METHOD FOR CONNECTING A FIBRE COMPOSITE COMPONENT TO A STRUCTURAL COMPONENT OF AN AIRCRAFT AND SPACECRAFT AND A CORRESPONDING ARRANGEMENT

• US 20130149501A1
• Filed: 06/01/2012
• Published: 06/13/2013
• Est. Priority Date: 12/08/2009
• Status: Active Application

• Current Assignee: Airbus Operations GmbH (Airbus SE)
• Original Assignee: Airbus Operations GmbH (Airbus SE)
• Inventors: PACCHIONE, Marco, FURFARI, Domenico
• Application Number:
  US13/486,444
• Publication Number:
  US 20130149501A1
• Time in Patent Office: Days
• US Class Current: 428/172
• CPC Class Codes: B23P 13/00 Making metal objects by operations essentially involving machining but not covered by a single other subclass making specific objects B23P15/00
• CPC Class Codes: B29C 35/02 Heating or curing, e.g. crosslinking or vulcanizing during moulding, e.g. in a mouldcold vulcanisation B29C35/18 ; vulcanising tyres, presses therefor B29D30/0601
• CPC Class Codes: B29C 65/344 being a woven or non-woven fabric or being a mesh
• CPC Class Codes: B29C 65/3444 being a ribbon, band or strip
• CPC Class Codes: B29C 65/3448 said ribbon, band or strip being perforated
• CPC Class Codes: B29C 65/3476 being metallic
• CPC Class Codes: B29C 65/364 being a woven or non-woven fabric or being a mesh
• CPC Class Codes: B29C 65/3644 being a ribbon, band or strip
• CPC Class Codes: B29C 65/3656 being a layer of a multilayer part to be joined, e.g. for joining plastic-metal laminates
• CPC Class Codes: B29C 65/368 with a polymer coating
• CPC Class Codes: B29C 65/483 Reactive adhesives, e.g. chemically curing adhesives
• CPC Class Codes: B29C 65/5021 being multi-layered
• CPC Class Codes: B29C 65/5057 positioned between the surfaces to be joined B29C65/5035 takes precedence
• CPC Class Codes: B29C 65/564 hidden in the joint, e.g. dowels or Z-pins B29C65/603 takes precedence
• CPC Class Codes: B29C 66/1122 Single lap to lap joints, i.e. overlap joints B29C66/45, B29C66/472, B29C66/52272 take precedence
• CPC Class Codes: B29C 66/30341 non-integral with the parts to be joined, e.g. making use of extra elements B29C65/562 takes precedence
• CPC Class Codes: B29C 66/43 Joining a relatively small portion of the surface of said articles B29C66/45 takes precedence
• CPC Class Codes: B29C 66/721 Fibre-reinforced materials B29C66/729 takes precedence
• CPC Class Codes: B29C 66/7212 characterised by the composition of the fibres
• CPC Class Codes: B29C 66/72141 Fibres of continuous length
• CPC Class Codes: B29C 66/72143 : Fibres of discontinuous lengths
• CPC Class Codes: B29C 66/73751 : the to-be-joined area of at least one of the parts to be joined being uncured, i.e. non cross-linked, non vulcanized
• CPC Class Codes: B29C 66/73752 : the to-be-joined areas of both parts to be joined being uncured
• CPC Class Codes: B29C 66/73753 : the to-be-joined area of at least one of the parts to be joined being partially cured, i.e. partially cross-linked, partially vulcanized
• CPC Class Codes: B29C 66/73754 : the to-be-joined areas of both parts to be joined being partially cured
• CPC Class Codes: B29C 66/73921 : characterised by the materials of both parts being thermoplastics
• CPC Class Codes: B29C 66/742 : to metals or their alloys
• CPC Class Codes: B29K 2101/12 : Thermoplastic materials
• CPC Class Codes: B29K 2105/06 : containing reinforcements, fillers or inserts
• CPC Class Codes: B29K 2105/24 : crosslinked or vulcanised
• CPC Class Codes: B29K 2105/243 : Partially cured
• CPC Class Codes: B29K 2105/246 : Uncured, e.g. green
• CPC Class Codes: B29K 2305/00 : Use of metals, their alloys or their compounds, as reinforcement
• CPC Class Codes: B29K 2307/00 : Use of elements other than metals as reinforcement
• CPC Class Codes: B29K 2307/04 : Carbon
• CPC Class Codes: B29L 2031/3076 : Aircrafts blades, propellers B29L2031/08
• CPC Class Codes: B32B 7/08 : by mechanical means
• CPC Class Codes: B64C 2001/0072 : from composite materials
• CPC Class Codes: F16B 15/0023 : Nail plates
• CPC Class Codes: F16B 15/0046 : from the body of the plate
• CPC Class Codes: Y02T 50/40 : Weight reduction
• CPC Class Codes: Y10T 29/49622 : Vehicular structural member making
• CPC Class Codes: Y10T 428/24612 : Composite web or sheet

Abstract

In the case of a method for connecting a fibre composite component to a structural component of an aircraft and spacecraft, a metal foil is provided as a transverse reinforcement element between the fibre composite component and the structural component. It is formed with at least one anchoring portion which protrudes from the surface facing the fibre composite component and is inserted between the fibre composite component and the structural component. A corresponding arrangement is produced in accordance with this method.

