WO2006072758A2 - Fibre metal reinforced composite structure - Google Patents
Fibre metal reinforced composite structure Download PDFInfo
- Publication number
- WO2006072758A2 WO2006072758A2 PCT/GB2005/004818 GB2005004818W WO2006072758A2 WO 2006072758 A2 WO2006072758 A2 WO 2006072758A2 GB 2005004818 W GB2005004818 W GB 2005004818W WO 2006072758 A2 WO2006072758 A2 WO 2006072758A2
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- fibre
- metal
- truss
- layers
- metallised
- Prior art date
Links
- 239000000835 fiber Substances 0.000 title claims abstract description 116
- 239000002131 composite material Substances 0.000 title claims abstract description 63
- 229910052751 metal Inorganic materials 0.000 title claims abstract description 49
- 239000002184 metal Substances 0.000 title claims abstract description 49
- 239000011347 resin Substances 0.000 claims abstract description 51
- 229920005989 resin Polymers 0.000 claims abstract description 51
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims description 44
- 229910052799 carbon Inorganic materials 0.000 claims description 43
- RTAQQCXQSZGOHL-UHFFFAOYSA-N Titanium Chemical compound [Ti] RTAQQCXQSZGOHL-UHFFFAOYSA-N 0.000 claims description 36
- 239000010936 titanium Substances 0.000 claims description 35
- 229910052719 titanium Inorganic materials 0.000 claims description 35
- 238000000034 method Methods 0.000 claims description 19
- 239000004744 fabric Substances 0.000 claims description 14
- 239000011152 fibreglass Substances 0.000 claims description 14
- 239000000463 material Substances 0.000 claims description 13
- 239000004593 Epoxy Substances 0.000 claims description 12
- 230000008569 process Effects 0.000 claims description 12
- 239000004411 aluminium Substances 0.000 claims description 7
- 229910052782 aluminium Inorganic materials 0.000 claims description 7
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 claims description 7
- 239000011888 foil Substances 0.000 claims description 5
- 239000004760 aramid Substances 0.000 claims description 4
- 229920003235 aromatic polyamide Polymers 0.000 claims description 4
- 239000010935 stainless steel Substances 0.000 claims description 4
- 229910001220 stainless steel Inorganic materials 0.000 claims description 4
- 229910000838 Al alloy Inorganic materials 0.000 claims description 3
- ZOXJGFHDIHLPTG-UHFFFAOYSA-N Boron Chemical compound [B] ZOXJGFHDIHLPTG-UHFFFAOYSA-N 0.000 claims description 3
- 229910000990 Ni alloy Inorganic materials 0.000 claims description 3
- 229910052796 boron Inorganic materials 0.000 claims description 3
- 239000011521 glass Substances 0.000 claims description 3
- -1 polypropylene Polymers 0.000 claims description 3
- 238000005507 spraying Methods 0.000 claims description 3
- 239000004642 Polyimide Substances 0.000 claims description 2
- 239000004743 Polypropylene Substances 0.000 claims description 2
- 239000000919 ceramic Substances 0.000 claims description 2
- XLJMAIOERFSOGZ-UHFFFAOYSA-M cyanate Chemical compound [O-]C#N XLJMAIOERFSOGZ-UHFFFAOYSA-M 0.000 claims description 2
- 150000002148 esters Chemical class 0.000 claims description 2
- 229920001903 high density polyethylene Polymers 0.000 claims description 2
- 239000004700 high-density polyethylene Substances 0.000 claims description 2
- 229920002577 polybenzoxazole Polymers 0.000 claims description 2
- 229920001721 polyimide Polymers 0.000 claims description 2
- 229920001155 polypropylene Polymers 0.000 claims description 2
- 239000007788 liquid Substances 0.000 description 21
- 230000008901 benefit Effects 0.000 description 10
- 238000001802 infusion Methods 0.000 description 9
- 238000004519 manufacturing process Methods 0.000 description 9
- 230000007797 corrosion Effects 0.000 description 7
- 238000005260 corrosion Methods 0.000 description 7
- 229920001971 elastomer Polymers 0.000 description 6
- 239000000806 elastomer Substances 0.000 description 6
- 230000003014 reinforcing effect Effects 0.000 description 6
- 238000010586 diagram Methods 0.000 description 5
- 238000012546 transfer Methods 0.000 description 4
- 239000011157 advanced composite material Substances 0.000 description 3
- 238000009833 condensation Methods 0.000 description 3
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- 238000010276 construction Methods 0.000 description 3
- 238000013461 design Methods 0.000 description 3
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- 238000002347 injection Methods 0.000 description 3
- 239000007924 injection Substances 0.000 description 3
- 238000009413 insulation Methods 0.000 description 3
- 239000002905 metal composite material Substances 0.000 description 3
- 238000000429 assembly Methods 0.000 description 2
