Methods of manufacturing composite structures

Abstract

A method of manufacturing a composite wing structure 200’ comprising a stiffener 220’ having an open side 224’ involves supporting a resin-infused fibre preform for forming the stiffener upon a suppor

Description

BACKGROUND OF THE INVENTION

[0001] The present invention concerns methods of manufacturing composite structures. More particularly, this invention concerns a method of manufacturing a composite structure comprising a stiffener for an aircraft wing.

[0002] An aircraft 100 having a wing 101 comprising a composite structure configured to provide stiffness to the wing 101 is shown in FIG. 1. The wing 101 has a wing tip 103 which, as can be seen in FIG. 2, curves upwardly in vertical Y direction when moving in a spanwise X direction, towards the tip 105 of the wingtip 103. The wingtip 103 is also swept rearwardly, as can be seen in FIG. 1, and has a twist angle T, which is illustrated schematically in FIG. 4 and is described in more detail below. Composite structures may form part of the airframe used to provide structural stiffness to prior art wings of this type.

[0003] A cross-sectional view of a composite structure 200 used to provide stiffness to the wing 105 at the wing tip 103 is shown schematically in FIG. 3. The composite structure 200 is part of a wing box, and comprises a wing skin 210 and two stiffeners 220 in the form of C-spars that are spaced-apart on the skin 210 in a chord-wise Z direction of the wing. The wing skm 210 may be an upper wmg skin or a lower wmg skin and, as will be understood by the skilled person, the composite structure 200 is configured such that a respective lower or upper wing skin, which is not shown in the figures, can be bonded or fastened to the upper-most surfaces 226 of the C-spars 220 to close the wing box.

[0004] The C-spars 220 extend in a spanwise X direction of the wing 101, perpendicular to the Y-Z plane indicated in FIG 3. Each C-spar 220 is formed by three structural elements 221, 222, 223. A first element 221 provides a foot that is bonded to the skin 210, a second element 222 extends away from the first element 221 to provide a web of the stiffener 220, and a third element 223 extends away from an end of the second element 222 to provide a flange of the stiffener 220 to which the upper wing skin will be attached. The third element 223 is spaced apart from the first element 221 and is shown in FIG. 3 as being oriented parallel to the first element. However, it will be appreciated that in typical wing boxes the orientation of the first element 221 and third element 223 may be dictated by the curvature of the aerodynamic surface of the wing, so may not be oriented parallel with one another. As such, the angle A between the first element 221 and the second element 222 and the angle B between the second element 222 and the third element 223, which are both labelled in FIG. 3, are not normally equal to 90 degrees and may vary along the spanwise direction of the wing. Together the three elements of the stiffener 221, 222, 223 are configured to provide the stiffener in the form of a C-spar 220 with a cross-section having a general C-shape. An open side 224 of the C-spar is defined between the first element 221 and the third element 223.

[0005] Prior art composite wing structures may of course have stiffeners with crosssections having other shapes, such as an I-shape, J-shape, T-shape or Z-shape. Some examples of stiffeners having those shapes are shown schematically in FIG. 5 where the depicted composite wing box is shown with a J-spar 260, a Z-stnnger 262, T-stringer 264, an I-stringer 266, and a J-stringer 268.

[0006] In prior art composite structures of the type described above, the stiffeners forming the spars or stringers may curve in the Y or Z directions, twist, and alternatively, or additionally taper in the spanwise X direction, depending on the shape of the wing that they form part of. As such, the stiffeners may have a relatively complex shape. This is illustrated by comparing the cross-section of the composite structure 200 shown in FIG. 3, which is taken at location LI indicated in FIG. 2, with the cross-section shown in FIG. 4, which is taken at location L2 in FIG. 2, closer the tip 105 of the wing tip 103. In the cross-section of FIG. 4, the C-spars 220 are closer together, have smaller cross-sectional dimensions, and have been twisted by the twist angle T of the wing, relative to the cross-section shown in FIG. 3.

