WING STRUCTURE

Abstract

An aircraft wing (6) is provided. The aircraft wing (6) comprises at least one structure (62), (64), 66 comprising: a foam core (626), (666); first and second carbon fibre composite layers (624a), (622a), (662a), (664a) respectively attached to top and bottom sides of the foam core to sandwich the foam core; and third and fourth carbon fibre composite layers (624b), (622b), (662b), (664b) respectively disposed adjacent to the first and second carbon fibre composite layers, wherein the total thickness of the structure is between 1 mm and 11 mm. An aircraft having the aircraft wing and a method of manufacturing a structure are also provided.

Claims

1: An aircraft wing, the aircraft wing comprising: at least one structure comprising: a foam core; first and second carbon fibre composite layers respectively attached to top and bottom sides of the foam core to sandwich the foam core; and third and fourth carbon fibre composite layers respectively disposed adjacent to the first and second carbon fibre composite layers, wherein the total thickness of the structure is between 1 mm and 11 mm.

2: The aircraft wing according to claim 1, wherein the at least one structure is an upper skin, a lower skin or a spar, or any combination thereof.

3: The aircraft wing according to claim 2, wherein the at least one structure is an upper skin or lower skin, and wherein: fibres in the first carbon fibre composite layer are arranged substantially orthogonally to fibres in the third carbon fibre composite layer and parallel to the plane of the first carbon fibre composite layer; fibres in the second carbon fibre composite layer are arranged substantially orthogonally to fibres in the fourth carbon fibre composite layer and parallel to the plane of the second carbon fibre composite layer; the fibres in the first carbon fibre composite layer or third carbon fibre composite layer are arranged at about 45±15 degrees to a spanwise axis of the aircraft wing; and the fibres in the second carbon fibre composite layer or fourth carbon fibre composite layer are arranged at about 45±15 degrees to the spanwise axis of the aircraft wing.

4: The aircraft wing according to claim 2, wherein the at least one structure is a spar, and wherein: fibres in the first carbon fibre composite layer are arranged substantially orthogonally to fibres in the third carbon fibre composite layer and parallel to the plane of the first carbon fibre composite layer; fibres in the second carbon fibre composite layer are arranged substantially orthogonally to fibres in the fourth carbon fibre composite layer and parallel to the plane of the second carbon fibre composite layer; the fibres in the first carbon fibre composite layer or third carbon fibre composite layer are arranged at about 45±15 degrees to the longitudinal axis of the spar; and the fibres in the second carbon fibre composite layer or fourth carbon fibre composite layer are arranged at about 45±15 degrees to the longitudinal axis of the spar.

5: The aircraft wing according to claim 2, wherein: the at least one structure includes an upper skin and a lower skin; the upper skin and lower skin are joined to form an aerofoil; and the wing comprises a spar disposed between the upper skin and lower skin in the longitudinal direction of the wing structure.

6: The aircraft wing according to claim 2, wherein the at least one structure do not have a further carbon fibre composite layer to the first, second, third and fourth carbon fibre composite layers.

7: The aircraft wing according to claim 1, wherein the foam core has a thickness of between 1 mm and 10 mm.

8: The aircraft wing according to claim 1, wherein the first, second, third and fourth carbon fibre composite layers are each between 10 μm and 50 μm thick.

9: The aircraft wing according to claim 2, wherein the at least one structure includes a spar that comprises an elongate panel, wherein top and bottom flanges of the panel curve away from the panel to couple the panel to the upper skin and lower skin.

10: The aircraft wing according to claim 2, wherein the at least one structure includes a spar that further comprises: respectively disposed adjacent to the third and fourth carbon fibre composite layers, fifth and sixth carbon fibre composite layers; respectively disposed adjacent to the fifth and sixth carbon fibre composite layers, seventh and eighth carbon fibre composite layers; respectively disposed adjacent to the seventh and eighth carbon fibre composite layers, ninth and tenth carbon fibre composite layers; respectively disposed adjacent to the ninth and tenth carbon fibre composite layers, eleventh and twelfth carbon fibre composite layers; respectively disposed adjacent to the eleventh and twelfth carbon fibre composite layers, thirteenth and fourteenth carbon fibre composite layers; and respectively disposed adjacent to the thirteenth and fourteenth carbon fibre composite layers, fifteenth and sixteenth carbon fibre composite layers; wherein the first, second, third, fourth, fifth, sixth, seventh, eighth, ninth, tenth, eleventh, twelfth, thirteen, fourteenth, fifteenth and sixteenth carbon fibre composite layers are each between 10 μm and 50 μm thick.