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- 230000004888 barrier function Effects 0.000 description 2
- 238000007596 consolidation process Methods 0.000 description 2
- 238000005553 drilling Methods 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000003628 erosive effect Effects 0.000 description 2
- 230000004313 glare Effects 0.000 description 2
- 239000003365 glass fiber Substances 0.000 description 2
- 230000035515 penetration Effects 0.000 description 2
- 230000002093 peripheral effect Effects 0.000 description 2
- 230000002787 reinforcement Effects 0.000 description 2
- 238000009745 resin transfer moulding Methods 0.000 description 2
- 239000003381 stabilizer Substances 0.000 description 2
- 229910000906 Bronze Inorganic materials 0.000 description 1
- 229920000049 Carbon (fiber) Polymers 0.000 description 1
- RYGMFSIKBFXOCR-UHFFFAOYSA-N Copper Chemical compound [Cu] RYGMFSIKBFXOCR-UHFFFAOYSA-N 0.000 description 1
- 208000032365 Electromagnetic interference Diseases 0.000 description 1
- 229920000914 Metallic fiber Polymers 0.000 description 1
- 229910000831 Steel Inorganic materials 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 239000010974 bronze Substances 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 239000000805 composite resin Substances 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 239000004020 conductor Substances 0.000 description 1
- 239000010949 copper Substances 0.000 description 1
- 229910052802 copper Inorganic materials 0.000 description 1
- KUNSUQLRTQLHQQ-UHFFFAOYSA-N copper tin Chemical compound [Cu].[Sn] KUNSUQLRTQLHQQ-UHFFFAOYSA-N 0.000 description 1
- 230000002708 enhancing effect Effects 0.000 description 1
- 229920001973 fluoroelastomer Polymers 0.000 description 1
- 239000006260 foam Substances 0.000 description 1
- 229910002804 graphite Inorganic materials 0.000 description 1
- 239000010439 graphite Substances 0.000 description 1
- 238000005470 impregnation Methods 0.000 description 1
- 230000008595 infiltration Effects 0.000 description 1
- 238000001764 infiltration Methods 0.000 description 1
- 239000012212 insulator Substances 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
- 238000002955 isolation Methods 0.000 description 1
- 238000005304 joining Methods 0.000 description 1
- 239000007769 metal material Substances 0.000 description 1
- 150000002739 metals Chemical class 0.000 description 1
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 1
- 239000003973 paint Substances 0.000 description 1
- 230000003071 parasitic effect Effects 0.000 description 1
- 229920001296 polysiloxane Polymers 0.000 description 1
- 230000008439 repair process Effects 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
- 239000010959 steel Substances 0.000 description 1
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Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/06—Fibrous reinforcements only
- B29C70/08—Fibrous reinforcements only comprising combinations of different forms of fibrous reinforcements incorporated in matrix material, forming one or more layers, and with or without non-reinforced layers
- B29C70/088—Fibrous reinforcements only comprising combinations of different forms of fibrous reinforcements incorporated in matrix material, forming one or more layers, and with or without non-reinforced layers and with one or more layers of non-plastics material or non-specified material, e.g. supports
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/06—Fibrous reinforcements only
- B29C70/08—Fibrous reinforcements only comprising combinations of different forms of fibrous reinforcements incorporated in matrix material, forming one or more layers, and with or without non-reinforced layers
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/06—Fibrous reinforcements only
- B29C70/08—Fibrous reinforcements only comprising combinations of different forms of fibrous reinforcements incorporated in matrix material, forming one or more layers, and with or without non-reinforced layers
- B29C70/083—Combinations of continuous fibres or fibrous profiled structures oriented in one direction and reinforcements forming a two dimensional structure, e.g. mats
- B29C70/085—Combinations of continuous fibres or fibrous profiled structures oriented in one direction and reinforcements forming a two dimensional structure, e.g. mats the structure being deformed in a three dimensional configuration
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/40—Shaping or impregnating by compression not applied
- B29C70/42—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles
- B29C70/44—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
- B29C70/443—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding and impregnating by vacuum or injection