[0007] When manufacturing a composite wing structure of the type described above, it is necessary to support a fibre preform for forming the stiffeners upon some form of tooling. The fibre preform may comprise a pre-preg or it may comprise dry fibres which are later infused with resin. The resin is then cured to produce the composite structure. A problem associated with manufacturing such composite structures is that it is not always possible to use one-piece tooling to support the fibre preform for forming a stiffener. The complex shape of the stiffener requires complementarity shaped tooling which in some cases would become trapped by the composite structure post-cure, such that it cannot be removed from the open side 224 of the stiffener 220. As such, prior art methods of manufacturing composite wing structures having such stiffeners rely on the use of complex multi-part tooling which is expensive to produce and time-consuming to assemble and disassemble.

[0008] The present invention seeks to mitigate the above-mentioned problems. Alternatively or additionally, the present invention seeks to provide an improved method of manufacturing a composite structure having a stiffener. SUMMARY OF THE INVENTION

[0009] The present invention provides a method of manufacturing a composite wing structure comprising a stiffener having an open side. The method comprises steps of supporting a resin-infused fibre preform for forming the stiffener upon a support structure and curing the resin to create the composite wing structure. The composite wing structure is configured such that the stiffener partially surrounds and encloses the support structure, thereby defining a volume in which the support structure is retained by the stiffener such that the support structure cannot be removed from the open side of the stiffener in one piece. The support structure comprises a monolithic sacrificial support element and the method comprises breaking the monolithic sacrificial support structure post-cure to permit removal of the monolithic sacrificial support structure from the volume defined by the stiffener.

[0010] It should be noted that the term “stiffener” as used herein should be understood to refer to a spar or to a stringer of a composite wing structure. The method of the invention may be used to manufacture any type of stiffener having an open side. The term “open side” as used herein refers to a stiffener having a cross-section that is not completely closed. For example, a stiffener having a square cross-section is completely closed and does not have an open side from which the sacrificial support structure can be accessed post-cure in order to break and remove the sacrificial support element. The skilled person will be aware of many types of stiffeners having open sides. For example, stiffeners having cross-sections having a general C-shape, T-shape, I-shape, J-shape, and Z-shape, such as those shown schematically in FIG. 5 and FIG. 11, all have open sides and are therefore suitable for being manufactured using a method according to the invention.

[0011] The method may comprise breaking the monolithic sacrificial support element into two or more pieces to permit removal of the monolithic sacrificial support element from the volume defined by the stiffener. This step may involve mechanically breaking the sacrificial support element into two or more pieces. In some embodiments of the invention a chemically sacrificial monolithic support element may be used, in which case a suitable solvent may be used to break down or dissolve the monolithic sacrificial support element. For example, some materials suitable for use as sacrificial support elements may be configured to be washed away with water, such as that sold under the product name “AquaCore” by Advanced Ceramics Manufacturing.

[0012] The resin-infused fibre preform may comprise fibres pre-impregnated with resin, also known as “pre-preg” Alternatively, the method of manufacturing may comprise a resin transfer moulding process (RTM) process, including a vacuum assisted resin transfer moulding (VARTM) process. As such, the manufacturing process may involve infusing the fibre preform with a liquid resin prior to the step of curing the resin.

[0013] The invention provides a method of manufacturing a composite wing structure having a stiffener with a complex shape that avoids the need for complex tooling assemblies. According to the invention, a fibre preform for forming the stiffener is supported upon a sacrificial support element during the curing process. Once the composite structure is cured and the sacrificial support element is trapped in place by the stiffener, the support element can be broken down and removed from the composite structure. The method therefore avoids the requirement for complex multi-part tooling to be used.

[0014] The composite structure may comprise part of a wing box, for example. As such, the stiffener may be a spar or a stringer of the wing box. In some embodiments, the method may be used to manufacture a composite wing structure having multiple stiffeners. The multiple stiffeners may comprise one or more spars. Alternatively or additionally, the multiple stiffeners may comprise one or more stringers.

[0015] The composite wing structure may have a spanwise direction and a chordwise direction. The spanwise direction may be perpendicular to the chordwise direction. Prior to curing, the fibre preforms used for forming the composite wing structure may comprise a spanwise direction and a chordwise direction. The spanwise direction and the chordwise direction of the fibre preform may correspond to the respective spanwise direction and chordwise direction of the composite wing structure.