11: The aircraft wing according to claim 2, wherein the at least one structure includes a spar that is disposed at between 24% and 36% Mean Aerodynamic Chord.

12: The aircraft wing according to claim 1, wherein each carbon fibre composite layer comprises 30 to 40 mass percent resin and 60 to 70 mass percent carbon fibre.

13: The aircraft wing according to claim 1, wherein the foam core is a polymer foam core.

14: The aircraft wing according to claim 1, wherein each carbon fibre composite layer has a glass transition temperature greater than 80 degrees Celsius.

15: An aircraft comprising the aircraft wing according to claim 1, wherein the aircraft wing has an aspect ratio greater than 17:1.

16: A method of manufacturing a structure, the method comprising: providing a mould tool having a mould surface; providing an uncured structure, the uncured structure comprising: a foam core; first and second carbon fibre composite layers respectively attached to top and bottom sides of the foam core to sandwich the foam core; and third and fourth carbon fibre composite layers respectively disposed adjacent to the first and second carbon fibre composite layers; applying the uncured structure to the mould surface of the mould tool so as to form an assembly; and curing the assembly so as to cure the uncured structure and mould the uncured structure against the mould surface, thereby producing the structure having a surface that is substantially the same shape as and contiguous with the mould surface, to provide an aircraft wing according to claim 1.

17: The method according to claim 16, wherein the structure is an upper skin or lower skin of the aircraft wing, the method further comprising: arranging fibres in the first carbon fibre composite layer to be orientated substantially orthogonally to fibres in the third carbon fibre composite layer and parallel to the plane of the first carbon fibre composite layer; arranging fibres in the second carbon fibre composite layer to be orientated substantially orthogonally to fibres in the fourth carbon fibre composite layer and parallel to the plane of the second carbon fibre composite layer; arranging the fibres in the first carbon fibre composite layer or third carbon fibre composite layer to be orientated at about 45±15 degrees to a spanwise axis of the aircraft wing; and arranging the fibres in the second carbon fibre composite layer or fourth carbon fibre composite layer to be orientated at about 45±15 degrees to a spanwise axis of the aircraft wing.

18: The method according to claim 16, wherein the structure is a spar, the method comprising: arranging fibres in the first carbon fibre composite layer to be orientated substantially orthogonally to fibres in the third carbon fibre composite layer and parallel to the plane of the first carbon fibre composite layer; arranging fibres in the second carbon fibre composite layer to be orientated substantially orthogonally to fibres in the fourth carbon fibre composite layer and parallel to the plane of the second carbon fibre composite layer; arranging the fibres in the first carbon fibre composite layer or third carbon fibre composite layer to be orientated at about 45±15 degrees to a longitudinal axis of the spar; and arranging the fibres in the second carbon fibre composite layer or fourth carbon fibre composite layer to be orientated at about 45±15 degrees to the longitudinal axis of the spar.

19: The aircraft wing according to claim 2, wherein the first, second, third and fourth carbon fibre composite layers are each between 10 μm and 50 μm thick.

20: The aircraft wing according to claim 3, wherein the first, second, third and fourth carbon fibre composite layers are each between 10 μm and 50 μm thick.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0062] Embodiments of the invention will now be described by way of example only and with reference to the accompanying drawings.

[0063] FIG. 1 is a perspective view of an aircraft having the skin structure according to the present invention;

[0064] FIG. 2 is a cross section of a wing structure having the skin structure according to the present invention;

[0065] FIG. 3 is a side view of a wing skin structure according to the present invention;

[0066] FIG. 4 is a side view of a spar structure according to the present invention; and

[0067] FIG. 5 shows test results for a wing structure having the skin structure according to the present invention.

DETAILED DESCRIPTION

[0068] Embodiments described herein generally relate to a skin and spar structures for use on aircraft. Primarily, the skin and spar structures are designed for use on monocoque aircraft, where the flight loads are distributed through and supported by the skin rather than internal structure of the airframe. The skin structure is designed to have a high strength to mass ratio.

[0069] FIG. 1 shows an aircraft 100 on which the skin structure is implemented. The aircraft 100 is a high-altitude long endurance (HALE) unmanned aircraft. However, the skin structure may also be implemented on other types of manned and unmanned aircraft, such as balloons and helicopters. HALE aircraft are those typically capable of flying as high as 18,288 metres (60,000 feet) with an endurance of 32 hours or more. They typically loiter at a low velocity. Medium-altitude long endurance (MALE) aircraft, are those aircraft typically designed to operate between 3,048 metres (10,000) feet and 9,144 metres (30,000 feet) for periods up to 48 hours.