Claims

1. A method for connecting a fibre composite component to a structural component of an aircraft and spacecraft, comprising the steps of:

providing a metal foil as a transverse reinforcement element between the fibre composite component and the structural component;

forming the metal foil with at least one anchoring portion which protrudes from the surface facing the fibre composite component (2); and

inserting the metal foil between the fibre composite component and the structural component.

2. The method of claim 1, wherein the metal foil is connected to the fibre composite component by curing of same, if the fibre composite component is uncured or is partially cured, wherein the at least one anchoring portion extends into the fibre composite component.

3. The method of claim 1, wherein the metal foil is connected to the fibre composite component by means of a first adhesive layer therewith, if the fibre composite component is cured, wherein the at least one anchoring portion extends into the first adhesive layer.

4. The method of claim 1, wherein the metal foil is connected to the structural component by means of a second adhesive layer, if the structural component is a metallic part or has a cured fibre composite structure.

5. The method of claim 4, wherein the metal foil is connected to the structural component by means of a further anchoring portion which protrudes from the surface facing the structural component and extends into the second adhesive layer.

6. The method of claim 1, wherein the method comprises the following additional method steps, if the structural component has an uncured or a partially cured fibre composite structure:

forming the metal foil with a further anchoring portion which protrudes from the surface facing the structural component; and

connecting the further anchoring portion to the structural component by curing of same, wherein the further anchoring portion extends into the structural component.

7. The method of claim 1, wherein the method comprises the following additional method steps, if the fibre composite component and the structural component each comprise a thermoplastic laminate:

forming the metal foil with at least one further anchoring portion which protrudes from the surface facing the structural component; and

welding the fibre composite component to the structural component after insertion of the metal foil, wherein the anchoring portions extend into the respectively associated component.

8. The method of claim 7, wherein during the welding procedure the metal foil is used at least partially for the purpose of coupling heat in, e.g. by induction.

9. The method of claim 1, wherein the metal foil is formed with at least one anchoring portion by means of at least one of the following:

punch-bending methods, high-speed metal removal, electron beam processing, additive layer manufacturing methods or welding of anchoring elements or anchoring pins or anchoring elements and anchoring pins.

10. A method for connecting a fibre composite component to a structural component of an aircraft and spacecraft, comprising the steps of:

providing a metal foil as a transverse reinforcement element in a connection portion of the fibre composite component and the structural component;

forming the metal foil with at least one anchoring portion which protrudes from a surface of the metal foil; and

disposing the metal foil on an outer side of the connection portion, wherein the at least one anchoring portion extends completely through the fibre composite component and extends into the structural component or extends completely through the structural component and extends into the fibre composite component.

11. The method of claim 10, wherein the metal foil is formed with at least one anchoring portion by means of at least one of the following:

punch-bending methods, high-speed metal removal, electron beam processing, additive layer manufacturing methods, welding of anchoring elements or anchoring pins or anchoring elements and anchoring pins.

12. An arrangement comprising a fibre composite component and a structural component of an aircraft and spacecraft, in which a metal foil is inserted in a connection portion of the fibre composite component and structural component as a transverse reinforcement element, wherein the metal foil comprises at least one anchoring portion which protrudes from a surface of the metal foil.

13. The arrangement of claim 12, wherein the metal foil comprises a further anchoring portion which protrudes from the surface of the metal foil facing the structural component.

14. The arrangement of claim 12, wherein the at least one anchoring portion or the further anchoring portion or at least one anchoring portion and the further anchoring portion are formed as punch-bent parts and comprise anchoring elements which are disposed substantially perpendicularly or at a predetermined angle with respect to the respective surface of the metal foil.

15. The arrangement of claim 12, wherein the at least one anchoring portion or further anchoring portion or the at least one anchoring portion and further anchoring portion comprise anchoring pins which are produced by means of electron beam processing or additive layer manufacturing methods or electron beam processing and additive layer manufacturing methods.

16. The arrangement of claim 12, wherein the at least one anchoring portion or further anchoring portion or the at least one anchoring portion and further anchoring portion comprise anchoring pins which are separately produced and welded.

17. The arrangement of claim 12, wherein the anchoring elements are provided with barbs or serrations or barbs and serrations.