[0016] In embodiments of the invention substantially all of the support structure may be formed by the monolithic sacrificial support element. In this case, “substantially all” may mean more than 90 or 95% by volume. 100% of the support structure may be formed by the monolithic sacrificial support element. Substantially all of the surfaces of the resin infused fibre preform for forming the stiffener which are supported by the support structure may be rested against the monolithic sacrificial support element. In other embodiments of the invention, the support structure may comprise a solid tooling element in addition to the monolithic sacrificial support element. More than 50% by volume of the support structure may be formed by the solid tooling element. More than 70% by volume of the support structure may be formed by the solid tooling element. The method may comprise removing the solid tooling element from the open side of the stiffener in one piece.

[0017] In embodiments of the invention, it may be advantageous for the support structure to comprise a solid tooling element where the stiffener is shaped to allow removal of a substantial part of the support structure from the open side of the stiffener in one piece. As such, the support structure may be only partially formed by a monolithic sacrificial support structure, which reduces the waste material associated with the monolithic support structure, which may be single use. The support structure may comprise a solid tooling element and two or more sacrificial support elements which are configured to be broken post-cure.

[0018] The method may comprise resting at least part of a surface of the resin-infused fibre preform for forming the stiffener against the solid tooling element. The method may comprise resting at least part of a surface of the resin-infused fibre preform for forming the stiffener against the monolithic sacrificial support element. The method may comprise resting a first surface of the resin-infused fibre preform against the solid tooling element. The method may comprise resting a second surface of the resin-infused fibre preform against the sacrificial support element.

[0019] In embodiments of the invention, the stiffener may be shaped such that the resin-infused fibre preform can be rested against a solid tooling element which can be removed from the open side of the stiffener in one piece. The solid tooling element may therefore be configurable to minimise the size of the sacrificial support element, and thereby the amount of waste material which results from the use of a sacrificial support element. In such cases, one or more surfaces of the resin-infused fibre preform may be rested against the solid tooling element. Additionally, resting the resin-infused fibre preform against the solid tooling element may provide surfaces of the stiffener which are formed against the solid tooling element with a smoother surface finish relative to the regions of the resin-infused fibre preform rested against the sacrificial support element. Furthermore, this arrangement may provide the possibility of using a solid tooling element, which may be stronger and stiffer than the sacrificial support element, to support the resin infused fibre preform in regions where tool strength and stiffness is important, and where the strength or stiffness of the monolithic support element may be inadequate.

[0020] In some instances, the material forming the sacrificial support element could provide the cured stiffener with a relatively rough surface finish. This may be the case where a foam is used, for example, because resin would seep into the pores of the foam during the curing process. This level of surface finish may not be desirable, particularly in regions of the stiffener where other parts of an airframe need to be bonded or otherwise fastened to the stiffener.

[0021] The support structure may comprise a surface finish element. The surface finish element may be provided upon a surface of the sacrificial support element. The surface finish element may be configured to provide the stiffener with a region having a surface finish that is different to the surface finish that would be provided by the sacrificial support element. The method may comprise resting a surface of the resin-infused fibre preform for forming the stiffener against a surface of the sacrificial support element and a surface of the surface finish element. The surface finish element may be provided between the resin-infused fibre preform for forming the stiffener and the monolithic sacrificial support element. The surface finish element may comprise a material chosen to provide a desirable surface finish. For example, the surface finish element may provide the stiffener with a region having a smooth surface finish, relative to the surface finish that would be provided by the sacrificial support element. This may be advantageous where the stiffener will be fastened to another part. The surface finish element may be configured to provide the stiffener with a region having a surface roughness that is tailored for bonding. Examples of suitable materials include metal, such as steel, composite, or a release film. In some cases, the surface finish element may comprise a relatively small discrete block such that the cured stiffener is provided with a relatively small localised region having a desired surface finish. However, the surface finish element can in principle take on any shape or size as required. In some embodiments of the invention, the monolithic sacrificial support element may be wrapped in a release film, or some other suitable material, to provide a surface finish element that provides the entire stiffener a smooth surface finish.

