Laminate and method of providing a laminate

Abstract

Laminate comprises layers (400, figure 5) (at least half) having continuous fibers (410, figure 4A) oriented along a length of the laminate and layers (500, figure 5) containing discontinuous fibres (

Description

TECHNICAL FIELD

[0001] The present invention relates to a laminate comprising layers of composite material, a method of providing the laminate, an aircraft structural component comprising the laminate, and a method of fabrication thereof. BACKGROUND

[0002] A laminate is structure comprising a plurality of superimposed layers of material or materials. Laminates comprising composite material are used in aerospace to provide a material with desirable properties for structural applications, such as in aircraft structural components.

[0003] Aircraft structural components, such as elongate aircraft structural components, are used in aircraft to provide structural support and strength to parts of the aircraft. Examples of these include aircraft wing stringers used in aircraft wings along their span, or support elements in the fuselage of the aircraft. SUMMARY

[0004] A first aspect of the present invention provides a laminate comprising layers of composite material, the layers of composite material comprising unidirectional-fibre layers, having continuous fibres orientated along a length of the laminate, and choppedstrand layers, having discontinuous fibres orientated in a plurality of directions within a plane of each of the respective chopped-strand layers, and wherein at least 50% of the layers of composite material in the laminate are the unidirectional-fibre layers.

[0005] Each of the layers in the laminate comprises a composite material, specifically a fibre composite material. Each of the fibre composite materials may comprise pre impregnated (also known as ‘pre-preg’) fibres, or alternatively may comprise ‘dry’ fibres. Dry fibres are supplied without a matrix and may be subsequently infused with a matrix material, such as a resin material.

[0006] Components made of laminates, such as for example composite laminates which comprise layers of composite material, may be used, for example, in aerospace applications, such as aircraft structural components. The composite laminates for use in aircraft structural components conventionally comprise unidirectional-fibre layers at respective varying orientations, such that adequate mechanical properties are provided in all directions. The present invention enables some of these to be replaced with choppedstrand layers. Up to half of the total number of layers may be replaced. Maintaining at least 50% of unidirectional-fibre layers ensures adequate mechanical properties of the laminate. The unidirectional-fibre layers may all comprise fibres aligned in one direction, or may comprise layers with fibres aligned in a plurality of respective directions.

[0007] Replacing some of the unidirectional-fibre layers with chopped-strand layers has the advantage of making use of shorter fibres, such as fibres that are generated as offcuts in production of the unidirectional-fibre layers. This helps to reduce waste and therefore has environmental and cost efficiency benefits. The laminate according to the present invention provides an alternative to the conventional all-unidirectional-fibre laminate for application in aircraft structural components. The presence of chopped-strand layers may also increase the permeability of the laminate during resin infusion, compared to a laminate with only uni directional-fibre layers, when the chopped-strand layers have a lower density compared to the unidirectional-fibre layers.

[0008] The unidirectional-fibre layers in the laminate may all be aligned in one direction, for example, in which greatest strength is needed for a particular application. An example of this would be unidirectional-fibre layers aligned with a loading axis of an elongate aircraft structural component.

[0009] Optionally, at least 60% of the layers of composite material in the laminate are the unidirectional-fibre layers.

[0010] Optionally, the fibres comprise carbon fibres.

[0011] Carbon fibres may have superior properties to other types of fibres that may be used in composite layers, such as for example glass fibres. The superior properties may include, for example, greater strength and greater strength-to-weight ratio. This may be especially beneficial in aerospace applications where high strength and low weight of structural components are beneficial for performance and efficiency. In some examples, at least a majority of the fibres are carbon fibres. In some examples, the fibres consist of carbon fibres.

[0012] Optionally, each of the chopped-strand layers has a thickness that is perpendicular to the plane of the respective chopped-strand layer, the thickness being smaller than a thickness of each of the unidirectional-fibre layers.

[0013] The smaller thickness of the chopped-strand layers compared to the thickness of the unidirectional-fibre layers has the benefit of a laminate that has an overall smaller thickness while preserving strength, for example, compared to a laminate made solely of unidirectional-fibre layers or a laminate comprising a combination of uni directional-fibre layers with chopped-strand layers with equal thickness to the unidirectional-fibre layer thickness. The smaller overall thickness of the laminate may result in a beneficial weight saving, for example in aircraft structural components, and may enable more slimline components to be manufactured to a functional, such as for example aerodynamic, and / or aesthetic benefit.

