WO2010144007A1 - Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler - Google Patents

Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler Download PDF

Info

Publication number
WO2010144007A1
WO2010144007A1 PCT/SE2009/050716 SE2009050716W WO2010144007A1 WO 2010144007 A1 WO2010144007 A1 WO 2010144007A1 SE 2009050716 W SE2009050716 W SE 2009050716W WO 2010144007 A1 WO2010144007 A1 WO 2010144007A1
Authority
WO
WIPO (PCT)
Prior art keywords
radius filler
composite
structural
nanostructure
parts
Prior art date
Application number
PCT/SE2009/050716
Other languages
French (fr)
Inventor
Per Hallander
Mikael Petersson
Björn Weidmann
Tommy Grankäll
Göte Strindberg
Pontus Nordin
Original Assignee
Saab Ab
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Saab Ab filed Critical Saab Ab
Priority to US13/377,054 priority Critical patent/US20120074265A1/en
Priority to CA2765138A priority patent/CA2765138A1/en
Priority to PCT/SE2009/050716 priority patent/WO2010144007A1/en
Priority to BRPI0924562A priority patent/BRPI0924562A2/en
Priority to EP09845906.8A priority patent/EP2440392B1/en
Publication of WO2010144007A1 publication Critical patent/WO2010144007A1/en