[0022] The sacrificial support element may comprise a foam. Foam comprises a particularly suitable material for the support element due to its low cost and the ease with which it is able to be broken up into smaller pieces. The foam may be an open or closed cell foam. A closed cell foam may be preferable in some embodiments of the invention in order to limit the penetration of uncured liquid resin into the foam. Any suitable foam may be used, provided that it can support the stiffener during the curing process and can be broken-up post-cure.

[0023] The method may also comprise the step of removing a piece of broken sacrificial support element from the volume defined by the stiffener via the open side of the stiffener.

[0024] The composite wing structure may be configured such that the orientation of a cross-section of the stiffener in a plane perpendicular to a spanwise direction of the composite wing structure changes along a spanwise direction of the composite wing structure. For example, the stiffener may twist along a span wise direction. Alternatively or additionally, the stiffener may be curved in a spanwise direction of the composite wing structure. The stiffener may curve in two perpendicular directions in a plane orientated perpendicularly to a spanwise direction of the wing structure.

[0025] For example, where the wing is a swept wing, the stiffener may be curved to follow the sweep of the wing. Alternatively or additionally, where the wing curves upwardly, as may happen for example at a wing tip, the stiffener may curve upwardly to follow the upward curve of the wing.

[0026] The composite wing structure may be configured such that the cross-sectional dimensions of the stiffener in a plane perpendicular to a spanwise direction of the composite wing structure change along the spanwise direction of the composite wing structure. The stiffener may taper towards the wing tip, for example. As such, one or more dimensions of the stiffener cross-section may reduce along the spanwise direction of the composite wing structure when moving from an inboard location to an outboard location of the composite wing structure. Alternatively, one or more dimensions of the stiffener cross-section may increase when moving from an inboard location to an outboard location of the composite wing structure. It should be understood that an “outboard” location is a position on the composite wing structure that is closer to the very tip of the wing tip than an “inboard” location.

[0027] The stiffener may comprise a first stiffener element and a second stiffener element. The first stiffener element may be spaced apart from the second stiffener element. The first stiffener element may form a foot of the stiffener and the second stiffener element may form a flange of the stiffener, or vice versa. The open side of the stiffener may be defined between the first stiffener element and the second stiffener element. The first and second stiffener elements may be connected by a third stiffener element. The third stiffener element may form a web of the stiffener. The first stiffener element, second stiffener element and third stiffener element may partially define the volume defined by the stiffener.

[0028] The resin-infused fibre preform for forming the stiffener may have elements corresponding to the elements of the cured stiffener that it will form. For example, the resin-infused fibre preform for forming the stiffener may comprise a foot for forming a foot of the stiffener. The foot of the resin-infused fibre preform for forming the stiffener may be supported on a resin-infused fibre preform for forming a substructure of the composite wing structure. The cured composite wing structure may therefore comprise a stiffener bonded to a substructure via a foot of the stiffener. The substructure may comprise a wing skin. The wing skin may be an upper wing skin. The wing skin may be a lower wing skin. The open side of the stiffener may be provided by a space between a flange of the stiffener and a foot of the stiffener. The open side of the stiffener may be provided by a space between a flange of the stiffener and the substructure.

[0029] Any suitable method of breaking the sacrificial support element may be used, as long as the composite wing structure itself is not damaged during the process. In some embodiments, the sacrificial support element may simply be removed by hand. It will be understood that the open side of the stiffener is required for providing access to the sacrificial support element in order to break and remove it. The step of breaking the sacrificial support element into two or more pieces may comprise using a high pressure water jet to break the sacrificial support element. The high pressure water jet may be introduced via the open side of the stiffener.