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/40—Shaping or impregnating by compression not applied
- B29C70/42—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles
- B29C70/44—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using isostatic pressure, e.g. pressure difference-moulding, vacuum bag-moulding, autoclave-moulding or expanding rubber-moulding
- B29C70/446—Moulding structures having an axis of symmetry or at least one channel, e.g. tubular structures, frames
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/40—Shaping or impregnating by compression not applied
- B29C70/42—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles
- B29C70/46—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs
- B29C70/48—Shaping or impregnating by compression not applied for producing articles of definite length, i.e. discrete articles using matched moulds, e.g. for deforming sheet moulding compounds [SMC] or prepregs and impregnating the reinforcements in the closed mould, e.g. resin transfer moulding [RTM], e.g. by vacuum
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/88—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts characterised primarily by possessing specific properties, e.g. electrically conductive or locally reinforced
- B29C70/882—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts characterised primarily by possessing specific properties, e.g. electrically conductive or locally reinforced partly or totally electrically conductive, e.g. for EMI shielding
- B29C70/885—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts characterised primarily by possessing specific properties, e.g. electrically conductive or locally reinforced partly or totally electrically conductive, e.g. for EMI shielding with incorporated metallic wires, nets, films or plates
-
- C—CHEMISTRY; METALLURGY
- C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
- C22C—ALLOYS
- C22C47/00—Making alloys containing metallic or non-metallic fibres or filaments
- C22C47/20—Making alloys containing metallic or non-metallic fibres or filaments by subjecting to pressure and heat an assembly comprising at least one metal layer or sheet and one layer of fibres or filaments
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3076—Aircrafts
- B29L2031/3085—Wings
Definitions
- 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.
- 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.
- 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.
- said resin is infused between first and second fibre layers and said truss structure, and between adjacent box structures.
- said first and second fibre layers each include at least one layer of fibre material.
- the first and second fibre layers each include two layers of fibre material.
- said fibre material is a carbon fibre sheet.
- 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.
- the metal is one of titanium, aluminium, aluminium alloy, stainless steel, and low thermal expansion nickel alloy.
- 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 TM (PBO), ceramic boron.
- Such fibres may be incorporated with metal.
- At least one shim plate is located with a minor surface thereof on a leading surface, in use, of said truss structure.
- said shim plate is oriented one of transversely and longitudinally of said composite structure.
- 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.
- 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.
- a fibreglass layer is provided between said first and second fibre layers and said truss structure.
- a fibreglass layer is provided on an outer surface of at least one of said first and second fibre layers.
- a metallised member is located on an outer surface of at least of one of said first and second fibre layers.
- 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.
- said composite structure is formed into components of an aircraft.
- 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.
- 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:
- 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.
- 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).
- 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.
- Titanium will fail at a much higher temperature than the fibre layers.
- 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.
- 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.
- a composite structure as defined above formed into components of an aircraft, such as a fuselage, floor panels, leading edge structures, and stabiliser skins.
- 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.
- the metallised surface is formed by a metal tape which is applied by an automatic tape laying process.
- 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.
