FIELD OF TECHNOLOGY

[0001] This invention relates to an aerostructure component, such as may be employed in the construction of a wing box. The invention further relates to: a wing box and/or a wing made from such a component; an aircraft incorporating such a component, wing box or wing; and to a method of manufacturing an aerostructure component.

BACKGROUND

[0002] A typical aircraft wing is subject to bending forces normal to its spanwise axis; and to torsion forces about its spanwise axis. The conventional wisdom is to react the bending loads by providing one or more spanwise structural beams, such as spars or stringers. Torsion forces are reacted by the skin of the wing. The skin, which is relatively thin, must be stabilised in order to prevent shear loads from causing buckling. The skin is normally stabilised by multiple internal ribs.

[0003] Figure 1 shows part of the span of a conventional aircraft wing 1 comprising a main spar 2, stringers 3 and ribs 4, all covered by a skin 5. Other conventional aircraft wings may have spars at the front and rear of the wing instead of, or in addition to, the main spar 2. As can be seen in the drawing, the wing has many ribs 4, each of which needs to be manufactured and fitted, using a set of close tolerance apertures and a corresponding set of robust fasteners. The process and assembly time for this can be considerable. Furthermore, the ribs 4 may also act as dams to fuel loaded in the wing 1 which leads to undrainable mass impact on the aircraft and complication to the fuel system. Therefore, it is desirable to be able to produce a wing having stiffer skins to resist torsional buckling so that the number of ribs may be reduced.

[0004] To this end, it has been proposed to employ wing skins comprising sandwich panels made of carbon fibre reinforced polymers (CFRP) having a core of foam or honeycomb. However, a problem which may be encountered with using such a sandwich material is that of ensuring the integrity of the bond between the skins and the core, both during manufacture and in service. Several factors may detrimentally effect the bond in service. For example, the core has a tendency to absorb water; in service, such water would freeze and potentially compromise the bond. In-flight impacts against the wing could cause the CFRP skins to undergo interlaminar shear, which would also affect the skin-to-core bond. Therefore, in order to employ such sandwich materials, suitable non-destructive testing (NDT) methods would also need to be developed to test the skin-to-core bond.

[0005] Furthermore, CFRP is not a good electrical conductor and so lightning strike protection in the form of metallic meshes must be provided, adding significantly to the weight and complexity of manufacture of the wing. All metallic fasteners employed in the construction of the wing would need to be electrically connected to the lightning strike protection.

BRIEF SUMMARY OF THE TECHNOLOGY

[0006] The invention provides an aerostructure component comprising a core sandwiched between two skins, each skin comprising a ply of a plurality of elongate reinforcing elements in a metal matrix and the core comprising a plurality of hollow spheres in a metal matrix. The component is inherently both lightweight and strong, prevents the ingress of moisture, contains embedded thermal insulation and is electrically conductive.

[0007] Preferably, the elongate reinforcing elements in the ply are arranged in a predetermined orientation so as to reinforce the component in that direction.

[0008] Advantageously, each skin comprises a plurality of plies; the elongate reinforcing elements of each ply are arranged in a predetermined orientation, and the elongate reinforcing elements in adjacent plies have different respective orientations. Thus, the component can be reinforced across several directions/orientations.

[0009] Preferably, the core also comprises a plurality of elongate reinforcing elements. Such elements can prevent cracks from forming within the core. [0010] Preferably, at least some of the elongate reinforcing elements extend into the core from a ply adjacent the core. This provides reinforcement at the boundary between the skin and the core and helps to prevent crack formation.

[0011] The elongate reinforcing elements comprise a selection of: aluminium oxide fibre; carbon fibre; and silicon carbide fibre.

[0012] The spheres are preferably of metal or ceramic.

[0013] The spheres may be of substantially the same diameter, or of a range of diameters.

[0014] The matrix material may be aluminium or titanium.

[0015] The invention further provides a component as aforedescribed in the form of a cover for an aircraft wing box, the cover comprising a side wall of a wing box and one of an upper or lower surface of a wing box.

[0016] The upper or lower surface of the wing box cover preferably further comprises regions of elongate reinforcing elements arranged in a span-wise orientation. Such regions of elongate reinforcing elements provide greater wing stiffness, thereby reducing the need for stringers and spars.

[0017] The side wall of the wing box cover preferably includes elongate reinforcing elements arranged in orientations of +/- 45° relative to a longitudinal axis of the wing box. Such reinforcing elements act as a truss-like structure to provide stiffness to the side walls.

