Composite panel

Abstract

A composite panel 100 for an aircraft structure comprising at least one metallic component, said composite panel comprising a composite laminate body, said composite laminate body that includes a plurality of dry-fibre plies 112 and a cured resin matrix 114 and a cover 120 applied to an edge (110a, fig 3a) of the composite laminate body for protecting the at least one metallic component against galvanic corrosion, said cover comprising a non-conductive material. The plurality of dry-fibre plies which make up the composite laminate body each comprise a plurality of conductive fibre reinforcements which terminate at said edge and wherein the cover is arranged so as to cap the respective ends of the plurality of conductive fibre reinforcements so as to substantially prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure during use. An aircraft structure, a method of manufacturing a composite panel for an aircraft structure and a method of sealing a trimmed edge of a composite aircraft panel are also claimed.

Description

FIELD OF THE INVENTION

[0001] The present invention relates to a composite panel for an aircraft structure comprising at least one metallic component, an aircraft structure comprising said composite panel, an aircraft comprising said aircraft structure and to a method of manufacturing said composite panel. BACKGROUND OF THE INVENTION

[0002] When conductive fibre-reinforced composite panels are used in applications in which the composite panels are in close proximity to metallic components, there is a risk that the metallic components of the aircraft structure can become exposed to the conductive fibres which can lead to galvanic corrosion.

[0003] To help prevent this issue, conductive fibre-reinforced composite panels intended for use in applications in which they are likely to be in close proximity to metallic components are often applied with a non-conductive seal around their edges.

[0004] Current methods used to seal the edges of conductive fibre-reinforced composite panels involve laying up a series of dry-fibre plies which are then impregnated and subsequently cured. Since reinforcing fibres are difficult to machine and cut, known composite panels often feature a rough edge after the initial cure that needs to be machined off. Known composite panels are therefore often made larger than is required for a given application and are machined to size after they have been impregnated and cured to achieve a machined edge. The machined edge is then sealed with a suitable resin.

[0005] However, such methods are generally undesirable. Firstly, designing a composite panel to be larger than is required adds unnecessary additional weight. Furthermore, the additional machining and sealing steps described above increase manufacturing times and cause material waste which increases production costs. There is also a risk that fibre fragments may be generated during edge machining which can float towards the machined edge and potentially compromise the resin seal.

[0006] It is therefore an aim of the present invention to provide a solution to at least one of the aforementioned setbacks associated with known sealing methods. SUMMARY OF THE INVENTION

[0007] According to a first aspect of the present invention, there is provided a composite panel for an aircraft structure comprising at least one metallic component, said composite panel comprising a composite laminate body, said composite laminate body comprising a plurality of dry-fibre plies and a cured resin matrix and a cover applied to an edge of the composite laminate body for protecting the at least one metallic component against galvanic corrosion, said cover comprising a non-conductive material, wherein the plurality of dry-fibre plies which make up the composite laminate body each comprise a plurality of conductive fibre reinforcements which terminate at said edge, and wherein the cover is arranged so as to cap the respective ends of the plurality of conductive fibre reinforcements so as to substantially prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure during use.

[0008] Advantageously, the provision a cover which is applied to an edge of the composite laminate body so as to cap the respective ends of the plurality of conductive fibre reinforcements helps to prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure thereby protecting said component against galvanic corrosion.

[0009] In exemplary embodiments, the conductive fibre reinforcements may be carbon fibre reinforcements.

[0010] In exemplary embodiments, the plurality of dry-fibre plies may be unidirectional dry-fibre plies.

[0011] In exemplary embodiments, the cover may comprise a base portion arranged so as to abut against the edge of the composite laminate body at which the plurality of conductive fibre reinforcements terminate and a pair of side portions which extend transversely away from the base portion, substantially parallel to a fibre direction of the plurality of dry-fibre plies.

[0012] In exemplary embodiments, the cover may be arranged such that the respective ends of the plurality of conductive fibre reinforcements are contained entirely between the pair of side portions.

[0013] Advantageously, containing the ends of the plurality of conductive fibre reinforcements entirely between the pair of side portions enables the cover to more effectively prevent an electrical contact from being formed between the respective ends of the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure during use.

[0014] In exemplary embodiments, the pair of side portions may ne tapered so as to define a ramped surface.

[0015] In exemplary embodiments, the composite laminate body may comprise at least one drop-off ply in a region of said ramped surface.

[0016] Advantageously, dropping-off one or more plies in the region of the ramped surface helps to reduce fibre compaction in the region of the cover which allows the resin matrix to better permeate through the composite laminate body during manufacture. Furthermore, dropping-off one or more plies in the region of the ramped surface also helps to avoid steps or “kinks” in the composite laminate body which is beneficial for stress distribution.

[0017] In exemplary embodiments, the cover may be a glass fibre-reinforced capping ply.

[0018] Advantageously, it has been found that glass fibre-reinforced capping plies are particularly effective at preventing an electrical contact from being formed between the respective ends of the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure due to their low conductivity. Furthermore, glass-fibre composites are relatively lightweight and hence the use of such a capping ply helps to minimise the weight of the composite panel.

[0019] In exemplary embodiments, the glass fibre-reinforced capping ply may have a single ply thickness.

[0020] Advantageously, providing a glass fibre-reinforced capping ply having a single ply thickness helps to further minimise the weight of the composite panel and also helps to avoid fibre compaction in the region of the cover.

[0021] In exemplary embodiments, the thickness of the glass fibre-reinforced capping ply may be less than a thickness of the dry-fibre plies which make up the composite laminate body.

[0022] In exemplary embodiments, the glass fibre-reinforced capping ply may have a thickness of approximately 1mm.