[0070] The aircraft 100 includes a payload 2 coupled to the front central part of a wing structure 6. The wing structure 6 includes a wing on either side of a central part. A fuselage 4 is coupled to the rear of the central part of the wing structure 6. An empennage 8 having tail surfaces for controlling the pitch and yaw of the aircraft 100 is coupled to the rear of the fuselage 4.

[0071] Due to the aircraft 100 being required to operate efficiently at high altitudes, the aircraft 100 is fitted with a wing structure 6 having a high aspect ratio. High altitudes are for example altitudes between about 16,000 metres and about 25,000 metres. Preferably, high altitudes are those between about 17,000 metres and about 21,000 metres. Wings with high aspect ratios provide more lift than low or moderate aspect ratio wings, and enable sustained endurance flight due to reduced drag. The aspect ratio is the ratio of the wing span to mean chord, equal to the square of the wingspan divided by the wing area. The wing aspect ratio of the aircraft 100 is preferably between about 17:1 and about 52:1. More preferably, the wing aspect ratio is between about 30:1 and about 40:1. For example, from the distal tip of each wing, the wing structure 6 is about 36 metres long. Preferably the wing span of the wing structure 6 is between about 30 and about 36 metres (for example about 35 metres). The mean chord of the wing structure 6 is about 1.2 metres. This results in an elongate wing structure.

[0072] Engines, batteries, and flight control systems are housed in nacelles in each wing of the wing structure 6, either side of the centre of the wing. Alternatively, the engines, batteries and flight control systems are housed in pods coupled to each wing of the wing structure 6.

[0073] FIG. 2 shows a cross section through the wing structure 6. The wing structure 6 includes a leading-edge strip 68. An upper skin 62 and lower skin 64 are joined or bonded together at the leading edge of the wing structure 6 by the leading-edge strip 68 to form an aerofoil shape. The upper skin 62 and lower skin 64 each comprise a skin structure as described with reference to FIG. 3.

[0074] In an alternative embodiment, the wing structure 6 comprises only one structure, such as the upper skin 62 or lower skin 64. Here, the single structure is moulded around a foam core to form an aerofoil shape.

[0075] The height of the leading-edge strip 68 is for example about 10-15 mm. The leading-edge strip 8 fits into a recess in the upper and lower portions. The leading-edge strip 8 is conformal with the wing structure external profile. The leading-edge strip 8 is continuous along the length of the wing structure 6.

[0076] An exemplary ply schedule for the leading-edge strip 68 is as follows:

TABLE-US-00001 Layer Fibre orientation Material 1 +45 degrees CRFC (carbon reinforced fibre composite) Mitsubishi Pyrofil HS40 fibre (carbon fibre)— about 20 grams per metre squared (g/m.sup.2) Thinpreg 402 epoxy—about 35% resin mass fraction 2 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 3 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 gsm Thinpreg 402 epoxy—about 35% resin mass fraction 4 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 5 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 6 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 7 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 8 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction Note: the zero-degree orientation reference is along the long axis of the part.
The leading-edge strip 68 also has a layer of Kapton on its whole outer surface (i.e. over the whole joint/junction).

[0077] A spar shear web 66 couples the upper skin 62 and lower skin 64 at approximately the midpoint of the aerofoil cross section. More specifically, the spar shear web 66 is located at about 30% Mean Aerodynamic Chord, which in the specific example is about 360 mm from the leading-edge strip 68 across the length of the wing structure 6. The distance between the leading and trailing edge of the wing, measured parallel to the normal airflow over the wing, is known as the chord. If the leading edge and trailing edge are parallel, the chord of the wing is constant along the wing's length. The width of the wing is greatest where it meets the fuselage at the wing root and progressively decreases toward the tip. As a consequence, the chord also changes along the span of the wing. The average length of the chord is known as the Mean Aerodynamic Chord. The spar shear web 66 extends through the longitudinal axis of the wing structure 6, i.e. from tip to tip. The spar shear web 66 takes the form of an I-beam. The spar shear web 66 carries the bending loads of the wing structure 6.

[0078] The upper and lower spar caps, i.e. the top and bottom parts of the !-shaped beam, are about 3 mm deep when consolidated. The spar caps are constructed using carbon fibre composite. In an exemplary embodiment, the spar caps are made from high-strength unidirectional carbon fibre composite, such as Mitsubishi Pyrofil MR70 (with the carbon fibres of the plies running in the direction of the long-axis (i.e. spanwise on the aircraft wing structure 6)). The spar caps are encapsulated within carbon fibre composite skins 662, 664. The width of the spar caps is varied between about 10 mm and about 20 mm locally along the wing structure 6 to optimise the mass of the spar shear web 66 as a function of the local wing loads.