[0030] The composite wing structure may be a composite wing tip structure. For example, the composite wing structure may comprise part of a wing box. DESCRIPTION OF THE DRAWINGS

[0031] Embodiments of the present invention will now be described by way of example only with reference to the accompanying schematic drawings of which: FIG. 1 shows a prior art aircraft with a wing comprising a composite structure; FIG. 2 is a schematic drawing showing the profile of the wing of the aircraft shown in FIG. 1; FIG. 3 is a schematic cross-sectional view of a composite structure used to provide stiffness to the wing tip of the wing shown in FIG. 1 taken at location LI in FIG. 2; FIG. 4 is a schematic cross-sectional view of the composite structure used to provide stiffness to the wing tip of the wing shown in FIG. 1 taken at location L2 in FIG. 2; FIG. 5 is a schematic cross-section drawing of a composite wing box comprising various types of prior art stiffener; FIG. 6 is a flow diagram showing the steps of a method according to an embodiment of the invention; FIG. 7 is a schematic cross-sectional view of a composite structure for providing stiffness to a wing tip that has been manufactured using a method according to an embodiment of the invention; FIG. 8 is a schematic drawing of dry fibre skin and stiffener preforms laid-up on a surface of a mould; FIG. 9 is a schematic drawing of the dry fibre skin and stiffener preforms enclosed in the mould, ready to be infused with resin and then cured; FIG. 10 is a schematic cross-sectional view of the cured composite structure wherein the sacrificial support elements are being removed using a high-pressure waterjet; FIG. 11 is a schematic partial cross-sectional view of a composite wing box comprising an outwards facing C-spar manufactured according to an embodiment of the invention; FIG. 12 is a schematic partial cross-sectional view of dry fibre stiffener and wing skm preforms used to manufacture the composite wing box of FIG. 11 laid-up with a support structure in a mould; and FIG. 13 is a schematic cross-sectional view of a support structure used to support a C-spar during a method of manufacturing the wing box of FIG. 11. DETAILED DESCRIPTION

[0032] The present invention provides a method of manufacturing a composite wing structure, such as a wing box of the type described above with reference to FIG. 2 to FIG. 4, where the complex multipart tooling of prior art manufacturing methods is replaced with a support structure formed at least in part by a monolithic sacrificial support element. For the avoidance of doubt, the term “monolithic” as used herein is used to describe the sacrificial support element as being formed in one piece, as opposed to being constructed from multiple pieces. Post-cure, where traditional one-piece tooling would be trapped m place by the cured composite structure, the sacrificial support element can be broken-up and removed from the composite structure.

[0033] An example of how a first embodiment of the invention can be used to manufacture the composite wing structure shown in FIG. 3 and FIG. 4 is described below with reference to the steps set out in FIG. 6. The reference numerals used in the following description to describe the composite structure 200’ produced according to the invention, which is shown in FIG. 7, are the same as those used to describe the prior art composite structure 200 described above with the addition of ’ to each reference numeral. It should be noted that, in the following example the composite wing structure is manufactured using a resin transfer moulding (RTM) process. The steps associated with RTM will be well understood by the skilled person and, as such, description of steps of the RTM process have been omitted where they do not relate to the invention.

[0034] With reference to FIG. 8, in a first step 601, a dry fibre skin preform 310 is laid-up on a surface of a lower half of a mould 501. In a second step 602, dry fibre stiffener preforms 320 are laid-up upon the skin preform 310 by supporting each of the stiffener preforms 320 upon respective support structures 400. Each monolithic support structure 400 comprises an elongate monolithic sacrificial support element 401 that has been cut or otherwise shaped to fit in the volume 225’ that will be defined by the respective structural elements 221’, 222’, 223’ of the cured stiffener 220’. The volume 225’ extends in the spanwise direction along the composite structure 200’ perpendicular to the Y’-Z’ plane indicated in FIG. 7. As such, each monolithic sacrificial support element 401 may curve, twist, and alternatively or additionally taper along its length in order to support its respective stiffener preform 320 in a shape that corresponds to the desired shape of the cured stiffener 220’. In the present embodiment, the monolithic sacrificial support elements 401 comprise a foam material. However, in other embodiments of the invention other materials may be used, provided that they are able to support the stiffener preforms during the curing process in a manner that produces cured stiffeners having a suitable tolerance, and provided that the monolithic sacrificial support structure can be easily broken and removed from the cured composite wing structure.