[0023] In exemplary embodiments, the thickness of each of the dry-fibre plies which make up the composite laminate body may be approximate 2.5mm.

[0024] In exemplary embodiments, the cover may be a polymer cap.

[0025] Advantageously, the provision of a polymer cap has been found to be effective at preventing an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component. Furthermore, the provision of a polymer cap can also afford protection to the composite laminate body against edge impacts.

[0026] In exemplary embodiments, the polymer cap may comprise a semi-flexible material, preferably Polyether ether ketone (PEEK).

[0027] Advantageously, the use of a semi-flexible polymer helps the cover to better withstand deformation stresses which may be applied to the cover during manufacture whilst still affording structural rigidity to the cover.

[0028] In exemplary embodiments, the composite laminate body may further comprise at least one non-conductive covering ply.

[0029] Advantageously, the provision of at least one non-conductive covering ply helps to prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure at the side portions of the composite laminate body, thereby further enhancing the galvanic corrosion protection afforded to the aircraft structure.

[0030] In exemplary embodiments, the at least one non-conductive covering ply may comprise a plurality of non-conductive fibre reinforcements.

[0031] In exemplary embodiments, the non-conductive fibre reinforcements may be orientated in the same direction as the conductive fibre reinforcements of the plurality of dry-fibre plies.

[0032] In exemplary embodiments, the at least one non-conductive covering ply may be a glass fibre-reinforced covering ply.

[0033] Advantageously, it has been found that glass fibre-reinforced covering plies are particularly effective at preventing an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure due to their low conductivity. Furthermore, glass-fibre composites are relatively lightweight and hence the use of such covering plies helps to minimise the weight of the composite panel.

[0034] In exemplary embodiments, the at least one non-conductive covering ply may comprise the same material as the cover.

[0035] In exemplary embodiments, the at least one non-conductive covering ply may underlap at least one of the side portions of the cover.

[0036] Advantageously, arranging the at least one covering ply so that it underlaps at least one of the side portions of the cover helps to more effectively prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure.

[0037] In exemplary embodiments, the at least one non-conductive covering ply and at least one of the side portions of the cover may together define a substantially step-free surface.

[0038] Advantageously, the provision of a substantially step-free surface helps to enhance the aerodynamic performance of the surface defined by the at least one non-conductive covering ply and the side portion of the cover.

[0039] In exemplary embodiments, the composite laminate body and the cover may be integrally formed.

[0040] In exemplary embodiments, the cured resin matrix may secure the cover to the composite laminate body.

[0041] According to a second aspect of the present invention, there is provided an aircraft structure comprising at least one metallic component and at least one composite component comprising the composite panel according to the first aspect of the present invention.

[0042] In exemplary embodiments, the at least one composite component may be one of an upper skin, a lower skin, a stringer, and / or a spar.

[0043] In exemplary embodiments, the aircraft structure may be an aircraft wing structure.

[0044] In exemplary embodiments, the at least one metallic component may be a rib or a pylon.

[0045] In exemplary embodiments, the at least one metallic component may comprise a light metal, preferably aluminium.

[0046] According to a third aspect of the present invention, there is provided an aircraft comprising the aircraft structure according to second aspect of the present invention.

[0047] According to a fourth aspect of the present invention, there is provided a method of manufacturing a composite panel for an aircraft structure comprising at least one metallic component, the method comprising the steps of: a) laying up a preform comprising a plurality of dry-fibre plies, wherein each of the dry-fibre plies comprise a plurality of conductive fibre reinforcements which terminate at an edge; b) applying a cover to the edge of the preform for protecting the at least one metallic component against galvanic corrosion, said cover comprising a non-conductive material and said cover being arranged so as to cap the respective ends of the plurality of conductive fibre reinforcements so as to substantially prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure during use; c) impregnating the preform with a resin matrix after the cover has been applied to the edge of the preform; and d) curing the resin matrix so as to secure the cover to the preform.

[0048] Advantageously, the aforementioned method allows a non-conductive cover to be applied to a composite panel during the initial lay-up and curing process. As such, the aforementioned method avoids having to perform separate edge machining and sealing processes after the composite panel has been cured thereby improving manufacturing times. Furthermore, since the aforementioned method does not require material to be machined off of the panel after cure, the panel does not need to be designed to be larger than is required for a given application and hence a weight saving benefit (as well as a reduction in material wastage) can also be achieved.

[0049] In exemplary embodiments, the cover may comprise a base portion and a pair of side portions which extend transversely away from the base portion, and step b) may comprise abutting the base portion of the cover against the edge of the composite laminate body such that that the respective ends of the plurality of conductive fibre reinforcements are contained entirely between the pair of side portions.

[0050] Advantageously, containing the respective ends of the plurality of conductive fibres entirely between the pair of side portions enables the cover to more effectively prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure during use.

[0051] In exemplary embodiments, the pair of side portions may be tapered such that each side portion defines a ramped surface, and, during step a), the plurality of dry-fibre plies may be laid up with at least one drop-off ply in a region of the ramped surface.

[0052] Advantageously, dropping-off one or more plies in the region of the ramped surface helps to reduce fibre compaction in the region of the cover which allows the resin matrix to better permeate through the composite laminate body during manufacture. Furthermore, dropping-off one or more plies in the region of the ramped surface also helps to avoid steps or “kinks” in the composite laminate body which is beneficial for stress distribution.

[0053] In exemplary embodiments, the cover applied to the edge of the preform during step b) may be a glass fibre-reinforced capping ply, and step b) may comprise wrapping the glass fibre-reinforced capping ply around the edge of the preform.