[0018] The side wall preferably also includes elongate reinforcing elements arranged in an orthogonal orientation with respect to the longitudinal axis of the wing box. This provides extra stiffness between the upper and lower surfaces of the wing box.

[0019] The side wall advantageously further comprises at least one mounting feature for the attachment of another component. This greatly simplifies the mounting of a leading edge and/ flight control surfaces to the wing box.

[0020] The invention further comprises an aircraft wing box comprising an upper cover and a lower cover, each of which comprises a wing box cover as aforedescribed. [0021] Preferably, each cover includes at least one contact surface arranged to abut a corresponding contact surface of the other cover, with the covers being fastened together at the abutting contact surfaces. This greatly simplifies assembly of the wing box.

[0022] The invention further comprises an aircraft including an aerostructure component, a wing box cover and/or an aircraft wing box constructed according to the invention.

[0023] The invention also provides a method of manufacture of an aerostructure component, a wing box cover or an aircraft wing box, the method comprising the steps of: forming at least one ply of elongate reinforcing elements arranged in a predetermined configuration; arranging a plurality of hollow spheres in a predetermined configuration; introducing a matrix material around the reinforcing elements and spheres such that the matrix material at least partially surrounds them; and solidifying the matrix material.

[0024] Preferably, the step of solidifying the matrix material includes the application of external pressure. This ensures a continuous metal matrix from top to bottom surface, and enables the component to be cast into near net shape.

BRIEF DESCRIPTION OF THE DRAWINGS

[0025] The invention will now be described, by way of example, with reference to the accompanying drawings in which:

Figure 1 is a perspective view of the interior of a conventional wing;

Figure 2 is a sectional view of an aerostructure component constructed according to the invention;

Figure 3 is a perspective view of part of a wing box made of a component constructed according to the invention;

Figure 4a is a side sectional view of a complete wing box including the part of Figure 3;

Figure 4b is an enlarged view of part A of the wing box of Figure 4a; Figure 5 is a plan view of an aircraft including an aerostructure constructed according to the invention; and

Figure 6 is a flow chart of a method of constructing an aerostructure component according to the invention.

DETAILED DESCRIPTION OF EMBODIMENTS OF THE TECHNOLOGY

[0026] In the foregoing description, the term “aero structure component” is intended to signify any component that makes up the airframe of an aircraft. This includes, but is not limited to: the panels that make up the fuselage; the empennage; wings (including winglets and wing tips); ribs, spars and stringers; flight control surfaces; fairings; fuel tanks; doors; landing gear; radomes; nacelles; and pylons.

[0027] Figure 2 is a simplified sectional view of an aerostructure component constructed according to the invention and indicated generally by the reference numeral 6. The parts making up the component 6 are not necessarily shown to scale in these drawings. The component 6 comprises a first skin 7, which typically forms an external surface of the component; and a second skin 8 which is typically an internal surface of the component. Each of the skins 7, 8 comprises a metal matrix composite (MMC) material comprising a plurality of plies 9a-9f of elongate reinforcing elements in a metal matrix 36. In this example, the external skin 7 is made up of plies 9a-9d; the internal skin 8 is made up of fewer plies 9e-9f. Any number of plies may be utilised for each skin.

[0028] The term “elongate” is intended to signify any structure having one dimension that is significantly longer than any other dimension. For example wires, fibres, tapes, threads and the like are all considered to be elongate reinforcing elements.

[0029] The elongate reinforcing elements may comprise, for example aluminium oxide fibres, silicon carbide fibres, galvanic-coated carbon fibres, or any other high strength fibres. The matrix material 36 is a metallic material, such as aluminium, titanium, or any other suitable metallic material. It is anticipated that the volume of matrix material 36 will be larger than the volume of elongate reinforcing elements in the component 6. [0030] The elongate reinforcing elements of each ply 9a-9f are unidirectional, arranged in a predetermined orientation. It should be noted that, in these drawings, the direction of a line indicating the presence of a ply does not necessarily correspond to the direction of the elongate elements in that ply. For example, the elongate elements of ply 9a are arranged at an angle of 45° into/out of the plane of the paper. The elongate elements in adjacent plies have different respective orientations. For example, the elongate elements of ply 9b are also arranged at an angle of 45° into/out of the plane of the paper, but are at 90° to the elongate elements of the adjacent ply 9a.

