GB2582148A

GB2582148A - Impact resistant panels

Info

Publication number: GB2582148A
Application number: GB1903365.3A
Authority: GB
Inventors: Tulloch William
Original assignee: Airbus Operations Ltd
Current assignee: Airbus Operations Ltd
Priority date: 2019-03-12
Filing date: 2019-03-12
Publication date: 2020-09-16
Also published as: GB201903365D0

Abstract

A leading edge skin panel (fig.2a,20) has a fibrous reinforcing structure 22,62 embedded in a metallic matrix material 21,61a,61b such as aluminium. The panel is configured to form a foremost aerodynamic surface of an aerofoil, which may be more resistant to impacts such as bird strike. The fibrous reinforcing structure may be a mesh or woven fabric structure containing aluminium oxide fibres, silicon carbide fibres and/or galvanically coated carbon fibres. The reinforcing structure may be distributed across the span of a nose portion (fig.2a,20b), or across the entire area of the panel. The reinforcing structure may be layer sandwiched between an inner and outer layer of metallic matrix material during manufacture, by simultaneously embedding it, either fully or partially, during an extrusion process.

Description

IMPACT RESISTANT PANELS

TECHNICAL FIELD

[0001] The present invention relates to leading edge skin panels for aerofoils, and to methods of manufacturing such skin panels.

BACKGROUND

[0002] The leading edge structures of aircraft aerofoils (wings, tailplanes, vertical stabilizers and the like) may he struck by birds during flight of the aircraft. It is therefore necessary for the aerofoils to be able to withstand damage caused by a bird strike sufficiently well to enable the aircraft to maintain flight at least long enough to reach a landing site.

[0003] For metallic wings, bird strike resistance has conventionally been achieved by forming the leading edge structure from intrinsically strong metallic materials such as high-strength aluminium alloys, and/or by increasing the amount of material comprised in the leading edge structure in order to strengthen it. Composite wings may be strengthened to withstand bird strikes by including auxiliary metallic or composite reinforcement structures and/or by increasing the thickness of the panels forming the leading edge structure. All of these conventional solutions add a significant amount. of weight and cost. to the wing (or other aerofoil), and may undesirably reduce the amount. of space available within the leading edge for housing aircraft systems.

[0004] An aerofoil leading edge structure which is able to withstand a bird strike and which is also relatively lightweight, low cost, space-efficient and easy to manufacture compared to current solutions is therefore desired.

SUMMARY

[0005] A first aspect of the present invention provides a leading edge skin panel for an aerofoil. The skin panel is configured to form an aerodynamic surface of a foremost. portion of the aerofoil, and comprises a fibrous reinforcing structure embedded in a metallic matrix material.

[0006] Optionally, the matrix material is aluminium.

[0007] Optionally, the fibrous reinforcing structure comprises a mesh. Optionally, the fibrous reinforcing structure is a woven structure. Optionally, the fibrous reinforcing structure comprises one or more of: aluminium oxide fibres; silicon carbide fibres; galvanically-coated carbon fibres.

[0008] Optionally, the fibrous reinforcing structure is distributed across substantially the entire spanwise length of the skin panel. Optionally, the skin panel comprises an upper portion, a lower portion, and a nose portion which connects the upper portion to the lower portion, and the fibrous reinforcing structure is distributed across substantially the entire area of the nose portion. Optionally, the fibrous reinforcing structure is distributed across substantially the entire area of the skin panel.

[0009] Optionally, the fibrous reinforcing structure is arranged as a layer and is disposed between an outer layer of the metallic matrix material and an inner layer of the metallic matrix material.

[0010] A second aspect of the invention provides an aerofoil comprising a leading edge skin panel according to the first aspect.

[0011] A third aspect of the invention provides an aircraft comprising an aerofoil according to the second aspect.

[0012] A fourth aspect of the invention provides a method of manufacturing a leading edge skin panel. The method comprises: - providing a fibrous reinforcing structure; - partially or entirely embedding the fibrous reinforcing structure within a metallic matrix material to form a metal matrix composite material; and -forming a leading edge skin panel from the metal matrix composite material.

[0013] Optionally, embedding the fibrous reinforcing structure within the metallic matrix material comprises sandwiching the fibrous reinforcing structure between two layers of the metallic matrix material.