[0054] Advantageously, it has been found that glass fibre-reinforced capping plies are particularly effective at preventing an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure due to their low conductivity. Furthermore, glass-fibre composites are relatively lightweight and hence the use of such capping plies helps to minimise the weight of the composite panel.

[0055] In exemplary embodiments, step a) may comprise laying up at least one non-conductive covering ply on top of and / or beneath the plurality of dry-fibre plies.

[0056] Advantageously, the provision of at least one non-conductive covering ply helps prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component at the side portions of the composite laminate body, thereby further enhancing the galvanic corrosion protection afforded to the aircraft structure.

[0057] In exemplary embodiments, during step b), the cover may be applied to the edge of the preform such that the at least one non-conductive covering ply underlaps at least one of the side portions of the cover.

[0058] Advantageously, arranging the at least one non-conductive covering ply so that it underlaps at least one of the side portions of the cover helps to more effectively prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component.

[0059] In exemplary embodiments, during step b), the cover may be applied to the edge of the preform such that the at least one non-conductive covering ply and the at least one side portion of the cover together define a substantially step-free surface.

[0060] Advantageously, arranging the at least one non-conductive covering ply and the at least one side portion of the cover such that they define a substantially step-free surface helps to enhance the aerodynamic performance of the composite panel.

[0061] In exemplary embodiments, steps c) and d) may be performed via Resin Transfer Moulding (RTM) or Vacuum-Assisted Resin Transfer Moulding (VARTM).

[0062] According to a fifth aspect of the present invention, there is provided a method of sealing a trimmed edge of a composite aircraft panel, the method comprising the steps of: a) laying up a preform comprising a plurality of plies, wherein each of said plies comprises a plurality of conductive fibres; b) cutting the plurality of conductive fibres to obtain a trimmed preform edge; c) applying a galvanically-inert cover to the trimmed preform edge, said cover comprising a non-conductive material; d) impregnating the preform with a resin matrix after the cover has been applied to the trimmed preform edge; and e) curing the resin-impregnated preform to obtain a composite aircraft panel.

[0063] The term “non-conductive” referred to herein is defined as a material having an impedance value not exceeding 2* 10‘9Q*cm2.

[0064] The term “semiflexible polymer” referred to herein is defined as a material having a modulus of elasticity between 2GPa and 4GPa.

[0065] The term “conductive” referred to herein is defined as a material having an having an impedance value is excess of 2*10'9Q*cm2.

[0066] BRIEF DESCRIPTION OF THE DRAWINGS

[0067] Embodiments of the invention will now be described with reference to the accompanying drawings, in which:

[0068] Figure lisa plan view of an aircraft;

[0069] Figure 2 is a schematic plan view of a wing of the aircraft illustrated in Figure i;

[0070] Figure 3a is an exploded side view of a composite panel according to an embodiment of the present invention, said composite panel comprising a composite laminate body and a cover;

[0071] Figure 3b is a schematic side view of the composite panel illustrated in Figure 3a in which the plies which make up the composite laminate body are visible;

[0072] Figure 4a is a schematic side view of a composite panel according to an alternative embodiment of the present invention in which the plies which make up the composite laminate body are visible;

[0073] Figure 4b is a schematic side view of a composite panel according to yet another embodiment of the present invention in which the plies which make up the composite laminate body are visible; and

[0074] Figures 5a to 5c illustrate a method of manufacturing a composite panel according to an embodiment of the present invention. DETAILED DESCRIPTION OF EMBODIMENT(S)

[0075] Figure 1 shows an aircraft 10. The aircraft 10 has a fuselage 11, and starboard and port fixed wings 12. An engine 13 is mounted to each wing 12 via a respective pylon 13a. The aircraft 10 is a typical jet passenger transport aircraft but the invention is applicable to a wide variety of fixed wing aircraft types, including commercial, military, passenger, cargo, jet, propeller, general aviation, etc. with any number of engines attached to the wings or fuselage. The invention is also applicable to other aircraft, such as helicopters. The aircraft also comprises starboard and port horizontal stabilisers 14 and a vertical stabiliser 15.

[0076] A schematic view of a wing box 20 of the starboard wing 12 is shown in Figure 2. The port wing is similar in construction and so only a description of the starboard wing is provided herein.

[0077] The wing 12 has a cantilevered structure with a length extending in a span-wise direction from a wing root 16 to a wing tip 17, with the root 16 being joined to the aircraft fuselage 11. The wing 12 has a leading edge 18 and a trailing edge 19 as shown in Figure 1. The leading edge 18 is at the forward end of the wing 12 and the trailing edge 19 is at the rearward end of the wing 12. The wing 12 comprises the wing box 20 and leading and trailing edge assemblies.

[0078] The wing 12 has a spanwise axis which extends in a direction from the wing root 16 to the wing tip 17, and a chordwise axis which extends in the direction from the leading edge 18 to the trailing edge 19.

[0079] The wing box 20 forms a structural assembly and includes forward and rearward spars 21, 22 extending in a spanwise direction; ribs 23 extending between the forward and rear spars 21, 22 in a chordwise direction; upper and lower skins 24 on the upper and lower sides of the wing box 20; and stringers 30 which extend in a spanwise direction to which the skins 24 are mounted.

[0080] The wing box 20 acts as an aircraft structure. Although a description is provided herein with reference to the wing box 20, it will be understood that the invention is also applicable to other aircraft structures.

[0081] Aircraft structures used within modern aircraft typically comprise a plurality of different components which may be manufactured from different materials.

[0082] For example, in the wing box 20 depicted in Figure 2, the forward and rearward spars 21, 22, the upper and lower skins 24 and the stringers 30 are all made of a fibre-reinforced composite material such as a carbon fibre-reinforced composite material.

