WO2006072758A2

WO2006072758A2 - Fibre metal reinforced composite structure

Info

Publication number: WO2006072758A2
Application number: PCT/GB2005/004818
Authority: WO
Inventors: Robert Samuel Wilson
Original assignee: Short Brothers Plc
Priority date: 2005-01-10
Filing date: 2005-12-14
Publication date: 2006-07-13
Also published as: WO2006072758A3; GB0525447D0; GB2421926A; GB2421926B

Abstract

A fibre metal reinforced structure has a fibre truss structure (3) sandwiched between first and second fibre layers (1, 2) and (5, 7). A metallised surface (6) is applied to an inner and/or outer surface of the truss structure and the composite structure is bonded together by resin that is infused through the composite and subsequently cured.

Description

FIBRE METAL REINFORCED COMPOSITE STRUCTURE

The present invention relates to a fibre metal reinforced composite structure and to a method of producing the same with particular, but not exclusive, application to components for use in aircraft construction.

Fibre metal composite structures began to be developed in the 1980's consisting of alternating sheets of aluminium and fibres, for example glass fibres. One example of such a composite structure is known as Arall and comprises thin sheets of aluminium alternated with thin sheets of aramid/epoxy. Another is known as Glare, which comprises alternating layers of glass/epoxy and aluminium. These composite structures were primarily developed as an alternative to aluminium in aircraft structures and have been adopted for use in the Airbus A380 aircraft. A fibre metal composite structure comprising alternating layers of titanium and unidirectional graphite has also been developed. Production of fibre reinforced composite structures by resin infusion of fibre structures formed around a number of mandrels is disclosed in WO03/103933.

The present invention seeks to provide a method of manufacturing metal enhanced fibre reinforced composite structures that provides improved performance over fibre reinforced composites and is less expensive and easier to produce than existing fibre metal laminate composites.

According to one aspect of this invention there is provided a fibre metal reinforced structure including at least one fibre truss structure and on opposing sides of said truss structure respective first and second fibre layers, wherein a metallised surface is applied to at least one of an inner and outer surface of said truss structure, and said composite structure is bonded together by resin that has been infused through the composite structure and cured.

Preferably, the truss structure is formed with a plurality of fibre braided box structures which may conveniently have a quadrilateral section, e.g. rectangular, square or trapezoidal. In said composite structure said resin is infused between first and second fibre layers and said truss structure, and between adjacent box structures. Conveniently, said first and second fibre layers each include at least one layer of fibre material.

Advantageously, the first and second fibre layers each include two layers of fibre material. Preferably, said fibre material is a carbon fibre sheet.

Advantageously, said metallised surface is at least one of a metal foil, shim, strip or tape, and expanded metal mesh, and metal coated fabric, and woven metal fibres/yarn and fibre fabric incorporating metal woven yarn.

Advantageously, the metal is one of titanium, aluminium, aluminium alloy, stainless steel, and low thermal expansion nickel alloy.

Conveniently, the fibre layers are made of fibres including at least one of carbon, carbon/epoxy, carbon/bismalemide, carbon/cyanate, ester, carbon/polyimide, glass, aramid, high density polyethylene, polypropylene, Zylon^R™ (PBO), ceramic boron. Such fibres may be incorporated with metal.

So as to strengthen said composite structure against birdstrike, preferably at least one shim plate is located with a minor surface thereof on a leading surface, in use, of said truss structure.

Conveniently, said shim plate is oriented one of transversely and longitudinally of said composite structure.

Preferably, a metallic surface is located one of between said first fibre layer and said truss structure, between said second fibre layer and said truss structure, and on an outer surface of said first and second fibre layers.

Advantageously, a metallic member is located between adjacent box structures with opposing major surfaces of said metallic member adjacent a respective box structure, wherein an interstitial space is provided between said metallic member and each adjacent box structure into which said resin is infused.

In an embodiment, a fibreglass layer is provided between said first and second fibre layers and said truss structure. In another embodiment, a fibreglass layer is provided on an outer surface of at least one of said first and second fibre layers. In a further embodiment, a metallised member is located on an outer surface of at least of one of said first and second fibre layers.

Advantageously, said metallised member and/or said metallised surface is one of a titanium sheet, a titanium spray coating, a perforated sheet and a metallised fabric.

Titanium is currently preferred as the metallised member from a weight, thermal, and galvanic property point of view.