[0014] Optionally, embedding the fibrous reinforcing structure within the metallic matrix material is performed simultaneously with forming a leading edge panel from the matrix composite material, using an extrusion process.

[0015] Optionally, the leading edge skin panel formed by the method is a leading skin panel according to the first aspect.

BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Embodiments of the invention will now he described, by way of example only, with reference to the accompanying drawings, in which: [0017] Figure 1 a is a schematic cross-section through an example aerofoil; [0018] Figure 2a is a schematic cross-section through an example leading edge skin panel according to the invention; [0019] Figure 2h is a close-up view of part of the example leading edge skin panel of Figure 2a; [0020] Figures 3a and 3b arc schematic views of part of an example leading edge skin panel according to the invention, before and after an impact; [0021] Figure 4 is a schematic front view of an example aircraft comprising a leading edge skin panel according to the invention; [0022] Figure 5 is a flow chart illustrating an example method of manufacturing a leading edge skin panel according to the invention; and [0023] Figure 6 shows an example skin panel according to the invention at three different stages during its manufacture.

DETAILED DESCRIPTION

[0024] The examples described herein relate to leading edge skin panels for aerofoils.

Each example leading edge skin panel according to the invention is configured to form an aerodynamic surface of a foremost portion of the aerofoil, and comprises a fibrous structure embedded in a metallic matrix material.

[0025] Metal matrix composites (M MCs) are composite materials comprising elements of a first material (such as fibres) distributed within a matrix of a second, metallic material. The presence of the first material elements typically changes one or more physical properties of the matrix material. Which properties arc altered and to what degree depends on the nature of the first material elements, and/or on the relative proportions of the two materials.

[0026] Embedding a fibrous structure in a metal matrix can confer significantly improved impact resistance properties on the resulting structure, as compared to those of the metal material without the embedded fibrous structure. This is because deformation of the fibrous structure due to an impact can absorb a significant amount of the impact energy. Significantly less energy must therefore be absorbed by the metallic matrix material. The amount of energy that the fibrous structure is able to absorb will depend on the type and arrangement of fibres in the fibrous structure, as will be explained in more detail below.

[0027] Embedding a fibrous structure in a metal matrix typically does not add a significant amount of weight, since the weight of the fibres is usually negligible compared to the weight of the matrix material. However; the increased impact resisting ability of the resulting composite structure can be significant. An aerofoil leading edge skin panel formed from a metal matrix composite material comprising a fibrous reinforcing structure has various advantages over known bird-strike resistant leading edge structures. In particular, the additional strength conferred by the fibrous reinforcing structure enables a lightweight, cheap and easily machinable metal such as aluminium to be used as the matrix material. A leading edge skin panel which has a simple structure (i.e. additional reinforcing structures external to the skin are not required) and is formed from aluminium would he simple and cost effective to manufacture, install, and to replace in service.

[0028] Figure 1 shows a cross-section through a simplified example aerofoil 1 having a conventional structure. A central torsion box is formed by a front spar 11, rear spar 12, upper cover 13 and lower cover 14. This box structure bears most of the load during operation of the aerofoil. A leading edge structure 15 is attached to the front spar, to form the forward part of the aerodynamic profile of the aerofoil. Similarly, a trailing edge structure 16 is attached to the rear spar, to form the rearward part of the aerodynamic profile of the aerofoil. Chordwise extending ribs (not shown) are also provided between the front and rear spars 11, 12, and in the leading and trailing edge structures 15, 16. The leading edge structure comprises a curved skin panel supported on the leading edge ribs. Conventional leading edge skin panels are formed either from metal, or from a resin-based composite material such as carbon fibre reinforced plastic (CFRP).

[0029] Figures 2a and 2h shows a chordwise cross-section through an example leading edge skin panel 20 according to the invention. The skin panel 20 is generally C-shaped, and has an upper portion 20a joined to a lower portion 20e by a nose portion 20b. The nose portion 20b has much greater curvature than the other two portions and includes the foremost point of an aerofoil in which the skin panel 20 is intended to be comprised. The upper portion 20a, nose portion 20b and lower portion 20c are all formed integrally in a single unitary structure. The thickness of the skin panel 20 may be less than or equal to a conventional metallic skin panel for an equivalent application.