[0083] Meanwhile, the ribs 23 which extend between the forward 21 and rear 22 spars and the pylons 13a to which the engines 13 are mounted are made from a metallic material, typically a light metal such as aluminium. As such, it is common for aircraft structures to comprise at least one metallic component and at least one composite component.

[0084] However, since the fibre reinforcements of the composite component(s) have a different half-cell potential to the metallic component(s) of the aircraft structure, there is a risk that if the conductive fibre reinforcements which make up the composite component(s) are left exposed and come into direct electrical contact with the metallic component(s) (e.g., a rib) in the presence of a suitable conductive electrolyte (e.g., water), a galvanic reaction may occur which could lead to galvanic corrosion of the metallic component(s) of the aircraft structure.

[0085] Figures 3a and 3b show a composite panel 100 for an aircraft structure comprising at least one metallic component, such as the wing box 20 depicted in Figure 2.

[0086] In some embodiments, the composite panel 100 may constitute, or form part of, the upper skin 24, the lower skin (not shown), the forward spar 21, the rear spar 22 and / or one or more of the stringers 30. However, it shall be appreciated that in other embodiments, the composite panel 100 may constitute, or form part of, a different component of the aircraft structure.

[0087] The composite panel 100 is made up of a composite laminate body 110 and a cover 120 applied to an edge 110a of the composite laminate body 110 for protecting the one or more metallic component(s) of the aircraft structure against galvanic corrosion during use.

[0088] The composite laminate body 110 comprises a length dimension (L), a width dimension (not shown in Figure 3a) and a thickness dimension (Tl) which is significantly less than the length (L) and width dimensions. It shall be appreciated that the thickness dimension (Tl) of the composite laminate body 110 illustrated in Figures 3a and 3b has been enlarged so as to better illustrated the ply stacking (as shall be described in detail below).

[0089] The composite laminate body 110 is provided as a laminate structure formed from a stack of dry-fibre sheets (commonly known as plies 112) which are held together with a cured resin matrix 114 which is fed, for example by infusion and / or injection, into the ply stack during manufacture (as shall be described later within this application).

[0090] The plurality of dry-fibre plies 112 which make up the stack of sheets (also know as a ply stack) each comprise a plurality of conductive fibre reinforcements (not shown) which terminate at the edge 110a.

[0091] In the illustrated embodiment, the conductive fibre reinforcements (which make up each of the plurality of dry-fibre plies 112) are unidirectional carbon-fibre reinforcements. In other words, the plurality of conductive fibre reinforcements which make up each ply are all aligned in the same direction.

[0092] As shown in Figure 3b, in the illustrated embodiment the conductive fibre reinforcements of each ply 112 extend continuously in a lengthwise direction across substantially the full length (L) of the composite laminate body 110 before terminating at the edge 110a and hence provide structural rigidity to the composite panel 100.

[0093] However, it shall be appreciated that in other embodiments, other types of conductive fibre reinforcements (such as aramid fibres) may be used. Furthermore, it shall be appreciated that in other embodiments, the plurality of fibre reinforcements (not shown) may be provided in a different (e.g., width wise) orientation.

[0094] It shall also be appreciated that in some embodiments, the plurality of plies which make up the stack of sheets may be multi-axial plies in which the fibre reinforcements (which make up each ply 112) are aligned in more than one direction.

[0095] The composite laminate body 110 illustrated in Figure 3b also comprises a pair of non-conductive covering plies 116 which form the outermost (e.g., top and bottom) plies of the ply stack and extend across the full length and width of the composite laminate body 110.

[0096] In other words, the stack of dry-fibre plies 112 (which are reinforced with the conductive fibre reinforcements) are sandwiched between a pair of non-conductive covering plies 116. The non-conductive covering plies 116 help to prevent the conductive fibre reinforcements from forming an electrical contact with the one or more metallic component(s) of the aircraft structure, thereby helping to enhance the galvanic corrosion protection afforded to said metallic components.

[0097] In the illustrated embodiment, the non-conductive covering plies 116 are a pair of glass-fibre reinforced covering plies and hence each covering ply 116 comprises a plurality of non-conductive (glass) fibre reinforcements (not shown).

[0098] However, it shall be appreciated that in other embodiments, other types of non-conductive covering ply may be utilised which may have a different (or no) type of reinforcement.

[0099] As shown in Figure 3b, the glass fibre reinforcements of the non-conductive covering plies 116 are unidirectional glass fibre reinforcements which extend continuously in a lengthwise direction across substantially the full length (L) of the composite laminate body 110, terminating at the edge 110a.

[0100] As such, the non-conductive fibre reinforcements (not shown) of the non-conductive covering plies 116 are orientated in the same direction as the conductive fibre reinforcements (not shown) of the conductive fibre reinforced composite plies 112.

[0101] However, it shall be appreciated that in other embodiments, the plurality of non-conductive fibre reinforcements (not shown) may be provided in a different (e.g., width wise) orientation and / or the non-conductive covering plies may be multi-axial plies.

[0102] It shall also be appreciated that in some embodiments, one or more of the non-conductive covering plies 116 may be omitted.

[0103] Considering now the cover 120, the cover 120 is made from a galvanically inert non-conductive material and is applied to the edge 110a of the laminate body 110 at which the plurality of conductive fibre reinforcements terminate.

[0104] As shown in Figure 3b, the cover 120 is arranged so as to cap the respective ends of the plurality of conductive fibre reinforcements so as to substantially prevent an electrical contact being formed between one or more of the conductive fibre reinforcements and the one or more metallic component(s) present within the aircraft structure during use, thereby helping to enhance the galvanic corrosion protection afforded to said metallic component(s).