[0079] The spar shear web 66, shown in FIG. 4, includes an about 5 mm thick foam/carbon fibre sandwich panel 666 disposed between the upper and lower spar caps, perpendicular to their longitudinal axis. The foam core of the sandwich is cut from Rohacell 31 IG foam. An exemplary ply schedule for the spar shear web 66 panel is as follows:

TABLE-US-00002 Fibre Layer orientation Material 1 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 2 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 3    0 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 4 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 5 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 6 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 7 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 8    0 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 9 n/a Film Adhesive GF736—about 25 g/m.sup.2 epoxy resin 10 n/a Rohacell 31IG sheet—about 5 mm thick foam 11 n/a Film Adhesive GF736—about 25 g/m.sup.2 epoxy resin 12    0 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 13 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 14 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 15 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 16 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 17    0 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 18 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction 19 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin mass fraction

[0080] In the table above, the zero-degree orientation is parallel to the longitudinal axis of the spar shear web 66. The fibres are arranged parallel to the plane of the ply.

[0081] Where the spar shear web 66 bonds to the upper skin 62 and lower skin 64, an about 5 mm radius curve is added to prevent sharp folding and consequently damage of the fibre in the composite skins.

[0082] The total thickness of the spar shear web 66 in the exemplary embodiment is about 5.4 mm. In other words, each of the sixteen layers of carbon fibre composite 662a-h, 664a-h is about 25 μm thick, and the foam core 666 is about 5 mm thick. In further embodiments, the foam core 666 has a thickness between about 1 mm and about 10 mm, and each layer of carbon fibre composite 662a-h, 664a-h has a thickness between about 10 μm and about 50 μm. In further embodiments, each layer of carbon fibre composite 662a-h, 664a-h has a thickness between about 15 μm and about 40 μm. In further embodiments, each layer of carbon fibre composite 662a-h, 664a-h has a thickness between about 20 μm and about 30 μm.

[0083] Although not shown in the Figures, ribs are disposed within the wing structure 6, in the lateral direction of the wing structure 6. The ribs are a carbon fibre composite and polymethacrylimide sandwich construction. The ribs are located at a regular 500 mm spacing along the wing structure 6 to provide accurate definition of the wing structure 6 and to increase pitching stiffness of the wing structure 6. Additional ribs are added at certain locations such as above the motor pods and fuselage joiner, where the wing structure 6 couples to the fuselage 4.

[0084] The ribs are Computer Numerical Controlled (CNC) machined from flat pre-cured sandwich panel material (pre-preg) that uses Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 and Thinpreg 402 epoxy—about 35% resin fraction skins and an about 50 g/m.sup.2 Redux 312 epoxy film adhesive to bond to the about 3 mm thick Rohacell 31 IG foam core.

[0085] Rohacell 31 IG is a closed-cell rigid foam based on polymethacrylimide (PMI) chemistry. It has a density of about 32 kg/m.sup.3, a compressive strength of about 0.4 MPa, compressive modulus of about 17 MPa, tensile strength of about 1.0 MPa, tensile modulus of about 36 MPa, shear strength of about 0.4 MPa and shear modulus of about 13 MPa. It would be appreciated that other foams having similar characteristics may be used, such as polyurethane foam or polyethylene foam.

[0086] The construction of the upper skin 62 of the wing structure 6 will now be described with reference to FIG. 3. The lower skin 64 is constructed using the same technique as will now be described, albeit using a different mould to achieve a different shape.

[0087] The upper skin 62 of the wing structure 6 is a composite sandwich panel constructed from a laminate of carbon fibre plies 622a (i.e. first carbon fibre composite layer), 622b (i.e. third carbon fibre composite layer), 624a (i.e. second carbon fibre composite layer), 624b (i.e. fourth carbon fibre composite layer), and a core material 626 disposed/sandwiched between two groups 622, 624 of plies. In an exemplary embodiment, the core material 626 is Rohacell 31 IG foam. The core material 626 is about 3 mm thick. The lower carbon fibre composite layer 624 comprises two carbon fibre composite plies 624a, 624b. Each carbon fibre composite ply 624a, 624b is about 25 μm thick. The upper carbon fibre composite layer 622 comprises two carbon fibre composite plies 622a, 622b. Each carbon fibre composite ply 622a, 622b is about 25 μm thick. The total thickness of the upper skin 62 in the exemplary embodiment is about 3.1 mm. In other words, each of the four layers (plies) of carbon fibre composite 622a-b, 624a-b is about 25 μm thick, and the foam core 626 is about 3 mm thick. In further embodiments, the foam core 626 has a thickness between about 1 mm and about 10 mm, and each layer of carbon fibre composite 622a-b, 624a-b has a thickness between about 10 μm and about 50 μm. In further embodiments, each layer of carbon fibre composite 622a-b, 624a-b has a thickness between about 15 μm and about 40 μm. In further embodiments, each layer of carbon fibre composite 622a-b, 624a-b has a thickness between about 20 μm and about 30 μm.