[0031] Sandwiched between the skins 7, 8 is a core 10 comprising a plurality of spheres 11 in a metal matrix 37. The spheres 11 may be of any diameter in the range of micrometres to millimetres. They may all be of substantially the same diameter or a selection of a range of diameters to facilitate denser packing of the spheres. The spheres 11 are hollow and are made from a shell of metal, ceramic or metal-coated ceramic. The spheres 11 may be of one material, such as alumina or silica carbide, or a plurality of different materials in the same composition. The spheres 11 may be coated to enable easier integration with the metal matrix 37. The metal matrix 37 for the core 10 may be same as the metal matrix 36 of the skins 7, 8, or may be of a different metallic material.

[0032] The core 10 further comprises a plurality of elongate elements 12. These elongate elements 12 are typically shorter than those in the skins 7, 8. This group of shorter elongate elements 12 may be of the same material as those in the plies 9a-9g or may be of a different material. In the core 10, the shorter elongate elements 12 are not arranged in predetermined orientations, but are instead oriented randomly within the core. The purpose of this group of elongate elements 12 is to arrest cracks. In time, the metal matrix composite may become fatigued and more prone to cracking. The core’s elongate elements 12 prevent the propagation of cracks within the component 6.

[0033] A proportion of the elongate elements from the plies closest to the core 9d, 9e, may be drawn out of the respective ply to provide another group of elongate reinforcing elements 13 in the core 10. This group of elongate elements 13 are transverse to the elongate elements in the plies, and are preferably aligned approximately orthogonal to them. This group of elongate elements 13 serves to prevent crack propagation in the transition region between the skins 7, 8 and the core 10. [0034] An aerostructure component constructed according to the invention has several advantages over conventional aerostructure components. For example, the component is inherently both lightweight and strong. The high modulus of the reinforcing elements results in a very stiff but low density structure, eminently suitable for aerospace applications.

[0035] The provision of a continuous metal matrix from top surface to bottom surface means that component is also water impermeable and no additional bonding processes are required between the skins and the core. Therefore, there is less of a problem of interlaminar peeling between the skins and the core than is encountered with conventional sandwich panels. The invention permits a hermetically sealed component having a porous core to be manufactured.

[0036] The provision of a core containing hollow metal or metal-ceramic spheres acts as an embedded layer of heat insulation. Therefore, a component constructed according to the invention may be utilised to store substances required to be kept at a predetermined range of temperatures or below a predetermined temperature threshold, such as liquefied gas.

[0037] A component constructed according to the invention is also inherently electrically conductive, thereby providing lightning strike protection. Furthermore, the component may be joined to other components using simple conventional low-cost bolting technologies. The component is also more easily recyclable than CFRP materials.

[0038] Aerostructure components constructed according to the invention can be cast to near net shape in a wide range of sizes, shapes and contours. For example, Figure 3 illustrates an aircraft component constructed according to the invention in the form of an upper cover 14 of a wing box 15. The complete wing box 15 (also known as a torsion box) comprising the upper cover 14 and a similar lower cover 16 is shown in Figure 4a. Figure 4b is a magnified view of a portion of the cross-section of the wing box that forms the upper surface of the wing box 15.

[0039] The upper cover 14 comprises an upper surface 14a and an orthogonal side wall 14b. As can be seen from Figure 4b, the upper cover 14 has a structure comprising an external skin 7 and an internal skin 8 similar to that described above with reference to Figure 2, each skin 7, 8 comprising a plurality of plies 9a-9d; 9e, 9f of elongate elements in a metal matrix 36. The skins 7, 8 form a sandwich with a core 10 comprising spheres 11 in a metal matrix 37. Shorter elongate elements 12 are distributed throughout the core, with some elongate elements 13 being drawn from the plies 9d, 9e immediately adjacent the core 10.

[0040] The wing box upper cover 14 further comprises a plurality of beams 17a- 17e (Figure 3). Each beam is formed from a bundle of plies 18 comprising unidirectional elongate elements in a metal matrix 38, with each elongate element extending spanwise along a major portion of the length of the upper cover 14. The beams 17a- 17e of the upper cover extend along the interior of the upper surface 14a of the wing box 15.

[0041] The metal matrices 36, 37, 38 preferably form one continuous metal matrix from top to bottom of the upper cover 14.