Advantageously, said composite structure is formed into components of an aircraft. According to a second aspect of this invention there is provided a process of making a fibre metal reinforced composite structure having at least one fibre truss structure and on opposing sides thereof respective first and second fibre layers, including the steps of forming a metallised surface to at least one of an inner and outer surface of said truss structure, infusing resin through the truss structure in a vacuum and curing the resin so as to bond the structure together.

Preferably, the metallised surface is formed by a metal tape which is applied by an automatic tape laying process.

A metallised member on an outer surface of at least one of the first and second layer is desired. A metallised surface improves the composite properties in a number of ways including:

1. Resistance to impact

2. Resistance to scratches

3. Elimination of the need for a strike protection mesh 4. Elimination of need for fibre break out plies (plies on surface of UD laminates which prevent splintering during drill break through)

5. Improved smoothness and surface finish

6. Better surface for paint adhesion/removal

7. Avoidance of porosity, the problem of delaminate glare (due to the non-porous foil trapped in the laminate trapping air, etc.)

8. Improved HIRF/EMI (High Intensity Radio Frequency/Electro Magnetic Interference) characteristics 9. Lower cost than an entirely metallic structure

10. Improved burn through characteristics

11. Improved core properties, e.g. compression, transverse shear, and bond strength to skin; and 12. Improved weight, corrosion and fatigue resistance with respect to solely metallic structures 13. The layers of titanium act as a barrier in a high temperature burst duct situation

A metallised surface on an inner surface of the imier and/or first and second fibre layers enhances impact properties of the composite structure. Further, any effects of thermal expansion mismatch, i.e. between the metal material and the fibre material, may be reduced. This is further enhanced by choosing metal and fibre materials having closely matching thermal expansion coefficients, as far as possible. Compensation for dissimilar thermal characteristics may also be achieved by examples of typical coefficient of thermal expansion (cte) (will vary with fibre orientations) Aramid/epoxy = 0, carbon/epoxy = 3.4, boron fibre/epoxy = 4.3, fibre glass/epoxy = 10, titanium = 10, steel = 10.8, aluminium = 23, low cte Nickel alloy = 3.4.

The foregoing has the benefit of facilitating the infusion/injection process while further enhancing the damage tolerance of the laminate. In addition, in some cases, for example fibreglass, can act as a galvanic corrosion isolation layer if aluminium alloy components are to be mechanically attached to the inner surface of the first or second layer.

The use of dissimilar materials may form a galvanic cell which can create corrosion in the presence of moisture. As a consequence, the choice of materials having their galvanic properties close to each other is preferred.

Advantages of the present invention can be illustrated by reference to a titanium and carbon/epoxy composite which is a preferred composite, although the advantages are also applicable to other metal fibre reinforced composites produced by the method of the invention. Resistance to impact/scratches/erosion

Titanium is harder than carbon/epoxy (resisting scratches, erosions, etc.) with similar stress/strain properties up to yield, after which strain to failure of the titanium is much better than the composite (resisting penetration).

Elimination of the need for a lightning strike protection mesh having HIRF/EMI characteristics.

Titanium is a much better electrical conductor than carbon/epoxy giving the lightning a clear path to dissipate its energy without significant damage to the advanced composite.

Better burn through resistance

Titanium will fail at a much higher temperature than the fibre layers.

Improved weight, corrosion and fatigue resistance in distinction to wholly metallic structures This invention capitalises on the excellent weight, corrosion resistance, and fatigue resistant properties of advanced composites with minimal penalty through the use of metallised material.

Improved smoothness and surface finish Shim/foil (0.13 mm) (0.005") thick is sufficient thickness to be unaffected by surface characteristics in the carbon fabric that can create small voids; the voids are filled with resin by the infusion process.

In addition to enhancement of the general performance properties of the composites, as outlined above, composites of the present invention also show improved resistance to bird strike damage. This aspect of performance can also be enhanced by applying metallic shim or tape to the walls of the truss structures transverse, for example substantially normal to a major plane of the composite. Such performance may further be enhanced by providing thicker metal, for example titanium plates, between the braids of the truss structures in a direction substantially normal to the direction in which an impact can be expected.

According to a feature of this invention there is provided a composite structure as defined above formed into components of an aircraft, such as a fuselage, floor panels, leading edge structures, and stabiliser skins.

According to another aspect of this invention there is provided a process of making a fibre metal reinforced composite structure having at least one fibre truss structure and on opposing sides thereof respective first and second fibre layers, including the steps of forming a metallised surface to at least one of an inner and outer surface of said truss structure, infusing resin through the truss structure and curing the resin so as to bond the structure together.