[0105] The cover 120 comprises a base portion 122 which is configured to abut against the edge 110a of the composite laminate body 110 at which the respective ends of the plurality of conductive fibre reinforcements terminate and a pair of side portions 124a,b which extend transversely away from the base portion 122 in a direction which is substantially parallel to the fibre direction of the dry-fibre plies 112 which make up the composite laminate body 110.

[0106] As shown in Figure 3b, the pair of side portions 124a,b are provided at either end of the base portion 122 and hence the cover 120 of the illustrated embodiment is substantially C-shaped. However, it shall be appreciated that in other embodiments, the cover 120 may comprise a different shape and / or configuration.

[0107] As with the composite laminate body 110, the base portion 122 of the cover 120 has a height dimension (H), a width dimension (not shown) and a thickness dimension (T2) which is significantly less than the height (H) and width dimensions.

[0108] The height dimension (H) of the base portion 122 is approximately equal to, or slightly greater than, the thickness (Tl) of the composite laminate body 110 and the width dimension (not shown) of the base portion 122 is approximately equal to, or slightly greater than, the width dimension of the laminate body 110.

[0109] As such, when the base portion 122 of the cover 120 is abutted against the edge 110a of the composite laminate body 110, the respective ends of the conductive fibre reinforcements (which make up the ply stack) are contained entirely between the pair of side portions 124a,b.

[0110] Advantageously, containing the ends of the conductive fibre reinforcements entirely between the pair of side portions 124a,b helps the cover 120 to more effectively prevent electrical contacts from being formed between the respective ends of the conductive fibre reinforcements and the one or more metallic component(s) of the aircraft structure, thereby further enhancing the galvanic corrosion protection afforded to said components during use.

[0111] Furthermore, as shown in Figure 3b, the cover 120 is applied to the edge 110a of the composite laminate body 110 such that the pair of covering plies 116, provided on the top and bottom surfaces of the ply stack, underlap the respective side portions 124a,b of the cover 120.

[0112] Advantageously, arranging the cover 120 such that the covering plies 116 underlap the side portions 124a,b of the cover helps to more effectively prevent electrical contacts from being formed between the plurality of conductive fibre reinforcements and the one or more metallic component(s) of the aircraft structure.

[0113] In addition, in the illustrated embodiment, the cover is arranged such that each covering ply 116 and corresponding side portion 124a,b of the cover 120 together define a substantially step-free surface thereby helping to enhance the aerodynamic performance of the composite panel 100.

[0114] In the embodiment illustrated in Figures 3a and 3b, the cover 120 is provided as a glass fibre-reinforced capping ply having a single ply thickness which is wrapped around the edge 110a of the composite laminate body 110 during manufacture.

[0115] As such, the cover 120 in the embodiment illustrated in Figures 3a and 3b comprises the same material as that of the non-conductive covering plies 116.

[0116] In the illustrated embodiment, the glass fibre-reinforced capping ply has a ply thickness of approximately 1mm whereas the carbon fibre-reinforced dry-fibre plies 112 which make up the composite laminate body 110 have a ply thickness of approximately 2.5mm. As such, in the illustrated embodiment, the thickness of the glass fibre-reinforced capping ply is less that the thickness of the dry-fibre plies 112 which make up the composite laminate body 110.

[0117] Advantageously, glass fibre-reinforced capping plies have been found to be particularly effective at preventing the formation of electrical contacts between the conductive fibre reinforcements and the one or more metallic component(s) of the aircraft structure due to their low conductivity.

[0118] Glass fibre-reinforced plies are also relatively lightweight and hence the use of such plies helps to minimise the weight of the composite panel 100. Furthermore, providing the cover as a glass-fibre reinforced capping ply having a single ply thickness helps to further minimise the weight of the composite panel and also helps to avoid compaction of the conductive fibre reinforcements proximal to the side portions 124a,b of the cover 120 which allows the resin matrix to better permeate through the composite laminate body 110 during manufacture (as shall be explained in greater detail below).

[0119] However, it shall be appreciated that in other embodiments, other types of cover 120 may be envisaged comprising a different structure, material and / or configuration.

[0120] An example of a composite panel 200 according to an alternative embodiment of the present invention shall now be described with reference to Figure 4a.

[0121] It shall be appreciated that the embodiment illustrated in Figure 4a has many features in common with the embodiment illustrated in Figures 3a and 3b and so, for the sake of conciseness, only the differences shall be described herein. Common features are denoted by corresponding reference numerals.

[0122] As can be seen in Figure 4a, the composite panel 200 is made up of a composite laminate body 210 and a cover 220 which is applied to an edge (not shown) of the composite laminate body 210 at which the plurality of conductive fibre reinforcements terminate.

[0123] However, unlike the embodiment illustrated in Figure 3 wherein the cover 120 is provided as a single glass-fibre reinforced capping ply, in the embodiment illustrated in Figure 4a, the cover 220 is a polymer cap which is pre-moulded from a semi-flexible polymer before being applied to the edge of the composite laminate body 210.

[0124] Advantageously, the aforementioned polymer cap has been found to be particularly effective at preventing the formation of electrical contacts between the conductive fibre reinforcements and the one or more metallic component(s) of the aircraft structure due to the low conductivity exhibited by such polymeric materials.

[0125] Furthermore, whilst the polymer cap is thicker (and hence slightly heavier) than the glass-fibre capping ply illustrated in Figure 3, unlike the glass fibre-capping ply the polymer cap has the added benefit of being able to afford impact protection to the edge of the composite laminate body 210.

[0126] The term “semi-flexible” is defined herein as a material having a modulus of elasticity between 2GPa and 4GPa.