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
    • 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
    • B29C65/00—Joining or sealing of preformed parts, e.g. welding of plastics materials; Apparatus therefor
    • B29C65/02—Joining or sealing of preformed parts, e.g. welding of plastics materials; Apparatus therefor by heating, with or without pressure
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/01—General aspects dealing with the joint area or with the area to be joined
    • B29C66/05—Particular design of joint configurations
    • B29C66/10—Particular design of joint configurations particular design of the joint cross-sections
    • B29C66/11—Joint cross-sections comprising a single joint-segment, i.e. one of the parts to be joined comprising a single joint-segment in the joint cross-section
    • B29C66/112—Single lapped joints
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/01—General aspects dealing with the joint area or with the area to be joined
    • B29C66/05—Particular design of joint configurations
    • B29C66/10—Particular design of joint configurations particular design of the joint cross-sections
    • B29C66/11—Joint cross-sections comprising a single joint-segment, i.e. one of the parts to be joined comprising a single joint-segment in the joint cross-section
    • B29C66/112—Single lapped joints
    • B29C66/1122—Single lap to lap joints, i.e. overlap joints
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/01—General aspects dealing with the joint area or with the area to be joined
    • B29C66/05—Particular design of joint configurations
    • B29C66/10—Particular design of joint configurations particular design of the joint cross-sections
    • B29C66/13—Single flanged joints; Fin-type joints; Single hem joints; Edge joints; Interpenetrating fingered joints; Other specific particular designs of joint cross-sections not provided for in groups B29C66/11 - B29C66/12
    • B29C66/131—Single flanged joints, i.e. one of the parts to be joined being rigid and flanged in the joint area
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/40—General aspects of joining substantially flat articles, e.g. plates, sheets or web-like materials; Making flat seams in tubular or hollow articles; Joining single elements to substantially flat surfaces
    • B29C66/41—Joining substantially flat articles ; Making flat seams in tubular or hollow articles
    • B29C66/43—Joining a relatively small portion of the surface of said articles
    • B29C66/434—Joining substantially flat articles for forming corner connections, fork connections or cross connections
    • B29C66/4344—Joining substantially flat articles for forming fork connections, e.g. for making Y-shaped pieces
    • B29C66/43441—Joining substantially flat articles for forming fork connections, e.g. for making Y-shaped pieces with two right angles, e.g. for making T-shaped pieces, H-shaped pieces
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/40—General aspects of joining substantially flat articles, e.g. plates, sheets or web-like materials; Making flat seams in tubular or hollow articles; Joining single elements to substantially flat surfaces
    • B29C66/49—Internally supporting the, e.g. tubular, article during joining
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/40—General aspects of joining substantially flat articles, e.g. plates, sheets or web-like materials; Making flat seams in tubular or hollow articles; Joining single elements to substantially flat surfaces
    • B29C66/49—Internally supporting the, e.g. tubular, article during joining
    • B29C66/496—Internally supporting the, e.g. tubular, article during joining using a support which remains in the joined object
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/70—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material
    • B29C66/72—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the structure of the material of the parts to be joined
    • B29C66/721—Fibre-reinforced materials
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/70—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material
    • B29C66/72—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the structure of the material of the parts to be joined
    • B29C66/721—Fibre-reinforced materials
    • B29C66/7214—Fibre-reinforced materials characterised by the length of the fibres
    • B29C66/72141—Fibres of continuous length
    • 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
    • B29C70/10—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres
    • B29C70/12—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of short length, e.g. in the form of a mat
    • B29C70/14—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of short length, e.g. in the form of a mat oriented
    • 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
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29D—PRODUCING PARTICULAR ARTICLES FROM PLASTICS OR FROM SUBSTANCES IN A PLASTIC STATE
    • B29D99/00—Subject matter not provided for in other groups of this subclass
    • B29D99/0003—Producing profiled members, e.g. beams
    • B29D99/0005—Producing noodles, i.e. composite gap fillers, characterised by their construction
    • 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
    • B29C35/00—Heating, cooling or curing, e.g. crosslinking or vulcanising; Apparatus therefor
    • B29C35/02—Heating or curing, e.g. crosslinking or vulcanizing during moulding, e.g. in a mould
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/70—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material
    • B29C66/71—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the composition of the plastics material of the parts to be joined
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/70—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material
    • B29C66/72—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the structure of the material of the parts to be joined
    • B29C66/721—Fibre-reinforced materials
    • B29C66/7212—Fibre-reinforced materials characterised by the composition of the fibres
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/70—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material
    • B29C66/73—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset
    • B29C66/737—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the state of the material of the parts to be joined
    • B29C66/7375—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the state of the material of the parts to be joined uncured, partially cured or fully cured
    • B29C66/73753—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the state of the material of the parts to be joined uncured, partially cured or fully cured 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
    • B29C66/73754—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the state of the material of the parts to be joined uncured, partially cured or fully cured 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 the to-be-joined areas of both parts to be joined being partially cured
    • 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
    • B29C66/00—General aspects of processes or apparatus for joining preformed parts
    • B29C66/70—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material
    • B29C66/73—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset
    • B29C66/739—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the material of the parts to be joined being a thermoplastic or a thermoset
    • B29C66/7394—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the material of at least one of the parts being a thermoset
    • B29C66/73941—General aspects of processes or apparatus for joining preformed parts characterised by the composition, physical properties or the structure of the material of the parts to be joined; Joining with non-plastics material characterised by the intensive physical properties of the material of the parts to be joined, by the optical properties of the material of the parts to be joined, by the extensive physical properties of the parts to be joined, by the state of the material of the parts to be joined or by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the material of the parts to be joined being a thermoplastic or a thermoset characterised by the material of at least one of the parts being a thermoset characterised by the materials of both parts being thermosets
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
    • B29K2105/00—Condition, form or state of moulded material or of the material to be shaped
    • B29K2105/06—Condition, form or state of moulded material or of the material to be shaped containing reinforcements, fillers or inserts
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
    • B29K2105/00—Condition, form or state of moulded material or of the material to be shaped
    • B29K2105/06—Condition, form or state of moulded material or of the material to be shaped containing reinforcements, fillers or inserts
    • B29K2105/16—Fillers
    • B29K2105/165—Hollow fillers, e.g. microballoons or expanded particles
    • B29K2105/167—Nanotubes
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
    • B29K2105/00—Condition, form or state of moulded material or of the material to be shaped
    • B29K2105/24—Condition, form or state of moulded material or of the material to be shaped crosslinked or vulcanised
    • B29K2105/243—Partially cured
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
    • B29K2307/00—Use of elements other than metals as reinforcement
    • 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
    • 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
    • B29L2031/3082—Fuselages
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/40—Weight reduction
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/24—Structurally defined web or sheet [e.g., overall dimension, etc.]
    • Y10T428/24479—Structurally defined web or sheet [e.g., overall dimension, etc.] including variation in thickness