- Figure 12 shows a 100 x enlargement of a fabric useful in this invention.
- FIG. 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.
- FIG. 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").
- a titanium shim 7 There is additionally provided a titanium shim 7.
- FIG 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.
- 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.
- 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.
- 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.
- 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.
- 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.
- 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.
- 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.
- Figure 6 shows shim plates 9 on the inside of the upper surface of the carbon braided box structures 3.
- FIGS 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).
- FIGs 8a and 8b show a composite assembly incorporating fibreglass layers.
- 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.
- a fibreglass layer 13 on the outer face of the upper skin formed by carbon fibre layers 1 and 2.
- 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.
- this drape limitation can be overcome by:
- titanium is initially placed applied to the mould either by a spray coating or pre-stretch forming to shape.
- 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:
- 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
- the fibreglass can also increase the insulation properties as, for example, shown in Figures 8a and 8b.
- 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.
- Titanium coated carbon fibres woven or braided into a fabric.
- 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.
- 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.
- 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.
- 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.
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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, ZylonR™ (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
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, ZylonR™ (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.
Applications Claiming Priority (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB0500408.0 | 2005-01-10 | ||
GB0500408A GB0500408D0 (en) | 2005-01-10 | 2005-01-10 | Metal fibre reinforced composite |
GB0516814A GB0516814D0 (en) | 2005-08-16 | 2005-08-16 | Fibre metal reinforced composite structure |
GB0516814.1 | 2005-08-16 |
Publications (2)
Publication Number | Publication Date |
---|---|
WO2006072758A2 true WO2006072758A2 (en) | 2006-07-13 |
WO2006072758A3 WO2006072758A3 (en) | 2006-12-21 |
Family
ID=35736111
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/GB2005/004818 WO2006072758A2 (en) | 2005-01-10 | 2005-12-14 | Fibre metal reinforced composite structure |
Country Status (2)
Country | Link |
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GB (1) | GB2421926B (en) |
WO (1) | WO2006072758A2 (en) |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP3056337A1 (en) * | 2015-02-13 | 2016-08-17 | Advanced Materials Technology Innovation Company | Composite laminate with reinforcement of metal mesh |
EP2046570B1 (en) * | 2006-08-01 | 2018-09-05 | Airbus Operations GmbH | Lay-up arrangement and structural arrangement and method for producing a component for aerospace |
CN112918687A (en) * | 2021-01-29 | 2021-06-08 | 东华大学 | Integral sealed cabin and preparation method thereof |
JP2023502142A (en) * | 2020-01-02 | 2023-01-20 | 中▲車▼青▲島▼四方▲機車車▼輌股▲分▼有限公司 | Composite car body, composite paving structure and paving method |
Families Citing this family (7)
Publication number | Priority date | Publication date | Assignee | Title |
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DE102006007428A1 (en) * | 2006-02-17 | 2007-08-30 | Airbus Deutschland Gmbh | Reinforcing material for locally reinforcing a component formed with a composite material and method |
CA2739477C (en) | 2008-10-03 | 2016-12-20 | Short Brothers Plc | Fibre reinforced composite structures and method of manufacture |
JP5959558B2 (en) * | 2014-03-13 | 2016-08-02 | アイシン高丘株式会社 | Composite structure and method for producing the same |
CN105882059A (en) * | 2015-01-26 | 2016-08-24 | 三鼎材料科技股份有限公司 | Structure for reinforcing composite with metal meshes |