[0014] Optionally, the laminate comprises a combined total number of the layers of composite material equal to 54.

[0015] Optionally, the laminate comprises 20 of the chopped-strand layers and 34 of the unidirectional-fibre layers.

[0016] Optionally, the laminate comprises a total number of fibres, and wherein at least eight percent of the total number of fibres are orientated in each of four directions, the four directions being: first and second mutually perpendicular directions lying in a plane, and third and fourth mutually perpendicular directions lying in the plane and orientated at 45 degrees to the first and second directions.

[0017] The at least eight percent of the total number of fibres may, for example, be within one or more of the chopped-strand layers. Alternatively, or additionally, the at least eight percent of the total number of fibres may be within one or more of the unidirectional-fibre layers. That is, in some cases, the at least eight percent of the total number of fibres are within a combination of the chopped-strand layers and the unidirectional-fibre layers.

[0018] Optionally, each of the chopped-strand layers is separated from a next adjacent chopped-strand layer of the laminate by at least one of the unidirectional-fibre layers, such as plural ones of the unidirectional-fibre layers.

[0019] Separating the adjacent chopped-strand layers may help increase interlaminar (i.e., between layers of a laminate) damage tolerance, by reducing instances of severe orientation change between two adjacent layers in the laminate.

[0020] Optionally, the laminate has a thickness in a range of from about 5mm to about 15mm.

[0021] A second aspect of the present invention provides an aircraft structural component comprising a web and a flange, wherein the flange extends from the web and comprises the laminate according to the first aspect.

[0022] Such an aircraft structural component will share the benefits discussed with respect to the laminate according to the first aspect.

[0023] Optionally, the aircraft structural component is an elongate structural component having a length, and the continuous fibres of the unidirectional-fibre layers align with the length of the elongate structural component.

[0024] This may help ensure that elongate aircraft structural component has adequate strength in the length direction, given by the continuous fibres of the unidirectional-fibre layers. This is beneficial at this is the direction that is most likely to experience loading, such as for example during bending of the aircraft structural component, in service.

[0025] Optionally, the aircraft structural component is a stringer.

[0026] A third aspect of the present invention provides an aircraft structural assembly, comprising: the aircraft structural component according to the second aspect; and a skin section affixed to the flange of the aircraft structural component.

[0027] A fourth aspect of the present invention provides an aircraft comprising the aircraft structural assembly according to the third aspect.

[0028] A fifth aspect of the present invention provides a method of providing the laminate according to the first aspect, the method comprising stacking the layers of composite material on top of each other.

[0029] A sixth aspect of the present invention provides a method of fabricating an aircraft structural component for an aircraft, the method comprising: providing a web; and providing a flange extending from the web and comprising the laminate according to the first aspect.

[0030] The method permits fabrication of the aircraft structural component according to the second aspect of the present invention.

[0031] Optionally, the aircraft structural component is an elongate aircraft structural component and has a length, and the providing the laminate comprises aligning the fibres of the unidirectional-fibre layers with the length of the aircraft structural component.

[0032] Optionally, in the fifth aspect or the sixth aspect, the method comprises infusing the laminate with a resin.

[0033] Optionally, in the fifth aspect or the sixth aspect, the method comprises providing a stitching interconnecting the layers of composite material.

[0034] Optionally, the flange comprises a wide flange region and a narrow flange region, the narrow flange region having a smaller width than the wide flange region, and wherein the wide flange region comprises the stitching interconnecting the layers of composite material.

[0035] Optionally, the stitching extends along a path in a direction with at least a component parallel to a width of the flange. The stitching interconnecting the layers of composite material is provided to hold the layers of composite material together, for example for transport, during manufacturing steps, during installation and / or in service. The stitching provides a benefit in increased strength of the flange along the path of the stitching. ‘Strength’ may refer to, for example, a degree of resistance to yield, fracture, fatigue, delamination and / or deformation under tensile loading, shear loading, loading in torsion, compression and / or bending. The stitching extending along a path in the direction with at least a component parallel to the width of the flange therefore results in increased strength of the flange in its width direction. The increased strength of the flange in its width direction may, for example, increase the flange’s resistance to loading in its width direction. The increased strength of the flange in its width direction may be tailored by changing a magnitude of the component of the direction of the path of the stitching in the flange width direction.