[0030] Figure 3h shows a dose-up view of a section A of the skin panel 20. It can he seen from Figure 2h that the skin panel 20 comprises a reinforcing structure 22 embedded within a matrix material 21. In the illustrated example the reinforcing structure 22 is formed as a layer, which is sandwiched between an outer layer of matrix material 21 and an inner layer of matrix material 21. The reinforcing structure 22 is therefore completely embedded within the matrix material 21. Other examples are possible in which the reinforcing structure is only partially embedded within the matrix material 21. In such examples the reinlorcing structure may he located at a surface of the matrix material 21.

[0031] The matrix material may in principle be any metallic material. However; preferably the matrix material is a relatively lightweight metallic material. Preferably the matrix material is relatively easy to machine, in order to facilitate manufacture and installation of the skin panel. In the particular example the matrix material is aluminium. Aluminium may additionally he advantageous for an aircraft application because it is galvanically compatible with neighbouring aircraft structures that the skin panel 20 would contact when installed on the aircraft.

[0032] The fibrous reinforcing structure 22 may comprise a plurality of fibres. Fibres comprised in the fibrous reinforcing structure may be interlinked or interwoven. The fibrous reinforcing structure 22 may comprise a mesh. In some examples the fibrous reinlorcing structure 22 is a woven structure. The fibrous reinforcing structure 22 may comprise a fabric. The fibrous reinforcing structure 22 may comprises one or more of: aluminium oxide fibres; silicon carbide fibres; galvanically-coated carbon fibres; or the like. Parameters of the fibrous reinforcing structure 22 including fibre type, fibre orientation, and fibre density may he selected in dependence on the level of impact force expected to be experienced by a particular example leading edge skin panel 20. The manner in which the fibres are interlinked may also be selected in dependence on the level of impact force expected to be experienced by a particular example leading edge skin panel 20. Custom fibre layups may be created in a similar manner as is known for forming carbon-fibre reinforced plastic (CRFP) components. hi some examples a pre-woven fibre product such as a textile woven from 3MTM NextelTM fibres can he used as the fibrous reinforcing structure 22.

[0033] The fibrous reinforcing structure 22 is distributed across substantially the entire spanwise length of the skin panel. The fibrous reinforcing structure 22 is distributed across substantially the entire area of the nose portion 20h of the skin panel 20. In some examples the fibrous reinforcing structure 22 is distributed across substantially the entire area of the skin panel 20. The distribution of fibres across a region of the skin panel 20 in which the fibrous reinforcing structure 22 is present may be substantially even. Alternatively, the fibres may be unevenly distributed over the skin panel 20, e.g. such that a higher density of fibres is provided in an area expected to receive a greater impact force. For example, the density of fibres near the foremost point of the skin panel 20 may he greater than in more rearward regions of the skin panel 20. In some examples the density of fibres in the nose portion 20b may he greater than the density of fibres in the upper portion 20a and/or the lower portion 20b.

[0034] Figures 3a and 3h show part of an example leading edge skin panel 30 according to the invention, before and after the skin panel 30 has received an impact force F. The skin panel 30 comprises a fibrous reinforcing structure 32 partially embedded in a metallic matrix material 31, such that the reinforcing structure is disposed on an outer surface of the matrix material 31. The reinforcing structure 32 comprises fibres woven into a grid-like mesh. The fibrous reinforcing structure 32 and the metallic matrix material 31 may have any of the features of the example reinforcing structure 22 and matrix material 21 described above.

[0035] In Figure 3a, the skin panel 30 has not yet received any impact force. The fibres of the reinforcing structure 32 are arranged evenly, in substantially straight lines. Figure 3b shows the skin panel 30 after an impact force F has been received by the skin panel 30. The impact force F (indicated by the block arrow) acted in a direction substantially normal to the outer surface of the skin panel 30. The impact force F causes the fibres of the reinforcing structure 32 to deform in a region 33 surrounding the location of the impact.. In particular, the fibres are bent inwardly in this region. Some of the fibres in the region 33 may he stretched. Some of the fibres in the region 33 may have snapped. Some or all of the fibres in the region 33 are axially displaced relative to the matrix material, compared to their position before the impact force F was received. The processes of stretching, snapping and displacing the fibres each absorb part of the impact energy. The matrix material 31 is also deformed in the region 33. The composite material formed by the matrix 31 and reinforcing structure 32 deforms as a uniform surface; however, each element may be design to fail in a different mode, depending on the requirements of a particular application. For example, upon receiving the impact force F the composite material may initially deform elastically, then some or all of the fibres of the reinforcing structure 32 may snap, then the matrix material 31 may plastically deform. The size of the deformed region 33 depends on the magnitude of the impact force F and on the area across which this force is applied.