Conveniently, the metallised surface is formed by a metal tape which is applied by an automatic tape laying process.

Embodiments of 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 a known composite fibre-reinforced truss stiffened structure,

Figure 2 is a perspective view of a fibre metal reinforced composite structure according to one embodiment of the present invention, Figure 3 is a perspective view of another fibre metal reinforced composite structure according to the present invention showing an arrangement of metal reinforcement effective for enhancement of bird strike resistance,

Figure 4 is a schematic diagram of preferred resin transfer infusion equipment suitable for producing composite structures according to the method of the present invention,

Figures 5a and 5b are schematic diagrams showing resin flow through the composite structure assembly during resin transfer of the structures shown in Figures 2 and 3 respectively,

Figure 6 is a perspective view of a further embodiment of a fibre metal reinforced composite structure according to the present invention, Figures 7a - 7c are schematic diagrams showing resin flow through alternative structures during resin transfer of composite assemblies which provide enhanced bird strike resistance,

Figure 8 is a perspective view of yet another fibre metal reinforced composite structure according to the present invention including additional fibre layers to compensate for thermal expansion disparities.

Figure 9 is a schematic diagram showing yet a further embodiment of the present invention offering further enhanced bird strike resistance,

Figures 10a and 10b are schematic diagrams showing an arrangement of lateral and longitudinal ribs for enhancement of bird strike resistance,

Figure 11 shows a perspective view of another composite structure according to this invention, and

Figure 12 shows a 100 x enlargement of a fabric useful in this invention.

In the Figures like reference numerals denote like parts. Referring to Figure 1, there is shown a known reinforced composite structure having an upper skin formed of carbon fibre sheets 1 and 2, plural carbon fibre braided box structures 3 laid side by side and lower skin carbon fibre sheets 4 and 5, the box structures acting as trusses.

In Figure 2 there is shown a fibre metal reinforced composite structure of the present invention having upper skin carbon fibre sheets 1 and 2, plural carbon fibre braided trusses, for example box structures 3 laid side by side and lower skin carbon fibre sheets 4 and 5. It is to be understood that this invention is not limited to two fibre sheets forming each of the upper and lower skins, but may have one or plural sheets forming each skin. The structure additionally has titanium foil, shim, strip, or tape 6 applied to an inner face, lower surface, of the carbon fibre braided boxes, The plain strip or tape may have a thickness in the range 0.08 - 0.18 mm (0.003" - 0.007"). There is additionally provided a titanium shim 7.

In Figure 3 there is shown another embodiment of a metal fibre reinforced composite structure of the present invention which advantageously provides improved resistance to bird strike damage and comprises upper skin carbon fibre sheets 1 and 2, carbon fibre braided box structures 3 and lower skin carbon fibre sheets 4 and 5. In addition, a titanium shim, strip or tape 8 is applied between the braided carbon fibre box structures transverse to, for example substantially normal to the major faces of the upper and lower skins.

Preferably, the structure is produced using low pressure resin transfer moulding equipment employing a resin transfer infusion process, such as shown substantially in Figure 4, and described in detail, in GB 2316036A.

A mould 26 has a hard base 28 having a laid-up region on which the elements of a composite panel structure are laid up as an assembly 27. An elastomeric bagging blanket 30 extends over the composite panel assembly 27 and sealingly cooperates with the hard base 28 at its outer peripheral edges to form a sealed enclosure 29 which encloses the composite panel structure assembly 27.

Typical elastomers for the blanket 30 include polyacrylic fluoro-elastomer or silicone, i.e. elastomers having good vacuum integrity.

The material of the bagging blanket 30 is selected to provide the blanket with a "soft" area 32 which enables the blanket to expand, and avoids the need for providing expansion folds or tucks to allow the blanket to expand.

A preferred material for combining with the elastomer to form the soft area 32 is a dry knit of, for example, glass fibres. The dry knit is bonded or mechanically keyed to the surface of the elastomer or encapsulated with the elastomer.

The blanket 30 also includes "semi-stiff areas 34, 36. The first semi-stiff area 34 forms a central region of the blanket 30 which registers with the fibre reinforced composite panel structure laid-up assembly 27 while the second semi-stiff area 36 forms the outer peripheral area of the blanket 30 which sealingly cooperates with the hard base 28. Conveniently, the semi-stiff area 34 contains as the elastomer reinforcement a prepreg fibre assembly having a coefficient of thermal expansion which is compatible with that of the fibre reinforced composite component to be formed thereby facilitating mould release.