[0127] An example of a suitable “semi-flexible” material from which the polymer cap may be formed is Polyether ether ketone (or PEEK). PEEK is particularly beneficial for use in the polymer cap of the present invention since it features the requisite modulus of elasticity (between 2GPa and 4GPa) and also is suitable for high heat applications which allows the polymer cap to withstand the curing process without deforming or losing its mechanical properties.

[0128] However, it shall be appreciated that the present invention is not so limited and so, in other embodiments, it shall be appreciated that the polymer cap may be formed using other “semi-flexible” materials.

[0129] As with the cover 120 illustrated in Figures 3a and 3b, the cover 220 (illustrated in Figure 4a) comprises a base portion 222 which is configured to abut against the edge of the composite laminate body 210 at which the respective ends of the plurality of conductive fibre reinforcements terminate and a pair of side portions 224a,b which extend transversely away from the base portion 222 in a direction which is substantially parallel to the fibre direction of the dry-fibre plies 212 which make up the composite laminate body 210.

[0130] However, whereas in Figures 3a and 3b the side portions 124a,b have a substantially uniform thickness, the cover 220 of Figure 4a comprises side portions 224a,b having a tapering thickness.

[0131] In other words, in the cover 220 illustrated in Figure 4a, a thickness (T3) at a root region of each side portion 224a,b (i.e., proximal to the base portion 222 of the cover 220) is greater than a thickness (T4) at a tip region of each side portion 224a,b (i.e., distal to the base portion 222 of the cover 220) and hence the thickness of each side portion 224a,b decreases as you move further away from the base portion 222 of the cover 220.

[0132] As shown in Figure 4a, the pair of side portions 224a,b are tapered so as to define a pair of ramped surfaces 226a,b on an inner side of the cover 220 which are configured to interface with the cover plies 216 of the composite laminate body 210.

[0133] Advantageously, the provision of ramped side portions 224a,b has been found to help reduce fibre compaction proximal to the side portions 224a,b of the cover 220 which allows the resin matrix to better permeate through the composite laminate body 210 during manufacture (as shall be explained in greater detail below).

[0134] Furthermore, the provision of ramped side portions 224a,b also helps to reduce the size of the “step” formed between the cover 220 and the composite laminate body 210 such that such that each covering ply 216 and corresponding side portion 224a,b together define a substantially step-free surface thereby helping to enhance the aerodynamic performance of the composite panel 200.

[0135] An example of a composite panel 300 according to another alternative embodiment of the present invention shall now be described with reference to Figure 4b.

[0136] It shall be appreciated that the embodiment illustrated in Figure 4b has many features in common with the embodiment illustrated in Figure 4a and so, for the sake of conciseness, only the differences shall be described herein. Common features are denoted by corresponding reference numerals.

[0137] Most notably, unlike the embodiment illustrated in Figure 4a wherein each ply 212 extends across the full length (L) of the composite laminate body 210, in the embodiment of Figure 4b a series of “drop off’ plies 318 (i.e., plies which do not extend across the full length (L) of the composite laminate body 310) are provided in the region of the ramped surfaces 326a,b.

[0138] Advantageously, the use of one or more “drop off’ plies 318 in the region of the ramped surfaces 326a,b avoids having to “squash” a large number of plies 312 between the side portions 324a,b of the cover 320 and hence helps to further reduce fibre compaction in the region of the cover which allows the resin matrix to better permeate through the composite laminate body 310 during manufacture.

[0139] Furthermore, dropping off one or more dry-fibre plies 318 in the region of the ramped surfaces 326a,b also helps to avoid steps or “kinks” which can be present when a large number of plies are “squashed” between the side portions 324a,b of the cover 320. By reducing the number of steps or “kinks”, the stress distribution across the composite laminate body 310 can be enhanced.

[0140] A method of manufacturing a composite panel 200 for an aircraft structure comprising at least one metallic component according to an embodiment of the present invention shall now be described with reference to Figures 5a-c.

[0141] It is important to note that the method described below is a “net edge” manufacturing method meaning that the edge of composite panel 100 obtained after the preform 130 and cover 120 have been co-cured is the edge that is present on the final product. In other words, the composite panel 100 does not feature a rough edge after manufacture that is later machined off.

[0142] “Net edge” manufacturing methods are beneficial since composite panels 100 manufactured via “net edge” methods do not require separate trimming and sealing operations post cure which helps to improve manufacturing times and also helps to reduce material wastage (thereby saving costs).

[0143] It shall also be appreciated that whilst the method illustrated in Figures 5a-c is a method for forming the composite panel 100 illustrated in Figures 3a and 3b, it shall be appreciated that corresponding methods may also be utilised for manufacturing composite panels according to other embodiments of the invention, such as the composite panels 200, 300 illustrated in Figures 4a and 4b.

[0144] In the illustrated embodiment, the composite panel 100 is manufactured by resin transfer moulding (or RTM) and uses a dry-fibre preform 130 which, when infused and cured, provides the composite laminate body 110 of the composite panel 100.

[0145] A preform (sometimes known as a charge) is a conventional term used to describe a component formed by a ply stack prior to said ply stack being supplied with a resin matrix.

[0146] Similarly, a “dry-fibre ply” is a sheet (or fabric) of reinforcing fibres (or tows) which have not yet been supplied with a resin matrix. As such, “dry-fibre plies” are distinguished over pre-impregnated (or “pre-preg”) plies which are a body of “preimpregnated” composite fibres which have been applied with a thermoset polymer resin matrix material prior to lay up.

[0147] In a first step of the method, a preform 130 comprising a plurality of dry fibre plies 112 is laid up on a first mould tool 140 as illustrated in Figure 5a.

[0148] In the embodiment illustrated in Figure 3, the composite panel 100 comprises a pair of top and bottom non-conductive covering plies 116. As such, during the lay-up step illustrated in Figure 5a, the bottom covering ply 116 is firstly placed on the first mould tool 140.