US10913210B2 (en) | 2017-02-23 | 2021-02-09 | Raytheon Technologies Corporation | Mold-less curing method of manufacturing a composite vessel assembly |
EP3595853A4 (en) | 2017-03-16 | 2020-12-23 | Guerrilla Industries LLC | Composite structures and methods of forming composite structures |
WO2022192355A1 (en) * | 2021-03-09 | 2022-09-15 | Guerrilla Industries LLC | Composite structures and methods of forming composite structures |
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GB635823A (en) * | 1945-06-01 | 1950-04-19 | Ernest Platton King | Improved method of manufacturing composite metal-fibrous structures |
US4617217A (en) * | 1983-09-19 | 1986-10-14 | Society Nationale Industrielle Aerospatiale | Beam or other element of great length of a composite material polymerized under heat and pressure |
US4934412A (en) * | 1987-03-10 | 1990-06-19 | Fischer Advanced Composite Components Gesellschaft M.B.H. | Cooling pipe line for an engine |
US5925277A (en) * | 1996-07-29 | 1999-07-20 | The Boeing Company | Annealed thermoplastic weld |
WO2002004201A1 (en) * | 2000-07-07 | 2002-01-17 | Delphi Technologies, Inc. | Cored contoured composite structural members and methods for making the same |
WO2003103933A1 (en) * | 2002-06-07 | 2003-12-18 | Short Brothers Plc | A fibre reinforced composite component and method to produce such component |
GB2410921A (en) * | 2004-02-10 | 2005-08-17 | Eurocopter Deutschland | Method for the manufacture of a hollow fibre-composite component and intermediate product therefor |
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US3091262A (en) * | 1957-08-19 | 1963-05-28 | Darworth Inc | Metal-fiber reinforced resin laminate |
GB8921076D0 (en) * | 1989-09-18 | 1989-11-01 | Shell Int Research | Fibre reinforced plastic grid |
US8262823B2 (en) * | 2003-09-04 | 2012-09-11 | The Boeing Company | Window skin panel and method of making same |
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2005
- 2005-12-14 GB GB0525447A patent/GB2421926B/en not_active Expired - Fee Related
- 2005-12-14 WO PCT/GB2005/004818 patent/WO2006072758A2/en not_active Application Discontinuation
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB635823A (en) * | 1945-06-01 | 1950-04-19 | Ernest Platton King | Improved method of manufacturing composite metal-fibrous structures |
US4617217A (en) * | 1983-09-19 | 1986-10-14 | Society Nationale Industrielle Aerospatiale | Beam or other element of great length of a composite material polymerized under heat and pressure |
US4934412A (en) * | 1987-03-10 | 1990-06-19 | Fischer Advanced Composite Components Gesellschaft M.B.H. | Cooling pipe line for an engine |
US5925277A (en) * | 1996-07-29 | 1999-07-20 | The Boeing Company | Annealed thermoplastic weld |
WO2002004201A1 (en) * | 2000-07-07 | 2002-01-17 | Delphi Technologies, Inc. | Cored contoured composite structural members and methods for making the same |
WO2003103933A1 (en) * | 2002-06-07 | 2003-12-18 | Short Brothers Plc | A fibre reinforced composite component and method to produce such component |
GB2410921A (en) * | 2004-02-10 | 2005-08-17 | Eurocopter Deutschland | Method for the manufacture of a hollow fibre-composite component and intermediate product therefor |
Cited By (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP2046570B1 (en) * | 2006-08-01 | 2018-09-05 | Airbus Operations GmbH | Lay-up arrangement and structural arrangement and method for producing a component for aerospace |
EP3056337A1 (en) * | 2015-02-13 | 2016-08-17 | Advanced Materials Technology Innovation Company | Composite laminate with reinforcement of metal mesh |
JP2023502142A (en) * | 2020-01-02 | 2023-01-20 | 中▲車▼青▲島▼四方▲機車車▼輌股▲分▼有限公司 | Composite car body, composite paving structure and paving method |
US11780507B2 (en) | 2020-01-02 | 2023-10-10 | Crrc Qingdao Sifang Co., Ltd. | Composite material vehicle body and composite material laying structure and laying method |
CN112918687A (en) * | 2021-01-29 | 2021-06-08 | 东华大学 | Integral sealed cabin and preparation method thereof |
Also Published As
Publication number | Publication date |
---|---|
GB2421926A (en) | 2006-07-12 |
GB2421926B (en) | 2010-03-10 |
WO2006072758A3 (en) | 2006-12-21 |
GB0525447D0 (en) | 2006-01-25 |
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