[0036] Optional features of any one of the aspects of the present invention may be applied equally to any other one of the aspects of the present invention, where appropriate. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Embodiments of the invention will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0038] Figure 1 shows a schematic view of an aircraft according to an example embodiment of the present invention,

[0039] Figure 2 shows an aircraft structural assembly according to an example embodiment of the present invention.

[0040] Figures 3A to 3C show, respectively, a side view of an aircraft structural component according to an example embodiment of the present invention, a cross-sectional view of the same, and a top-down view of the same.

[0041] Figures 4A and 4B show, respectively, a unidirectional-fibre composite layer, and a chopped-strand composite layer.

[0042] Figure 5 shows a laminate according to an example embodiment of the present invention.

[0043] Figure 6 shows a method of providing a laminate according to an example embodiment of the present invention.

[0044] Figure 7 shows a cross-sectional view of the aircraft structural component of Figure 3A comprising the laminate of Figure 5.

[0045] Figure 8 shows a partial cross-sectional view of the aircraft structural component of Figure 3 A.

[0046] Figure 9 shows an example method of fabricating an aircraft structural component, according to an example embodiment of the present invention. DETAILED DESCRIPTION

[0047] Figure 1 shows an aircraft 100 according to an example of the invention. The aircraft has multiple structural assemblies that benefit from improved aircraft structural components. Such assemblies are included in the fuselage 110, the wings 120, and the tail 130.

[0048] A cross section through one of the wings 120 of the aircraft 100 is shown in Figure 2. The wing 120 is an aircraft structural assembly and comprises a skin section 210 defining its outer surface, spars 230 (shown in dashed lines in Figure 2) and ribs 240 (only one of which is visible in Figure 2) supporting the skin section 210, and a series of stringers 300. The rib 240 has openings 242 through which the stringers 300 pass. The stringers 300 are affixed to the inner surface 212 of the skin section 210 at their respective flanges 320 (described below). This is achievable by mechanical joining, for example by mechanical fasteners such as rivets or bolts (not shown). Alternatively, other joining methods such as adhesives or welding can be envisaged. A stringer is an example of an elongate aircraft structural component.

[0049] An aircraft structural component according to an embodiment of the present invention is shown in Figure 3 A. In this example embodiment, the component is one of the stringers 300 shown in Figure 2. In other embodiments, the aircraft structural component may be another component, such as a spar. The stringer 300 (and the rest of the stringers 300) is provided in the wing 120 to provide structural support and must therefore be adapted to withstand the loads the wing 120 is exposed to both on the ground and in flight.

[0050] The stringer 300 comprises a web 310. A flange 320 extends approximately perpendicularly from a first end 314 of the web 310, and the web 310 also has a second end 312 distal from the flange 320. The flange 320 has a length L and a thickness T perpendicular to the length L. The stringer 300 is depicted schematically in a cross-sectional view in Figure 3B, where it is shown that the web 320 also has a width W. The width W is perpendicular to both the thickness T and the length L, and is also perpendicular to the web 310. The length L is parallel to a length of the overall stringer 300.

[0051] A further view of the stringer 300 is shown in Figure 3C, viewed top-down in a direction of the thickness T of the flange 320, with the web 310 shown in dashed lines. In this embodiment, the web 310 and the flange 320 provide the stringer 300 with a upside down ‘T’-shape cross section, as best understood from Figure 3B. Other arrangements are possible, and may be desirable for different structural applications. In other examples, the cross section may, for example, have an ‘L’-shape or an ‘Q’-shape (which can be an omega or ‘acorn’ shape).

[0052] The stringer 300 is made of a carbon fibre composite material arranged in a laminate according to an example of the present invention. Composite materials are employed for aircraft components as they are lightweight and have favourable mechanical properties, such as strength and stiffness.

[0053] Composite materials, such as carbon fibre composite materials or glass fibre composite materials, may be formed from layers stacked to form a laminate structure. An example of a layer of composite material is shown in Figure 4A, and is a unidirectional-fibre layer 400. The unidirectional-fibre layer 400 comprises carbon fibres 410 arranged side-by-side and aligned parallel to each other. A unidirectional-fibre layer 400 is stronger in a direction along the fibres 410 than in other directions. A composite laminate or layup structure would conventionally be composed solely of such unidirectional-fibre layers at different relative orientations to provide adequate strength in all directions.