[0042] The lower cover 16 of the wing box comprises a lower surface 16a and an orthogonal side wall 16b. The lower cover 16 is of a similar construction to the upper cover 14. Beams 19a- 19e of the same construction may also be provided in the lower cover 16, extending spanwise along the interior of the lower surface 16a of the wing box 15.

[0043] The beams 17, 19 are arranged to react bending loads experienced by the wing in use. The provision of these beams 17, 19 reduces the need for stringers and spars in the wing, or may even remove the need for stringers and spars completely. This provides a significant savings in assembly time of the wing box, as well as a considerable reduction in weight.

[0044] By varying the number of plies 18 in each beam 17, 19, and the dimensions of each ply, beams of differing cross sections and lengths may be produced, in dependence on the bending loads expected to be experienced by each part of the completed wing in use. The number of beams may also be varied. The beams need not be uniformly spaced chordwise along the wing box.

[0045] Referring back to Figure 3, the upper cover 14 further comprises a side wall 14b, which makes up a side of the completed wing box. The side wall 14b is of a similar construction to the top surface 14a of the upper cover i.e. it comprises skins of elongate elements and a core of spheres, both in a metal matrix. However, the elongate elements are arranged in unidirectional bundles in the metal matrix in the directions shown in the drawing. These elongate elements form a truss-like structure: some bundles of elongate elements are at an angle of approximately +45° to the longitudinal axis of the wing box; some bundles of elongate elements are arranged at an angle of approximately -45° to the longitudinal axis of the wingbox; and some bundles of elongate elements are arranged orthogonally to the longitudinal axis i.e. these fibres extend between the upper and lower surfaces of the wing box. The lower cover 16 preferably has a similar arrangement of elongate elements in a metal matrix making up a truss-like structure in the opposite side wall 16b of the wing box.

[0046] In order to make up a complete wing box 15, the upper cover 14 and lower cover 16 are brought together and may be simply bolted together. In order to facilitate this assembly process, each cover 14, 16 may be provided with flanges arranged to form contact surfaces with similar flanges in the other cover. For example, the upper cover 14 has a first flange 20 that is an extension of its upper surface 14a, and a second flange 21 that extends outwardly at right angles to the end of the side wall 14b. The lower cover 16 has a flange 22 that extends outwardly at right angles to the end of the side wall 16b and is arranged to abut the first flange 20 of the upper cover 14. The other flange 23 of the lower cover 16 extends from the lower surface 16a and is arranged to contact the second flange 21 of the upper cover 14. The flanges 20-23 may be formed wholly or in part from a metallic compound with fewer or no reinforcing elements in order to simplify the drilling and bolting processes.

[0047] The sidewall 14b of the upper cover 14 also includes a mounting feature in the form of a pair of lugs 24, as shown in Figure 3. The lugs 24 are a point of attachment for components that make up the completed wing, such as a leading edge structure in the form of a D-nose. The lugs 24 are formed integrally with the sidewall 14b of the cover 14. The lugs 24 comprise elongate reinforcing elements in a metal matrix. The elongate elements are wound in loops around each aperture 24a, 24b of the lug 24 to provide structural strength and to transfer loads from the lugs to the wing box. Only one set of lugs 24 is shown in this drawing for clarity, but many sets of lugs or other mounting features may be provided on either or both covers 14, 16 in dependence on the desired configuration of the completed wing. For example, other mounting features may be provided on the covers to provide points of attachment for other components, such as other leading edge structures; trailing edge structures; and flight control surfaces, such as slats, flaps, ailerons or spoilers. Each of these components may be pre-equipped with control systems and/or actuators, or they may be provided and attached separately. The provision of attachment points integral to the wing covers allows high-load components to be simply fastened onto the wing box structure. The components attached to the wing box may also be constructed according to the invention, or more conventionally-formed components may be employed.

[0048] The wing constructed according to the invention is lightweight and strong, capable of reacting both bending forces and torsion forces experienced by the wing in use. The wing may be constructed with fewer or no ribs, although it is envisaged that a rib may be required at the root of the wing and at its tip. The removal of ribs from the wing represents a substantial saving in material costs of both ribs and associated components, reduced assembly times and a reduced requirement for fasteners.

[0049] The wing constructed according to the invention is electrically conductive, such that it provides lightning strike protection without the need for metallic meshes or foils that may add to the weight and complexity of the wing. The invention removes the need for bonding strips and similar structures for the fasteners.

[0050] The invention also gives greater flexibility in the location of a fuel tank in the wing: the tank may to be installed at any desired location within the wing box.