[0035] In the present embodiment, one of the support structures 400 comprises a surface finish element 700 which has been provided on a surface of the monolithic sacrificial support element 401 such that, during the curing process, a surface of one of the stiffeners 220’ is formed against the surface finish element 700. In the present embodiment, the surface finish element is a block comprising, for example, steel or another material that provides the stiffener 220’ with a localised region having a smooth surface finish, relative to the regions where no surface finish element 700 is present. The use of a surface finish element 700 may be particularly advantageous where the material forming the sacrificial support elements 401 results in the cured stiffener having an undesirable surface finish. For example, where a sacrificial support element 401 comprises a foam, the porous surface of the foam may result in the cured stiffener having relatively rough surface finish due to resin seeping into the pores of the foam during the curing process. This relatively rough surface finish may be unsuitable for use as an interface between the stiffener and another part of the airframe where the stiffener is to be bonded or fastened to another part of the airframe.

[0036] The surface finish element may itself have a relatively smooth surface such that it is configured to provide the stiffener with a relatively smooth surface finish where the stiffener is to provide a fastening interface. In embodiments where the stiffener is to provide a bonding surface, the surface finish element may itself have a tailored surface roughness such that it is configured to provide the stiffener with a surface having a complementary roughness that is suitable for bonding. However, in some embodiments, a relatively rough surface finish provided by the sacrificial support structure itself may form a suitable bonding surface, in which case no surface finish element may be needed.

[0037] In some embodiments, the surface finish element 700 may be provided by a release film, or a layer of some other material that produces a desired surface finish. In some embodiments of the invention, the support structure 400 may be wrapped in a release film, or some other suitable material, forming a surface finish element that provides the entire stiffener 220’ with a desired surface finish.

[0038] In the following step 603, the upper mould half 502 is placed over the top of the fibre preforms 310, 320, as shown in FIG. 9. The fibre preforms 310, 320 are then infused with resin and cured in a manner that will be well understood by the skilled person to produce the cured composite wing structure 200’. The skilled person will of course be aware of many different types of fibre and resin suitable for forming the composite structure 200’.

[0039] As described above, the shape of the cured stiffeners 220’ is such that the support structures 400 are trapped and retained within the volumes 225’ defined by structural elements 221’, 222’, 223’. The support structures 400 therefore cannot be removed from the open sides 224’ of the stiffeners 220’ in one piece. Therefore, in a final step 604 shown schematically in FIG. 10, the monolithic sacrificial support elements 401 are broken into multiple pieces 405’ using a high-pressure waterjet 800 and removed from the open sides 224’ of the stiffeners 220’. As will be apparent, this step also permits removal of the surface finish element 700. It will also be apparent that, while a high-pressure water jet 800 provides a convenient means of removing the monolithic sacrificial support elements 401, they can be broken and removed in other ways. For example, they could simply be pulled out by hand. In other embodiments, the monolithic sacrificial support elements may be formed from a material configured to be washed away by a solvent, for example water.

[0040] An example of how a second embodiment of the invention can be used to manufacture the composite wing structure 200” shown in FIG. 11, which comprises a stiffener in the form of an outward facing C-spar 220”, will now be described with reference to FIG. 12. In this case it should be noted that the composite structure 200” is shown in FIG. 11 with an upper wing skin 211”. The reference numerals used in the following description to describe the method of manufacturing the composite structure 200” produced according to the second embodiment of the invention are the same as those used to describe the method of manufacturing the composite structure 200’ according to the first embodiment of the invention but with ” following each reference numeral. Again, it should be noted that in the following example the composite wing structure 200” is manufactured using a resin transfer moulding (RTM) process and that description of steps of the RTM process have been omitted where they do not relate to the invention.