[0054] A second example layer of composite material is shown in Figure 4B, and is a chopped-strand layer 500. The chopped-strand layer 500 comprises carbon fibres 510 that are relatively shorter than fibres in the unidirectional-fibre layer 400. Each of the fibres 510 may, for example, be 10mm to 100mm long. The fibres 510 may be sourced from off cuts of the uni directional-fibre layers 400, for example. The fibres 510 are arranged within a plane of the layer 500 but are aligned in random directions relative to each other, or at least in a plurality of directions. Thus, a single chopped-strand layer 500 has approximately equal strength in all directions within the plane of the chopped-strand layer 500. The chopped-strand layer comprises eight percent of fibres orientated in each of four directions, the four directions being: first and second mutually perpendicular directions lying in a plane, and third and fourth mutually perpendicular directions lying in the plane and orientated at 45 degrees to the first and second directions.

[0055] A laminate 600 according to an example embodiment of the present invention is shown in Figure 5. The laminate 600 comprises plural ones of each of the unidirectional-fibre layer 400 and the chopped-strand layer 500. In the laminate 600, the unidirectional fibres 410 in all the uni directional-fibre layers 400 are all aligned in one direction, into the page in Figure 5, and the chopped-strand fibres 510 within the chopped-strand layers 500 provide strength in the other directions.

[0056] In this example, the laminate 600 comprises one or two of the unidirectional-fibre layers 400 between pairs of the chopped-strand layers 500, and terminates at each of two opposite faces of the laminate 600 with respective ones of the chopped-strand layers 500. The laminate 600 comprises 34 unidirectional-fibre layers 400 and 20 chopped-strand layers 500, with a total number of the layers of composite material being 54. In this example, each of the chopped-strand layers 500 has a smaller thickness than each of the unidirectional-fibre layers 400, the thickness being a dimension perpendicular to the plane of the respective chopped-strand layer. This means the overall thickness of the laminate 600 is smaller than a corresponding laminate comprising no chopped-strand layers 500 and in which unidirectional-fibre layers 400 are arranged and used to provide strength along all directions. In this example embodiment, the laminate has a thickness of about 10mm.

[0057] In other embodiments, fewer or more of the unidirectional-fibre layers 400 and / or the chopped-strand layers 500 may be provided in the laminate, and / or different ratios of the unidirectional-fibre layers 400 and the chopped-strand layers 500 may be present. However, at least half, and preferably at least 60%, of the total number of layers are to be unidirectional-fibre layers 500. In some examples, different stacking sequences of the unidirectional-fibre layers 400 and the chopped-strand layers 500 are possible. Preferably, pairs of immediately-adjacent ones of the chopped-strand layers 500 are separated by at least one of the unidirectional-fibre layers 400. In other examples, the unidirectional fibres 410 in all the uni directional-fibre layers 400 may not all be aligned in a single direction; they may be orientated in two or more different directions. In other examples, the laminate may have a different thickness, for example the laminate may have a thickness of from 5mm to 15mm.

[0058] An example method 800 of providing a laminate according to an example embodiment of the present invention is shown in Figure 6. In this example method 800, the laminate is the laminate 600 shown in Figure 5 and discussed above. In other embodiments of the method of providing a laminate, the laminate may be an alternative laminate according to an embodiment of the present invention, such as one of those previously described.

[0059] The method 800 comprises providing 810 uni directional-fibre layers. In this example, these are the unidirectional fibre layers 400 discussed above and shown in Figure 4A. In other examples, they may be other unidirectional-fibre layers.

[0060] The method comprises providing 820 chopped-strand layers. These are, in this example, the chopped-strand layers 500, but may be other alternative chopped-strand layers in other embodiments of the method.

[0061] The providing 810 the unidirectional-fibre layers and the providing 820 the chopped-strand layers is performed such that the layers of composite material of the laminate 600 are built up one by one, alternating the providing 810 the unidirectional-fibre layers and the providing 820 chopped-strand layers until a desired stacking sequence is achieved. That is, the layers of composite material 400, 500 are stacked on top of each other. In the example method 800, wherein the laminate is the laminate 600, the method comprises building the layers of composite material ‘bottom-up’ as shown in Figure 5.