[0051] The completed wing 25a is shown on an aircraft 26 in plan view in Figure 5. The aircraft comprises first and second aircraft wings 25a, 25b constructed according to the invention. It will be appreciated that the aerostructure component of the present invention may comprise other features of the aircraft, such as: the panels that make up the fuselage 27; some or all of the parts of the empennage 28 (tail plane, tail fin, elevators and rudder); the engine nacelles 29; the pylons 30 that mount the engines to the wings; the nose cone 31; fairings; landing gear; doors; and/or the flight control surfaces 32 (ailerons, spoilers, flaps, slats, airbrakes etc).

[0052] Figure 6 is a flow chart illustrating an example method of manufacturing an aerostructure component constructed according to the invention. Such a component may be, for example, any of the components mentioned above.

[0053] In one embodiment of a casting method according to the invention, the first step 33 comprises “laying up” of the elongate reinforcing elements and the hollow metal ceramic spheres. This step may be broken down into a series of sub-steps. Firstly, the elongate reinforcing elements making up the skins 7, 8 are placed inside respective moulds. The elongate elements may be conveniently embedded in a fabric or tape that can simply be laid in plies the mould. A plurality of longitudinal elements may be pre-woven or braided into an interlinked arrangement before being arranged in the mould. The laying of the elongate elements may be performed manually, or using an automated tool. Any technique known for arranging fibres in the manufacturing of carbon fibre reinforced plastic (CFRP) components may be used to arrange the elongate reinforcing elements.

[0054] Special mould features may be provided for mounting features, such as the lugs 24. For example, a mandrel may be provided; a bundle of elongate reinforcing elements may be simultaneously wound around the mandrel to form a plurality of loops. Alternatively, a single longitudinal element may be wound around the mandrel multiple times until a desired number of loops is formed.

[0055] Elongate elements 13 are then drawn out of each uppermost ply so that they are upstanding from it.

[0056] The spheres 11 and shorter elongate elements 12 are placed inside one or both moulds. The hollow spheres may be arranged inside the mould, such as through vibrating, to pack the spheres into a best attainable close-packed density and to ensure a distribution of the shorter elongate reinforcing elements throughout the core. Alternatively, the spheres 11 and shorter elongate elements 12 may be made into a preform by using wax binders.

[0057] The next step 34 in the manufacturing process is that of introducing the metal matrix material or materials 36, 37, 38. One way in which this can be done is by injecting liquid metal into the mould at several entry points. In this embodiment, the mould incorporating the spheres and reinforcing material is first pre-heated. Preferably, the pre-heat temperature is approximately equal to the casting temperature of the matrix-forming liquid metal in order to prevent premature solidification of the matrix before complete filling of the mould.

[0058] The matrix-forming liquid metal is cast into the mould in such a manner as to fill the voids around the hollow spheres and reinforcing elements while avoiding disturbance of the spheres and reinforcing elements within the mould. As mentioned above, wax binders may be used to hold the spheres and reinforcing elements in place. In some embodiments, it may be useful to use screens, pegs or other similar means, for maintaining the arrangement of the spheres and elongate reinforcing elements within the mould as the liquid metal is introduced.

[0059] Squeeze casting is a step 35 in the process by which the molten metal solidifies under applied external pressure that is maintained until the end of solidification. The closed dies are positioned between the plates of a hydraulic press. By pressurizing the liquid metals while they solidify, the desired near net shape can be achieved in sound castings. The applied pressure and the instant contact of the molten metal with the die surface produce a rapid heat transfer condition that yields a fine-grain casting with mechanical properties approaching those of a wrought product. The cooling of the mould can be effected through atmospheric cooling or through more controlled cooling methods.

[0060] The component may be formed by other suitable processes known to the skilled person, such as powder metallurgy or diffusion bonding.

[0061] The component may then be machined, drilled and assembled as required. As mentioned above, the parts that are intended to be drilled may be formed with fewer or no reinforcing elements so as to facilitate drilling.

[0062] Variations may be made without departing from the scope of the invention. For example, other elongate reinforcing elements suitable for use in a component constructed according to the invention include galvanic-coated carbon fibres, basalt fibres, boron fibres, or any other high strength fibres. Other suitable matrix materials include alloys of titanium, aluminium, steel, copper and nickel, and metal ceramics materials. Further variations of the invention will be apparent to the person skilled in the art.