[0036] Figure 4 shows an example aircraft 400. The aircraft 400 has a pair of wings 402a, 402b joined to a fuselage 401. The aircraft 400 also has a pair of tailplanes (or horizontal stabilizers) 403a, 403b, and a vertical stabilizer 404. The wings 402a, 402b, tailplanes 403a, 403b, and vertical stabilizer 404 all comprise aerofoils which could get. struck by birds during flight of the aircraft. To enable the aircraft to sufficiently withstand such bird strike events, the leading edge of each wing 402a, 402h, tailplane 403a, 403h and the vertical stabilizer 404 is formed by one or more leading edge skin panels according to the invention. Other examples are envisaged in which a leading edge skin panel according to the invention is comprised in one or more of the aerofoils of the aircraft 400, but not necessarily in all of the aerofoils of the aircraft 400. In such examples aerofoils not comprising a leading edge skin panel according to the invention may be reinforced or protected against bird strike in sonic other manner.

[0037] Figure 5 is a flow chart illustrating an example method 500 of manufacturing a leading edge skin panel, which may be a leading edge skin panel according to the invention. In a first block 501, a fibrous reinforcing structure is provided. The fibrous reinforcing structure may have any of the features of the example reinforcing structures 22, 32 described above. Providing the reinforcing structure 22 may comprise forming a woven or otherwise interlinked structure from a plurality of fibres. In some examples, the fibrous reinforcing structure 22 may be sourced as a pre-woven (or otherwise interlinked) structure. Providing the fibrous reinforcing structure 22 may comprise arranging the reinforcing structure 22 on an item of manufacturing equipment suitable for forming a metallic component, such as a casting mould, an extrusion machine, or the like.

[0038] In a second block 502, the fibrous reinforcing structure is partially or entirely embedded within a metallic matrix material to form a metal matrix composite material. The metallic matrix material may have any of the features of the example metal matrix materials 21, 31 described above. In some examples, embedding the fibrous reinforcing structure within the metallic matrix material comprises sandwiching the fibrous reinforcing structure between two layers of the metallic matrix material. In some examples, embedding the fibrous reinforcing structure in the metallic matrix material comprises flowing liquid matrix material around the fibrous reinforcing structure. In some examples the fibrous reinforcing structure is pressed into the matrix material, e.g. during a process of extruding the matrix material. The metal matrix composite material formed as a result of performing block 502 may be in the form of a sheet, a block, a billet, or may be substantially in the form of a leading edge skin panel.

[0039] In a third block 503, a leading edge skin panel is formed from the metal matrix composite material. The skin panel so formed is a leading edge skin panel according to the invention. The skin panel may have any of the features of the example skin panels 20, 30 described above. In examples in which the metal matrix composite material is in the form of a sheet, block 503 may comprise bending the sheet into the shape of a leading edge skin panel (that is, a D-nose shape). In some examples, block 503 may comprise creep/stretch forming the sheet into the shape of a leading edge skin panel. In examples in which the metal matrix composite material is in the form of a billet, forming the leading edge skin panel may comprise extruding the billet into the shape of a leading edge skin panel. In examples in which the metal matrix composite material is substantially in the form of a leading edge skin panel following block 502, performing block 503 may comprise machining the metal matrix composite material, e.g. to create fastener holes, improve the surface smoothness, or the like. In any of the examples, performing block 503 may comprise performing finishing operations such as fettling, polishing, the creation of fastener holes, or the like.

[0040] In some examples blocks 502 and 503 may be performed simultaneously. For example, if an extrusion process is to be used to form the leading edge skin panel, the fibrous reinforcing structure and a billet of metallic matrix material may be fed simultaneously into an extrusion machine, such that the reinforcing structure becomes pressed into the matrix material as a result of the extrusion process.