[0062] The method 800 comprises providing 830 a stitching connecting the layers of composite material. The providing 830 the stitching comprises tufting together all of the layers of the laminate 600. In other embodiments, the providing 830 the stitching may be omitted, or may comprise sewing or tufting only some of the layers of the laminate.

[0063] The method 800 comprises infusing 840 the laminate with resin, by injecting the resin into the layers of composite material. The infusing the laminate with resin may be omitted in some examples of the method, for example if pre-preg fibre layers are used which are coated with resin before they are laid up or stacked on top of each other.

[0064] Figure 7 shows a cross-sectional view of the aircraft structural component 300 of Figure 3B, to show how the laminate 600 is arranged within the aircraft structural component 300. In other examples, the laminate in the aircraft structural component 300 may be a laminate according to a different embodiment of the present invention.

[0065] The laminate 600 is arranged, in this example, such that the unidirectional fibre layers 500 have fibres 510 aligned with the direction of the length L of the aircraft structural component, into the page in Figure 7. This is not shown for clarity. The laminate 600 is depicted schematically with a reduced number of layers shown, for clarity also.

[0066] The laminate 600 is arranged such that the layers of composite material in the laminate 600 are stacked in a direction of the thickness T of the flange 320, and bent through approximately 90° at the first end 314 of the web 310 to form the web 310 such that the layers within the web 310 are stacked in a direction of the width W of the flange 320 (or thickness of the web 310). In this example, the component 300 is formed by two L-shaped parts formed in this manner, joined at the respective webs, as shown. The two L-shaped parts may be joined at the respective webs by, for example, co-curing or mechanical fastening.

[0067] Other configurations of incorporating a laminate according to an embodiment of the invention into an aircraft structural component are envisaged. For example, in some other embodiments, the aircraft structural component comprises the laminate in only a portion of the component, such as in only the web or in only the flange.

[0068] An example close-up of the flange 320 of the component 300 is shown in Figure 8. In this example, the flange 320 comprises stitching 330 (i.e., a line of stitches) though the thickness T of the flange 320. The stitching 330 connects the layers of composite material of the laminate 600. The stitching 300 lies at an angle 0 from an axis A that is parallel to the thickness T of the flange 320. Thus, the stitching 330 lies along a path with a direction with a component in the direction of the width W and with a component in the direction of the thickness T. By varying angle 9, the component in the direction of the width W may be varied. A greater component in the direction of the width W results in a greater strength provided to the flange 320 in that direction. In the example of Figure 8, angle 0 is about 30°. In other examples, angle 9 may be any other angle, such as an angle greater than 0° and smaller than 90°, such as for example 45°.

[0069] In alternative examples of the component 300, the flange 320 may comprise no stitchings or may comprise a plurality of stitchings (i.e., a plurality of lines of stitches). The plurality of stitchings may be orientated at dissimilar angles 9, and may have a component in a direction of the length L of the flange 320. The stitchings may be located in all of, or only parts of, the flange 320 and / or the web 310. The parts of the flange 320 where the stitching or stitchings are located may be, for example, flange grow-outs where the width W of the flange 320 is greater than in a remainder of the flange 320. The flange 320 may comprise a plurality of such grow-outs spaced apart along the length L of the flange 320. Such grow-outs may be provided, for example, to accommodate attachment sections where the component is to be attached to an aircraft skin section, such as outlined with respect to Figure 2 for example, and may therefore benefit from additional strength provided by the stitching.

[0070] Figure 9 shows an example method 1000 of providing an aircraft structural component, which in this case is the component 300 described above. In alternative embodiments, the method may be a method of fabricating another component, according to an embodiment of the invention.

[0071] The method 1000 comprises providing 1010 a web, which in this example is the web 310.

[0072] The method 1000 further comprises providing 1020 a flange, which in this example is the flange 320. The providing 1020 a flange comprises providing the flange with wide flange regions, which are the grow-outs of the flange 320. In alternative examples, such wide flange regions may be omitted.