[0041] A dry fibre skin preform 310” is laid-up on a surface of a lower half of a mould 501”. A dry fibre stiffener preform 320” is laid-up upon the skin preform 310” by supporting the stiffener preform 320” upon a support structure 400”. In the example described above substantially all of each of the support structures 400, except for the surface finish element 700, are formed by a monolithic sacrificial support element 401. However, in this case, the support structure 400” comprises a central solid tooling element 402 and two monolithic sacrificial support elements 4011, 4012. A first monolithic sacrificial support element 4011 is positioned below the solid tooling element 402 such that the first sacrificial support element 4011 supports the weight of the solid tooling element 402. A second monolithic sacrificial support element 4012 is positioned above and rests upon the central solid tooling element 402. As can be seen in FIG. 12, each of the sacrificial support elements 4011, 4012 have a triangular cross-section and the solid tooling element 402 has a rectangular cross-section that together form a support structure 400” that fits in the volume 225” that will be defined by the respective structural elements 221”, 222”, 223” of the cured C-spar 220”. Therefore, upper and lower parts of the stiffener preform 320” are supported against the sacrificial support elements 4011, 4012 and a central part of the stiffener preform 320 is supported against the solid tooling element 402.

[0042] An upper mould half is then placed over the top of the fibre preforms 310”, 320”, the fibre preforms 310”, 320” are then infused with resin and cured to produce the cured composite wing structure 200”. During the curing process, surfaces of the first and third structural elements 221”, 223” of the C-spar 220 are therefore formed against the respective first and second monolithic sacrificial support elements 4011, 4012. A central part of a surface of the second structural element 222” is formed against the central solid tooling element 402. Lower and upper parts of the surface of the second structural element 222” are formed against the respective first and second monolithic sacrificial support elements 4011,4012.

[0043] As can be seen in FIG. 11, the shape of the cured C-spar 220” is such that the support structure 400” will be trapped within the volume 225” enclosed by the C-spar 220” due to the acute angles made by the first structural element 221” with the second structural element 222” and by the second structural element 222” with the third structural element 223”. The support structure 400” therefore cannot be removed from the open side 224” of the C-spar 220” in one piece. In the present example, the support structure 400” is configured such that the solid tooling element 402 can be removed in one piece by extracting the solid tooling element 402 from the volume 225” along a horizontal direction V. Once the solid tooling element 402 is removed, the monolithic sacrificial support elements 4011, 4012 can be washed away or broken up using a high-pressure waterjet or removed by some other means, depending on the material used for the sacrificial support elements 4011, 4012.

[0044] An alternative support structure 400”’ for use in place of the support structure 400” in order to manufacture the C-spar 220” according to the method of the second embodiment of the invention is shown in FIG. 13. In this case, the central solid tooling element 402 and the lower monolithic sacrificial support element 4011 of the support structure 400” have been replaced by a single solid tooling element 402”’ having a cross-section in the shape of a parallelogram. The support structure 400’” also comprises a single monolithic sacrificial support element 401’” positioned above and supported upon the solid tooling element 402”’. During the curing process, a surface of the first structural element 221” and first part of a surface of the second structural element 222” of the C-spar 220” are therefore formed against the solid tooling element 402’”. A second part of the surface of the second structural element 222” and a surface of the third structural element 223” are formed against the monolithic sacrificial support element 401”’. As can be seen in FIG. 13, the solid tooling element 402’” is shaped such that it can be removed in one-piece post-cure by extracting the solid tooling element 402 from the volume defined by the C-spar 220” along the direction V”, in a direction parallel with the first structural element 221” of the C-spar. Once the solid tooling element 402”’ is removed, the monolithic sacrificial support element 401”’ can then be broken up and removed.

[0045] Whilst the present invention has been described and illustrated with reference to manufacture of composite wing tip structures 200’, 200” comprising stiffeners 220’, 220” in the form of C-spars, it will be appreciated by those of ordinary skill in the art that the invention lends itself to many different composite structures not specifically illustrated herein.

[0046] In other embodiments, a composite structure manufactured according to an embodiment of the invention may of course have fewer or more than two stiffeners, and the stiffeners may be spars or stringers. Furthermore, the invention can be used to manufacture any suitably shaped stiffener, provided that the cross-section of the stiffener has an open side through which a monolithic sacrificial support element can be accessed in order to break and remove the monolithic sacrificial support element. Other suitable stiffeners that will be known to the skilled person may have T-shaped, I-shaped, J-shaped, or Z-shaped cross-sections. Where a T- or I-shaped stiffener is manufactured, two separate support structures according to the invention may be required to support the stiffener preform during the curing process. In such cases, the monolithic sacrificial support elements may be positioned either side of a web of the stiffener preform in order to support flanges that extend on either side of the web.