The equipment further comprises a liquid resin supply line 38, of which one end thereof is connected to a liquid resin inlet port 40 in the laid-up region of the hard base 28, and which, at its opposite end, is connected to a liquid resin supply 42. In addition, the equipment also includes a vacuum supply line 44 which at one end thereof is connected to a vacuum outlet port 46 in the laid-up region, and which, at the opposite end thereof, is connected to vacuum generation means (not shown) for applying a vacuum pressure of typically 4 millibar to the sealed laid-up region.

The application of vacuum pressure to the sealed enclosure causes liquid resin to be drawn or injected into the sealed enclosure 29 from the liquid supply 42 to form a liquid resin/reinforcing fibre laid-up assembly system in the sealed enclosure 29. This injection of liquid resin into the sealed enclosure can also be assisted by applying positive pressure to the liquid resin in the liquid resin supply 42. Application of vacuum pressure to the sealed enclosure 29 further acts to prevent air becoming trapped in the liquid resin/reinforcing fibre laid-up assembly system. The application of vacuum pressure in the sealed enclosure through the vacuum outlet port 46 results in liquid resin being drawn into the laid-up region and impregnating the laid-up assembly 27.

The mould 26 is located in an autoclave 24 to control the temperature in the sealed enclosure so that the viscosity of the liquid resin is maintained at a reduced value which allows wet-out of the laid-up assembly.

After completion of the liquid resin injection stage, the autoclave 24 is used to apply an external pressure to the blanket 30 to cause the blanket 30 to apply a consolidating force to the liquid resin/reinforcing fibre laid-up assembly in the sealed enclosure 29 while maintaining control of the temperature in the sealed enclosure to keep the liquid resin at a reduced viscosity so as to enable full impregnation of the reinforcing fibre laid-up assembly with liquid resin. For this consolidation stage the vacuum pressure is withdrawn and the liquid resin inlet port 40 is used for ejecting excess liquid resin from the sealed enclosure 29 to a resin dump (not shown) under the action of a consolidating force applied to the liquid resin/reinforcing fibre laid-up assembly by the blanket 30.

After the consolidation stage has been completed the external pressure applied to the blanket and the temperature in the sealed enclosure 29 are controlled by a compressor 48 and heaters 50 to cure the liquid resin impregnated into the reinforcing fibre laid-up assembly 27 and thereby to form a metal fibre reinforced resin composite structure.

The metal shim, strip, or tape can be applied to any surface of the composite structure provided that a resin flow path is present so as to permit infusion of resin throughout the complete structure assembly during resin infusion. Figures 5a and 5b show resin flow 51, 52, 53 through the component assemblies, as shown in Figures 2 and 3 respectively.

By placing metal shim, strip or tape throughout and around the trusses or skins, various design options are possible. Figure 6 shows shim plates 9 on the inside of the upper surface of the carbon braided box structures 3.

Enhancement of bird strike resistance is greatest when the metal shim is applied either inside and/or outside (between) the walls of the carbon fibre box structures, i.e. normal to the direction of impact during forward motion of an aircraft comprising composite structures of the invention. Figures 7a - 7c show resin flow through a composite assembly including titanium shim 10 applied to inner walls of the braided box structures 3 (Figures 7a, 7b) and titanium shim 8 applied between carbon fibre box structures 3 (Figure 7c).

As noted above, differences in thermal expansion characteristics can be compensated by application of metallic layers. In addition, fibre reinforced plies which, when infused with resin, exhibit different thermal expansion properties, i.e. with a higher coefficient thermal expansion process, for example fibreglass epoxy, can also be incorporated. Figures 8a and 8b show a composite assembly incorporating fibreglass layers. In Figure 8a there is shown a fibreglass layer 12 incorporated between the inner skin formed by the lower walls of the carbon box structures 3 and the lower outer skin formed by carbon fibre layers 4 and 5. In Figure 8b there is shown a fibreglass layer 13 on the outer face of the upper skin formed by carbon fibre layers 1 and 2.

In addition to the enhancement of bird strike resistance already described, resistance can be further enhanced by addition of thicker titanium plates between the carbon fibre box structures forming the truss walls, as shown in Figure 9, where titanium plates 14 are incorporated between the carbon fibre box structures 3. The plates act as a series of mini ribs or riblets arranged in the direction of impact to absorb impact energy, the ribs being arranged with the major surfaces thereof substantially normal to the direction of impact. The ribs can be arranged laterally or longitudinally depending on the arrangement of the carbon fibre box structures in the component assembly. Figures 10a and 10b show a lateral and longitudinal arrangements of ribs 15 respectively (the fibre braided box structures not being shown for clarity). The titanium plates/ribs between the fibre braided box truss structures can replace multi rib designs currently employed which require mechanical fastening to component structures. Composite construction can be simplified since the riblets of the present invention do not require flanges or mechanical fastening.

