugated
trailing edge control surface fully extended.
is retracted (δ=0◦;c=0) and fully deflected and extended
(δ=12◦;c=2.5%c). While no major differences can be
observed for the lift data a significant reduction in the zero-lift
drag coefficients is measured at both air flow speeds. While
at this point it cannot be stated how close these zero-lift drag
coefficients are to a NACA 0024 with a smooth morphing
trailing edge control surface, it is certain that a discontinuous
segmented skin improves the aerodynamic performance of the
MSP concept.
Another way to reduce the drag increase might be to use
a different, thicker, aerofoil. It is to be noted though that
the NACA0024 is a fairly thick aerofoil compared to most
conventional aerofoils. Nevertheless for a thicker aerofoil the
pressure drag should dominate (Anderson 2001) and hence
the skin friction drag created by the corrugations (they can
be considered as an extremely rough surface) becomes less
important. Thus the MSP concept might be an interesting
alternative solution for trailing edge control surfaces of thick
aerofoils. However further work is necessary to confirm this
assumption.
Apart from the increase in drag generated from
the corrugated surface, some additional limitations were
experienced during the wind tunnel testing. When the MSPs
were not under tension significant out-of-plane deflections and
vibrations were observed at wind speeds above 20 m s−1.
This suggests that the concept may only be viable for small,
light and slow air vehicles. Should it be applied to larger
air vehicles, corresponding morphing ribs to support the
MSP without restricting their deformation might be necessary.
Furthermore, while an elastic chord length change of 4%
chord with no trailing edge control surface deflection was
achieved during the out of wind tunnel testing this may not
be enough for aircraft flight control. This chord length change
is variable subject to increasing the width of the MSP or the
transverse elastic structural strain capability of the MSP. The
latter can only be achieved at the expense of the out-of-plane
stiffness thus possibly requiring development of corresponding
morphing ribs.
A further point of discussion is the actuation power
required to morph the MSPs. Ideally the power required
to actuate a morphing skin and hold it in a desired shape
should be minimal. The holding power depends mainly on
the actuation system unless the skin shape can be ‘locked’,
i.e. shape memory materials (McKnight and Henry 2005, Reed
et al 2005). In the presented concept the lead screw allowed
for a zero holding power. The actuation power depends mainly
on the stiffness of the skin material. There has to be a
compromise between low skin stiffness (e.g. elastomeric based
concepts) for good morphing characteristics and low actuation
power and high stiffness (e.g. stiffness tailored composites)
for good load transfer. In theory shape memory materials do
not have this issue but their stiffness changes are typically
triggered by changes in temperature (Lendlein and Kelch 2002)
and thus require additional power; furthermore during the
low stiffness stage of the skin, the out-of-plane displacements
limit the operational airspeed, similar to the MSP concept.
Thus overall the MSP concept is believed to compare fairly
9

Smart Mater. Struct. 19 (2010) 124009 C Thill et al
similarly to other morphing skin concepts on an actuation
power requirements comparison.
5. Conclusions
This paper summarizes the main findings of extensive
structural and aerodynamic studies on corrugated composites
for morphing structures undertaken over the last few years.
Arranged in a corrugated sandwich structure their extremely
orthotropic properties can be effectively used in a morphing
trailing edge control surface application. A wide envelope
exists in which these structures can be designed to suit
the precise morphing skin requirements. For acceptable
aerodynamic performance the corrugation geometry and form
need to be designed adequately or an external discontinuous
segmented skin can be used to improve surface smoothness.
This study has combined the design, manufacture and
test of a wind tunnel demonstrator aerofoil with a morphing
trailing edge control surface using MSPs made from corrugated
sandwich structures. Chord length changes of up to 4% chord
and control surface deflections of up to 12◦, while maintaining
a continuous aerodynamic surface, were demonstrated in out of
wind tunnel testing. Wind tunnel testing showed the concept to
work well under aerodynamic loads at low speeds. The use of
a discontinuous segmented skin showed a significant reduction
in the drag generated by the corrugated trailing edge region
without altering the structural performance. The concept could
be further improved by the development of corresponding
morphing ribs to provide additional skin support.
Acknowledgments
The authors would like to thank the UK EPSRC (Engineering
and Physical Sciences Research Council) for their support and
funding of this work through the University Technology Strate-
gic Partnership—SMARTCOMP (EP/D03423X/1). Special
thanks go to the Faculty of Engineering workshop staff without
whose efforts this project would have been impossible: Lee
Winter, Russ Eyre and Alan Bishop; John Morrison, Michael
Powell, Chris Hunt and John Byles; Ian Chorley and Mike
Jones; Mark Fitzgerald; and to members of the Aerospace
Department academic staff: Drs Giuliano Allegri, Panagiotis
Margaris and Neil Taylor.
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