[0149] Once the non-conductive covering ply 116 has been placed on the first mould tool 140, the plurality of conductive fibre-reinforced dry-fibre plies 112 are stacked on top of the covering ply 116 so as to obtain a preform 130 having a desired shape and thickness.

[0150] The number of plies 112 laid up to form the preform 130 are typically selected so as to achieve a volume fraction of 60% in the cured composite laminate body 110 (meaning that 60% of the volume of the composite laminate body 110 is occupied by the reinforcing fibres with the remaining 40% being occupied by the cured resin matrix 114).

[0151] Advantageously, a volume fraction of 60% has been found to provide a good balance between fibre / resin adhesion and structural rigidity in the final cured composite laminate body 110.

[0152] Once the plurality of conductive fibre-reinforced dry-fibre plies 112 have been laid up, the second (or top) covering ply 116 is laid up on top of the stack such that the conductive fibre-reinforced dry-fibre plies 112 are sandwiched between the non-conductive covering plies 116.

[0153] In some embodiments, such as that which is illustrated in Figure 4b, the preform 130 may also be laid up with one or more drop off plies 318 in order to help further reduce fibre compaction within the preform 130, although it shall be appreciated that in other embodiments the provision of one or more “drop off’ plies may be omitted.

[0154] Furthermore, it shall also be appreciated that in some embodiments, one or more of the non-conductive covering plies 116 may also be omitted and hence in some embodiments the preform 130 may simply comprise a stack of conductive dry-fibre plies 112.

[0155] In order to obtain a preform 130 having the desired shape, the stack of dry-fibre plies 112, 116,118 which make up the preform 130 are also trimmed, for example using an ultrasonic knife, which results in the preform 130 featuring at least one trimmed edge 110a.

[0156] In some embodiments, the plurality of dry-fibre plies 112, 116, 118 may be trimmed individually before they are laid up in the ply stack. Alternatively, in other embodiments, the plurality of dry-fibre plies 112, 116 may first be laid up to obtain the preform 130 and then the pre-form 130 may subsequently trimmed to shape in a single operation after all of the dry-fibre plies 112, 116, 118 have been laid up.

[0157] Once the preform 130 has been laid up and trimmed, the cover 120 is applied to the trimmed edge 110a of the preform 130 at which the plurality of conductive fibre reinforcements terminate as shown in Figure 5b. It is important to note that this step (wherein the cover 120 is applied to the trimmed edge 110a of the preform 130) is performed prior to the application of the resin matrix.

[0158] As discussed previously, the cover 120 is applied onto the trimmed edge 110a of the preform 130 such that the base portion 122 of the cover 120 abuts against the respective ends of the plurality of conductive fibre reinforcements and such that the respective ends of the plurality of conductive fibre reinforcements are entirely contained between the pair of side portions 124a,b of the cover 120 with the pair of covering plies 116, provided on the top and bottom surfaces of the preform 130, underlapping the respective side portions 124a,b of the cover 120.

[0159] This arrangement effectively caps the respective ends of the plurality of conductive fibre reinforcements thereby substantially preventing any electrical contacts from being formed between the plurality of conductive fibre reinforcements and the one or more metallic component(s) of the aircraft structure during use.

[0160] In the illustrated embodiment, the cover 120 comprises a glass fibre-reinforced capping ply (such as that which is illustrated in Figures 3a and 3b) and hence the cover 120 is applied via wrapping the glass fibre-reinforced capping ply (or plies) around the trimmed edge 110a of the preform 130.

[0161] Alternatively, in embodiments wherein the cover 220, 320 is provided as a polymeric cap (see Figures 4a and 4b) the polymeric cap may be manufactured in advance, for example via injection moulding or another suitable method, before being pushed onto the trimmed edge 110a of the preform 130 after lay-up.

[0162] Once the cover 120 has been applied to the trimmed edge 110a of the preform 130, a second mould tool 150 is then fitted as shown in Figure 5c, and a curable resin matrix 114 is fed, for example by infusion and / or injection, into the cavity 160 between the mould tools 140, 150. In the illustrated embodiment, the curable resin matrix 114 is a thermosetting epoxy-resin. However, it shall be appreciated that in other embodiments, different types of resin may be used.

[0163] It shall also be appreciated that in some embodiments, the composite panel 100 may be manufactured by vacuum-assisted resin transfer moulding (or VARTM) and hence at the step illustrated in Figure 5c the preform 130 may be vacuum bagged prior to resin infusion / injection rather than being applied with a second mould tool 150.

[0164] The first 140 and second 150 mould tools hold the composite panel 100 between them. Thus, the composite panel 100 is compressed, consolidated, shaped and / or moulded by the first 140 and second 150 mould tools during the resin transfer moulding process. In embodiments wherein the composite panel 100 is manufactured via VARTM, the composite panel 100 is compressed, consolidated, shaped and / or moulded by a single mould tool 140.

[0165] As the curable resin matrix 114 is fed into the cavity 160, the preform 130 and the cover 120 become co-infused with the curable resin matrix 114.

[0166] The assembly is then heated so that the resin matrix 114 cures to co-cure the cover 120 and the preform 130 to each other.

[0167] As such, the composite laminate body 110 and the cover 120 of the composite panel 100 obtained via the aforementioned method are integrally formed with the cured resin matrix 114 securing the cover 120 to the composite laminate body 110.

[0168] Where the word 'or' appears this is to be construed to mean 'and / or' such that items referred to are not necessarily mutually exclusive and may be used in any appropriate combination.

[0169] Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims.