[0041] Figure 6 illustrates a particular example method of manufacturing a leading edge skin panel according to the invention. The skin panel resulting from this example method has a fibrous reinforcing structure 62 completely embedded within a metallic matrix material 61a, 61b. The embedding is achieved by sandwiching the fibrous reinforcing structure 62 between two layers 61a, 616 or the metallic matrix material. Part (i) of Figure 6 shows the reinforcing structure 62 being arranged on a first layer 61a of the matrix material. The reinforcing structure 62 may he retained on the first layer 61a of matrix material using any suitable technique. For example, the first layer 61a may be oriented horizontally, such that the reinforcing structure 62 remains in a desired position due to gravity.

[0042] Part (ii) of Figure 6 shows the second layer of matrix material 61h being arranged on the reinforcing structure 62. The reinforcing structure 62 is thereby sandwiched between the two layers 61a, 61b of matrix material. In the illustrated example, the first and second layers 61a, 61b of matrix material are the same thickness, such that a substantially equal amount of matrix material 62 is provided on either surface of the reinforcing structure in the resulting metal matrix composite material. The layers may be adhered together using any suitable technique, such as friction stir welding, bonding, hot isostatic press (HIP) forming, or the like.

[0043] Part (iii) of Figure 6 shows the resulting metal matrix composite material. In this example, the metal matrix composite material is in the form of a sheet, having the same thickness as is desired for the skin panel being manufactured. This sheet can then be formed into a leading edge skin panel according to the invention by bending, using any suitable technique known in the art. Further finishing operations such as fettling, polishing, and the creation of fastener holes may be performed, as required, once the metal matrix composite material has been bent into a desired shape.

[0044] Although the invention has been described above with reference to one or more preferred examples or embodiments, it will be appreciated that various changes or modifications may he made without departing from the scope of the invention as defined in the appended claims.

[0045] Where the term "or" has been used in the preceding description, this term should he understood to mean "and/or", except where explicitly stated otherwise.

Claims

CLAIMS: 1. A leading edge skin panel for an aerofoil, the skin panel configured to form an aerodynamic surface of a foremost portion of the aerofoil and comprising a fibrous reinforcing structure embedded in a metallic matrix material.
2. A leading edge skin panel according to claim 1, wherein the matrix material is aluminium.
3. A leading edge skin panel according to claim 1 or claim 2, wherein the fibrous reinforcing structure comprises a mesh.
4. A leading edge skin panel according to any preceding claim, wherein the fibrous reinforcing structure is a woven structure.
5. A leading edge skin panel according to any preceding claim, wherein the fibrous reinforcing structure comprises one or more of: aluminium oxide fibres; silicon carbide fibres; galvanically-coated carbon fibres.
6. A leading edge skin panel according to any preceding claim, wherein the fibrous reinforcing structure is distributed across substantially the entire spanwise length of the skin panel.
7. A leading edge skin panel according to any preceding claim, where-in the skin panel comprises an upper portion, a lower portion, and a nose portion which connects the upper portion to the lower portion, and wherein the fibrous reinforcing structure is distributed across substantially the entire area of the nose portion.
8. A leading edge skin panel according to any preceding claim, wherein the fibrous reinforcing structure is distributed across substantially the entire area of the skin panel.
9. A leading edge skin panel according to any preceding claim, wherein the fibrous reinforcing structure is arranged as a layer and is disposed between an outer layer of the metallic matrix material and an inner layer of the metallic matrix material.
10. An aerofoil comprising a leading edge skin panel according to any of claims 1 to 9.
11. An aircraft. comprising an aerofoil according to claim 10.
12. A method of manufacturing a leading edge skin panel, the method comprising: providing a fibrous reinforcing structure; partially or entirely embedding the fibrous reinforcing structure within a metallic matrix material to form a metal matrix composite material; and forming a leading edge skin panel from the metal matrix composite material.
13. A method according to claim 12, wherein embedding the fibrous reinforcing structure within the metallic matrix material comprises sandwiching the fibrous reinforcing structure between two layers of the metallic matrix material.
14. A method according to claim 12, wherein embedding the fibrous reinforcing structure within the metallic matrix material is performed simultaneously with forming a leading edge panel from the matrix composite material, using an extrusion process.
15. A method according to any of claims 12 to 14, wherein the leading edge skin panel formed by the method is a leading skin panel according to any of claims 1 to 9.