[0073] In embodiments of the method of providing an aircraft structural component, the providing 1010 the web may be preceded by, followed by, or happen concurrently with, the providing 1020 the flange. In alternative embodiments of the method of providing an aircraft structural component, only one of the providing 1010 a web and the providing 1020 a flange comprises forming the web or the flange from a laminate according to an embodiment of the present invention, such as for example the laminate 600. The other of the web and flange would be provided in a different manner, such as through use of a different laminate, moulding material, or the like.

[0074] The method 1000 of this embodiment comprises infusing 1030 the laminate 600 in the web 310 and the flange 320 of the component 300 with a resin. The infusing 1030 the component with resin may be omitted in some embodiments, such as if pre-preg fibres are used instead of dry fibres.

[0075] The method 1000 of this embodiment comprises providing 1040 a stitching. This is achieved by sewing with a tufting needle positioned at an angle of the stitching. In other examples, the stitching may be achieved in a different way, such as by an automated sewing process, or may be omitted.

[0076] In embodiments of the method of providing an aircraft structural component, the infusing the component with resin and the providing the stitching may happen in either order, when both are performed. The infusing with resin may comprise curing the resin to harden the resin.

[0077] It is to be noted that the term “or” as used herein is to be interpreted to mean “and / or”, unless expressly stated otherwise.

Claims

1. A laminate comprising layers of composite material, the layers of composite material comprising unidirectional-fibre layers, having continuous fibres orientated along a length of the laminate, and chopped-strand layers, having discontinuous fibres orientated in a plurality of directions within a plane of each of the respective choppedstrand layers, and wherein at least 50% of the layers of composite material in the laminate are the unidirectional-fibre layers.

2. The laminate according to claim 1, wherein at least 60% of the layers of composite material in the laminate are the unidirectional-fibre layers.

3. The laminate according to claim 1 or claim 2, wherein the fibres comprise carbon fibres.

4. The laminate according to any one of the preceding claims, wherein each of the chopped-strand layers has a thickness that is perpendicular to the plane of the respective chopped-strand layer, the thickness being smaller than a thickness of each of the unidirectional-fibre layers.

4. The laminate according to any one of the preceding claims, wherein the laminate comprises a combined total number of the layers of composite material equal to 54.

5. The laminate according to claim 4, wherein the laminate comprises 20 of the chopped-strand layers and 34 of the unidirectional-fibre layers.

6. The laminate according to any one of the preceding claims, wherein the laminate comprises a total number of fibres, and wherein at least eight percent of the total number of fibres are orientated in each of four directions, the four directions being: first and second mutually perpendicular directions lying in a plane, and third and fourth mutually perpendicular directions lying in the plane and orientated at 45 degrees to the first and second directions.

7. The laminate according to any one of the preceding claims, wherein each of the chopped-strand layers is separated from a next adjacent chopped-strand layer of the laminate by at least one of the unidirectional-fibre layers.

8. The laminate according to any one of the preceding claims, wherein the laminate has a thickness in a range of from about 5mm to about 15mm.

9. An aircraft structural component comprising a web and a flange, wherein the flange extends from the web and comprises the laminate according to any one of the preceding claims.

10. The aircraft structural component according to claim 9, wherein the aircraft structural component is an elongate structural component having a length, and wherein the continuous fibres of the unidirectional-fibre layers align with the length of the elongate structural component.

11. The aircraft structural component according to claim 9 or claim 10, wherein the aircraft structural component is a stringer.

12. An aircraft structural assembly, comprising:the aircraft structural component according to any one of claims 9 to 11; and a skin section affixed to the flange of the aircraft structural component.

13. An aircraft comprising the aircraft structural assembly according to claim 12.

14. A method of providing the laminate according to any one of claims 1 to 8, the method comprising stacking the layers of composite material on top of each other.

15. A method of fabricating an aircraft structural component for an aircraft, the method comprising:providing a web; andproviding a flange extending from the web and comprising the laminate according to any one of claims 1 to 8.

16. The method according to claim 15, wherein the aircraft structural component is an elongate aircraft structural component and has a length, and the providing the laminate comprises aligning the fibres of the unidirectional-fibre layers with the length of the aircraft structural component.

17. The method according to any one of claims 14 to 16, wherein the method comprises infusing the laminate with a resin.

18. The method according to any one of claims 14 to 17, wherein the method comprises providing a stitching interconnecting the layers of composite material.19

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