Composites produced by methods of the present invention can be used to replace some composites currently manufactured entirely of titanium. Titanium is very difficult to manufacture, i.e. form, drill, etc. Composite components produced by the methods of the present invention use a relatively thin shim, strip, tape or foil which is much easier to bend to shape or cut.

Metallic shim will not drape readily into '3D' shapes. The methods of the present invention are particularly useful in production of Aerospace components that are either flat or have simple '2D' curvatures, for example:

1. Fuselage skins 2. Floor panels

3. Leading edge structures

4. Stabilizer skins

In certain cases this drape limitation can be overcome by:

1. Choosing a design where the metallic layer is only applied to the outer surface (Figures 8a and 8b, inclusion of fibreglass plies) and

2. Where the titanium is initially placed applied to the mould either by a spray coating or pre-stretch forming to shape.

3. The application of perforated shim or metallised fabric (as hereinafter described).

The application of a metallic shim to strategic areas of a resin infused truss stiffened skin can also be achieved with a pre-impregnated (prepreg) outer skin, where the truss structure is produced by a resin film infusion technique.

The metal fibre reinforced composites of the present invention can advantageously be employed in the manufacture of fuselage and leading edge structures. Composite structures using trusses provide a number of advantages when employed in construction of fuselages including ease of joining; repair benefits including the ability to drain moisture ingress in a through penetration event, systems access (pneumatic, electrical), etc. The metal fibre composite structures produced according to the present invention offer further advantages in production of an advanced composite pressurised fuselage. These advantages include:

Significant weight saving (from 20% to 35%).

Excellent fatigue and corrosion resistance.

Increased internal volume (space) for a given outside diameter. Potential to pressurise at 1 ,829m (6,000 ft.) instead of 2,438m

(8,000 ft.) (no corrosion issues due to condensation).

Opportunity for component integration and reduce fastener count.

There are significant differences between the inside cabin temperature (room temperature to 3O⁰C) and the outside temperature (-55⁰C) which can result in condensation. The present invention has two significant benefits in that any condensation can be drained away (independent of impact damage) (it is claimed that some commercial aircraft can carry many thousands of kilograms (tons) of parasitic moisture in insulation because of this) and the cavity will act as a better lighter insulator than foam insulation.

Titanium outer skin and fibreglass inner skin plies, such as shown in Figures

8a and 8b, may be advantageously employed in this application because, as well as the other benefits, the fibreglass can also increase the insulation properties as, for example, shown in Figures 8a and 8b.

In addition, the application of a titanium shim to the other internal skin surfaces may be advantageously employed to neutralise any change in coefficient of thermal expansion effects caused by the difference in external and internal temperatures, as shown in Figure 11, where fibreglass layer 16 is applied to the outer surface of the upper skin having carbon fibre layers 1 and 2, and shim plates 17 and 18 are applied to the inner upper and inner lower surfaces of the carbon box structures 3. A titanium shim plate 19 is applied to the outer face of the lower skin formed of carbon fibre layers 4 and 5.

In all the foregoing embodiments where dry fibre glass layers are used, the following alternatives may be employed. 1. An expanded mesh made of, for example, titanium or other metals such as stainless steel, copper, bronze or the like, such mesh being made by Dexmet Corporation of Branford, Connecticut, U.S.A.

2. Titanium coated carbon fibres woven or braided into a fabric.

3. A fabric or braid manufactured using metal fibres/yarns, e.g. stainless steel, titanium, etc., as shown in Figure 12.

4. A carbon fibre fabric or braid with some selected metallic yarns. Because the resin is able to flow through the foregoing, there is greater freedom to interleave layers without affecting the infusion process.

In the embodiments of Figures 2, 3, 6, 8 A, 8B, the provision of a titanium shim may also be replaced by the above-noted expanded mesh, titanium coated carbon fibres woven into a fabric, fabric manufactured using metal fibres/yarns, or carbon fibre fabric with selected metal yarns. The advantages of such a structure are: Surface treatment may not be required (or may not be as critical), since due to the resin infiltration into the mesh or metallic fibre, a metallic bond is achieved. There is a potential for weight saving.