[0088] The upper skin 62 and lower skin 64 are manufactured within a female mould so that the outer surface of the wing structure 6 is the smooth moulded surface. The moulding occurs using a ‘single step’ approach, in an oven with 1 atmosphere pressure. All composite parts are post-cured such that the glass transition temperature (Tg) is greater than about 80° C.

[0089] An exemplary ply schedule for the upper skin 62 is as follows:

TABLE-US-00003 Fibre Layer orientation Material 1 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin fraction 2 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin fraction 3 n/a GF736 epoxy film adhesive about 25 g/m2 epoxy resin 4 n/a Rohacell 31IG, about 3 mm foam 5 n/a GF736 epoxy film adhesive about 25 g/m.sup.2 epoxy resin 6 −45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin fraction 7 +45 degrees CRFC Mitsubishi Pyrofil HS40 fibre—about 20 g/m.sup.2 Thinpreg 402 epoxy—about 35% resin fraction Note: fibre orientation 0-degree reference is spanwise across the wing.

[0090] Orienting the fibres in the carbon fibre composite plies 622a, 622b, 624a, 624b such that they are orthogonal to each other (but not parallel or orthogonal to a spanwise axis (i.e. an axis running in the spanwise direction) of the wing structure 6) tends to improve the torsional stiffness of the upper skin 62, and hence wing as a whole. While orthogonal orientation is preferable, similar advantage tends to be achieved through having fibres arranged at other angles, such as about 70 degrees (for example, fibres in one layer 622a arranged at about −30 degrees and fibres in the other layer 622b arranged at about +40 degrees). Other examples include the fibres in adjacent carbon fibre plies 622a, 622b being arranged at about 50 degrees to each other (for example, fibres in one layer 622a arranged at about −10 degrees and fibres in the other layer 622b arranged at about +40 degrees; about 40 degrees to each other (for example, fibres in one layer 622a arranged at about −30 degrees and fibres in the other layer 622b arranged at about +10 degrees); about 70 degrees to each other (for example, fibres in one layer 622a arranged at about −35 degrees and fibres in the other layer 622b arranged at about +35 degrees); and about 80 degrees to each other (for example, fibres in one layer 622a arranged at about −30 degrees and fibres in the other layer 622b arranged at about +40 degrees. The fibres are arranged parallel to the plane of the respective ply.

[0091] The skin structure of the upper skin 62 is substantially the same as the skin structure of the fuselage 4, engine pods and empennage 8.

[0092] FIG. 4 demonstrates the effectiveness of the described skin structure of the wing structure 6. At the furthest point from the fuselage, the wing structure deflects by about 0.43 metres (1.14% of wing span) when the aircraft 100 is performing a +2.5 g manoeuvre at an equivalent airspeed (EAS) of 11.1 ms.sup.−1. This is a relatively small deflection, given the whole mass of the structure forming the aircraft 100 is about 42.6 kg.

[0093] The present disclosure tends to provide a wing structure that is strong yet light enough to be suitable for high-altitude long-endurance flight.

[0094] While a fixed wing aircraft 100 has been described, it would be readily appreciated that the skin structure could be applied to a different type of vehicle. For example, instead of a wing structure, the described skin structure could be applied in a similar manner to the rotor blade of a helicopter.

[0095] Where, in the foregoing description, integers or elements are mentioned that have known, obvious, or foreseeable equivalents, then such equivalents are herein incorporated as if individually set forth. Reference should be made to the claims for determining the true scope of the present disclosure, which should be construed so as to encompass any such equivalents. It will also be appreciated by the reader that integers or features of the disclosure that are described as optional do not limit the scope of the independent claims. Moreover, it is to be understood that such optional integers or features, while of possible benefit in some embodiments of the disclosure, may not be desirable, and can therefore be absent, in other embodiments.