Claims

1. A composite panel for an aircraft structure comprising at least one metallic component, said composite panel comprising:a composite laminate body, said composite laminate body comprising a plurality of dry-fibre plies and a cured resin matrix; anda cover applied to an edge of the composite laminate body for protecting the at least one metallic component against galvanic corrosion, said cover comprising a non-conductive material,wherein the plurality of dry-fibre plies which make up the composite laminate body each comprise a plurality of conductive fibre reinforcements which terminate at said edge, andwherein the cover is arranged so as to cap the respective ends of the plurality of conductive fibre reinforcements so as to substantially prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure during use.

2. The composite panel according to claim 1, wherein the cover comprises:a base portion arranged so as to abut against the edge of the composite laminate body at which the plurality of conductive fibre reinforcements terminate, anda pair of side portions which extend transversely away from the base portion, substantially parallel to a fibre direction of the plurality of dry-fibre plies; andwherein the cover is arranged such that the respective ends of the plurality of conductive fibre reinforcements are contained entirely between the pair of side portions.

3. The composite panel according to claim 2, wherein the pair of side portions are tapered so as to define a ramped surface, and wherein the composite laminate body comprises at least one drop-off ply in a region of said ramped surface.

4. The composite panel according to any of claims 1 to 3, wherein the cover is a glass fibre-reinforced capping ply.

5. The composite panel according to claim 4, wherein the glass fibre-reinforced capping ply has a single ply thickness.

6. The composite panel according to any of claims 1 to 3, wherein the cover is a polymer cap.

7. The composite panel according to claim 6, wherein the polymer cap comprises a semi-flexible material, preferably Polyether ether ketone (PEEK).

8. The composite panel according to any preceding claim, wherein the composite laminate body further comprises at least one non-conductive covering ply.

9. The composite panel according to claim 8, wherein the at least one non-conductive covering ply is a glass fibre-reinforced covering ply, and preferably wherein the at least one non-conductive covering ply comprises the same material as the cover.

10. The composite panel according to claim 8 or claim 9, when dependent on claim 2 or claim 3, wherein the at least one non-conductive covering ply underlaps at least one of the side portions of the cover.

11. The composite panel according to any of claims 8 to 10, when dependent on claim 2 or claim 3, wherein the at least one non-conductive covering ply and at least one of the side portions of the cover together define a substantially step-free surface.

12. The composite panel according to any preceding claim, wherein the composite laminate body and the cover are integrally formed, and preferably wherein the cured resin matrix secures the cover to the composite laminate body.

13. An aircraft structure comprising:at least one metallic component; andat least one composite component comprising the composite panel according to any preceding claim.

14. The aircraft structure according to claim 13, wherein the at least one composite component is one of an upper skin, a lower skin, a stringer, and / or a spar.

15. The aircraft structure according to claim 13 or 14, wherein the aircraft structure is an aircraft wing structure.

16. The aircraft structure according to any of claims 13 to 15, wherein the at least one metallic component is a rib or a pylon.

17. The aircraft structure according to any of claims 13 to 16, wherein the at least one metallic component comprises a light metal, preferably aluminium.

18. An aircraft comprising the aircraft structure according to any of claims 13 to 17.

19. A method of manufacturing a composite panel for an aircraft structure comprising at least one metallic component, the method comprising the steps of:a) laying up a preform comprising a plurality of dry-fibre plies, wherein each of the dry-fibre plies comprise a plurality of conductive fibre reinforcements which terminate at an edge;b) applying a cover to the edge of the preform for protecting the at least one metallic component against galvanic corrosion, said cover comprising a non-conductive material and said cover being arranged so as to cap the respective ends of the plurality of conductive fibre reinforcements so as to substantially prevent an electrical contact from being formed between the plurality of conductive fibre reinforcements and the at least one metallic component of the aircraft structure during use;c) impregnating the preform with a resin matrix after the cover has been applied to the edge of the preform; andd) curing the resin matrix so as to secure the cover to the preform.

20. The method according to claim 19, wherein the cover comprises a base portion and a pair of side portions which extend transversely away from the base portion, and wherein step b) comprises abutting the base portion of the cover against the edge of the composite laminate body such that that the respective ends of the plurality of conductive fibre reinforcements are contained entirely between the pair of side portions.

21. The method according to claim 20, wherein the pair of side portions are tapered such that each side portion defines a ramped surface, and wherein, during step a), the plurality of dry-fibre plies are laid up with at least one drop-off ply in a region of the ramped surface.

22. The method according to any of claims 19 to 21, wherein the cover applied to the edge of the preform during step b) is a glass fibre-reinforced capping ply, and wherein step b) comprises wrapping the glass fibre-reinforced capping ply around the edge of the preform.

23. The method according to any of claims 19 to 22, wherein step a) comprises laying up at least one non-conductive covering ply on top of and / or beneath the plurality of dry-fibre plies.

24. The method according to claim 23, when dependent on claim 20 or 21, wherein, during step b), the cover is applied to the edge of the preform such that the at least one non-conductive covering ply underlaps at least one of the side portions of the cover and / or wherein, during step b), the cover is applied to the edge of the preform such that the at least one non-conductive covering ply and at least one side portion of the cover together define a substantially step-free surface.

25. A method of sealing a trimmed edge of a composite aircraft panel, the method comprising the steps of:a) laying up a preform comprising a plurality of plies, wherein each of said plies comprises a plurality of conductive fibres;b) cutting the plurality of conductive fibres to obtain a trimmed preform edge;c) applying a galvanically-inert cover to the trimmed preform edge, said cover comprising a non-conductive material;d) impregnating the preform with a resin matrix after the cover has been applied to the trimmed preform edge; ande) curing the resin-impregnated preform to obtain a composite aircraft panel.

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