[0047] Additionally, while the specific embodiments of the invention described herein involve an RTM process for manufacturing the composite wing structure, other out-of-autoclave processing methods may be used. For example, in some embodiments, vacuum assisted resin transfer moulding (VARTM) may be used. It is also within the scope of the invention for monolithic sacrificial support elements to be used when manufacturing composite structures using pre-preg.

[0048] Where m the foregoing description, integers or elements are mentioned which 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 invention, 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 invention that are described as preferable, advantageous, convenient or the like are optional and do not limit the scope of the independent claims. Moreover, it is to be understood that such optional integers or features, whilst of possible benefit m some embodiments of the invention, may not be desirable, and may therefore be absent, in other embodiments.

[0049] The term ‘or’ shall be interpreted as ‘and / or’ unless the context requires otherwise.

Claims

1. A method of manufacturing a composite wing structure comprising a stiffener having an open side, the method comprising:- supporting a resin-infused fibre preform for forming the stiffener upon a support structure, the support structure comprising a monolithic sacrificial support element,- curing the resin to create the composite wing structure, the composite wing structure being configured such that the stiffener partially surrounds and encloses the support structure, thereby defining a volume in which the support structure is retained by the stiffener such that the support structure cannot be removed from the open side of the stiffener in one piece, and- breaking the monolithic sacrificial support element to permit removal of the support structure from the volume defined by the stiffener.

2. The method of claim 1, wherein the support structure comprises a solid tooling element in addition to the monolithic sacrificial support element and the method comprises removing the solid tooling element from the open side of the stiffener in one piece.

3. The method of claim 2, comprising resting a first surface of the resin-infused fibre preform against the solid tooling element and resting a second surface of the resin-infused fibre preform against the sacrificial support element.

4. The method of any of claims 1 to 3, wherein the support structure comprises a surface finish element provided upon a surface of the sacrificial support element, the surface finish element being configured to provide the stiffener with a region having a surface finish that is different to the surface finish that would be provided by the sacrificial support structure, and the method comprises resting a surface of the resin-infused fibre preform against a surface of the sacrificial support element and resting a surface of the resin-infused fibre preform against a surface of the surface finish element.

5. The method of any of claims 1 to claim 4, wherein the sacrificial support structure comprises a foam.

6. The method of any of claims 1 to 5, comprising removing a piece of broken sacrificial support structure from the volume defined by the stiffener via the open side of the stiffener.

7. The method of any of claims 1 to 6, wherein the composite wing structure is configured such that the orientation of a cross-section of the stiffener in a plane perpendicular to a spanwise direction of the composite wing structure changes along a spanwise direction of the composite wing structure.

8. The method of any of claims 1 to 7, wherein the stiffener is curved in a spanwise direction of the composite wing structure.

9. The method of claim 8, wherein the stiffener curves in two perpendicular directions in a plane orientated perpendicularly to a spanwise direction of the wing structure.

10. The method of any of claims 1 to 9, wherein the composite wing structure is configured such that the cross-sectional dimensions of the stiffener in a plane perpendicular to a spanwise direction of the composite wing structure change along the spanwise direction of the composite wing structure.

11. The method of any of claims 1 to 10, wherein the stiffener comprises a first stiffener element spaced apart from a second stiffener element, and wherein the open side of the stiffener is defined between the first stiffener element and the second stiffener element.

12. A method according to any of claims 1 to 11, wherein the resin-infused fibre preform for forming the stiffener comprises a foot for forming a foot of the stiffener, and the foot of the resin-infused fibre preform for forming the stiffener is supported on a resin-infused fibre preform for forming a substructure of the composite wing structure such that thecured composite wing structure comprises a stiffener bonded to the substructure via a foot of the stiffener.

13. A method according to any of claims 1 to 12, wherein breaking the sacrificial support structure comprises using a high pressure waterjet to break the sacrificial support structure.

14. The method of any of claims 1 to 13, wherein the composite wing structure is a composite wing tip structure.21

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