There is a potential to overcome some of the drape limitations of the shim/plate.

The metal fibre reinforced composite structure of the present invention can also be advantageously employed in the manufacture of leading edge structures. The truss structure can carry anti-icing fluid or bleed air. Perforations for release of anti- icing fluid or bleed air can conveniently be produced in the outer skin of the component using laser drilling, such as is described in GB 2364366A. Preferably, the laser drilling system is an Excimer, such as described in GB 2364366A. The perforations can be used for release of anti-icing fluid (hot air or liquid) or for aerodynamic reasons (preventing shock waves or laminar flow) or acoustic attenuation purposes. The layers of titanium will act as a barrier in a high temperature burst duct situation. The present invention may be applied to the production of, essentially, any fibre metal composite structure, but is particularly useful in production of fibre reinforced truss structures for use in aircraft and for structures such as those described in WO03/103933.

Claims

CLAIMS:

1. A fibre metal reinforced structure including at least one fibre truss structure and on opposing sides of said truss structure respective first and second fibre layers, wherein a metallised surface is applied to at least one of an inner and outer surface of said truss structure, and said composite structure is bonded together by resin that has been infused through the composite structure and cured.

2. A structure as claimed in claim 1, wherein the truss structure is formed with a^' plurality of fibre braided box structures.

3. A structure as claimed in claim 2, wherein each structure has a quadrilateral section.

4. A structure as claimed in any preceding claim, wherein said resin is infused between first and second fibre layers and said truss structure, and between adjacent box structures.

5. A structure as claimed in any preceding claim, wherein said first and second fibre layers each include at least one layer of fibre material.

6. A structure as claimed in any preceding claim, wherein the first and second fibre layers each include two layers of fibre material.

7. A structure as claimed in any preceding claim, wherein said fibre material is a carbon fibre sheet.

8. A structure as claimed in any preceding claim, wherein said metallised surface is at least one of a metal foil, shim, strip or tape, and expanded metal mesh, and metal coated fabric, and woven metal fibres/yarn and fibre fabric incorporating metal woven yarn.

9. A structure as claimed in claim 8, wherein the metal is one of titanium, aluminium, aluminium alloy, stainless steel, and low thermal expansion nickel alloy.

10. A structure as claimed in claim 1, wherein the fibre layers are made of fibres including at least one of carbon, carbon/epoxy, carbon/bismalemide, carbon/cyanate, ester, carbon/polyimide, glass, aramid, high density polyethylene, polypropylene, Zylon^R™ (PBO), ceramic boron.

11. A structure as claimed in claim 10, wherein the fibres incorporate metal.

12. A structure as claimed in any preceding claim, wherein at least one shim plate is located with a minor surface thereof on a leading surface, in use, of said truss structure.

13. A structure as claimed in claim 12, wherein said shim plate is oriented one of transversely and longitudinally of said composite structure.

14. A structure as claimed in any preceding claim, wherein a metallic surface is located one of between said first fibre layer and said truss structure, between said second fibre layer and said truss structure, and on an outer surface of said first and second fibre layers.

15. A structure as claimed in any of claims 1 — 13, wherein a metallic member is located between adjacent box structures with opposing major surfaces of said metallic member adjacent a respective box structure, wherein an interstitial space is provided between said metallic member and each adjacent box structure into which said resin is infused.

16. A structure as claimed in any preceding claim, wherein a fibreglass layer is provided between said first and second fibre layers and said truss structure.

17. A structure as claimed in any of claims 1 - 15, wherein a fϊbreglass layer is provided on an outer surface of at least one of said first and second fibre layers.

18. A structure as claimed in any of claims 1 — 15, wherein a metallised member is located on an outer surface of at least of one of said first and second fibre layers.

19. A structure as claimed in claim 18, wherein said metallised member and/or said metallised surface is one of a titanium sheet, a titanium spray coating, a perforated sheet and a metallised fabric.

20. A structure as claimed in any preceding claim, wherein said composite structure is formed into components of an aircraft.

21. A process of making a fibre metal reinforced composite structure having at least one fibre truss structure and on opposing sides thereof respective first and second fibre layers, including the steps of forming a metallised surface to at least one of an inner and outer surface of said truss structure, infusing resin through the truss structure in a vacuum and curing the resin so as to bond the structure together.

22. A process as claimed in claim 21, wherein the metallised surface is formed by a metal tape which is applied by an automatic tape laying process.