Definitions

  • Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler
  • the present invention relates to a composite radius filler according to the preamble of cla im 1 and a method of producing an aircraft structure according to claim 10.
  • structural composite parts such as semi-cured and pre-formed plastic beams, resin shells etc.
  • the gap is formed by at least two rounded sections of rounded or curved areas of the respective structural composite part facing each other.
  • the composite radius fillers are thus arranged between the rounded sections of the structural composite parts (such as composite laminates) for filling the gap between structural composite parts' rounded areas.
  • the radius filler is thus applied in the aircraft structure for preventing deformation and/or weakening of the area near the gap or cavity otherwise present between the structural composite parts of the aircraft structure.
  • the radius filler is often used to fill the gap between outer rounded edges of adjacent structural composite parts being bonded to each other.
  • the structural composite parts will thus, bonded together, constitute a rigid and strong aircraft structure for use in an aircraft, wherein the structural composite parts must be of a satisfactory strength.
  • the aircraft structure is in this application defined as a specific structure of an aircraft, such as a wing, a fuselage, a rudder, a flap, an aileron, a fin, a tailplane etc.
  • the aircraft structure consists of at least two assembled, and bond together, three-dimensional structural composite parts.
  • the structural composite parts are formed (e.g. hot drape forming or mechanical forming) earlier into three-dimensional structural composite parts and are thereafter assembled together to form the aircraft structure.
  • the three-dimensional structural composite parts are bonded to each other by means of a bonding interlayer material, i.e. an adhesive.
  • This adhesive is today used for holding together the aircraft structure, i.e. the assembly of the structural composite parts.
  • the adhesive can be a melt-bondable adhesive resin, such as an epoxy.
  • the applicant of this application also has filed an application reciting a bonding interlayer material comprising a nanostructure provided for connecting at least two structural composite parts (e.g. a wing beam and wing shell) together for building an aircraft structure (e.g. a wing) of an aircraft. Earlier such beam and shell connections were made by pure resin, screws, rivets etc.
  • Each structural composite part can be comprised of fibre reinforced plies being laid onto each other before for the forming of a stack. The stack is then formed into said structural composite part and semi-cured for effective handling.
  • Composite radius fillers of today which being comprised of traditional resin filler material without any fibres, serve for filling up the gap between the structural composite parts assembled together to form the aircraft structure.
  • Such a composite radius filler weakens the finished aircraft structure, especially within the interface area between the composite radius filler and the surface of the structural composite parts.
  • the composite radius filler according to prior art thus may tear apart and crack within said interface area due to tear and shear strain. Also, the radius filler may separate from the structural composite part due to heavy loads or stress.
  • the structural composite part of the aircraft structure is defined in this application as a specific three-dimensional structural composite part being used together with at least another specific three-dimensional structural composite part for building the aircraft structure.
  • the three-dimensional structural composite part can be a beam, a shell, a wing rib, a bulkhead, a nose cone shell, a frame, a web stiffener etc.
  • the structural composite part can be made of a stack of pre-preg plies (fiber layers impregnated with resin before being placed on a temporary support by means of for example an Automatic Tape Laying-machine).
  • the stack has plies with different fiber directions.
  • the stack is thereafter moved to a forming tool for forming the stack into a three-dimensional shape, or at least a two- dimensional shape.
  • a force generated from a forming medium e.g. vacuum bag or rollers
  • the benefit with the gliding effect or sliding between the plies is several, especially the lack of wrinkles.
  • the finished formed structural composite part is thereafter moved to an assembly and curing tool for assembly and curing together with at least another finished formed three-dimensional structural composite part with the radius fillers arranged in specific gaps or cavities between adjacent structural composite parts.
  • US 4 331 495 describes structural composite parts being joined, wherein the gaps created by the rounded surfaces are filled with radius fillers, by sewing the structural composite parts together within the area of said rounded areas and the radius filler for preventing "peel” forces tending to separate the structural composite parts at their interface.
  • US 5 650 229 describes a preformed radius filler which is adapted to fill a gap between composite laminates of a structural composite part.
  • the radius filler comprises unidirectional fibers and a stabilizer as injected resin.
  • US 5 827 383 describes a radius filler comprising resin and fibers, wherein the radius filler is further strengthened by means of pins.
  • radius fillers suffer from that the peripheral area of the radius filler, which area is adjacent the interface between the radius filler and the structural composite part rounded surface, has a tendency to crack when the structural composite part is under heavy load and/or stress.
  • an integrated monolithic structure such as an aircraft structure, being comprised of at least three assembled structural composite parts and at least one composite radius filler filling a gap between said at least three assembled structural composite parts.
  • Th is has been achieved by the composite radius filler defined in the introduction being characterized by the features of the characterizing part of claim 1.
  • the aircraft composite integrated structure as a whole can have an optimal strength since the composite radius filler by means of the present invention will be considered as a structural part of the aircraft composite integrated structure.
  • the reinforced composite radius filler will thereby provide the aircraft composite structure a strength in x-, y-, z-direction.
  • the composite radius filler will have a reduced thermal linear expansion and the viscosity can be controlled by the amount of nanostructure for an optimal handling of the composite radius filler under production .
  • the aircraft composite integrated structure (aircraft structure) as a whole will thus have an optimal impact strength.
  • the nanostructure primarily is arranged in the periphery of the composite radius filler for reinforcement of an interface area between the composite radius filler and the structural composite parts.
  • the aircraft structure can be reinforced by means of a small addition of material, meaning that the overall weight of the aircraft can be reduced.
  • the nanostructure located within the peripheral area of the radius filler hinders that cracks will appear in the radius filler near the interface between the radius filler and the structural composite part during heavy stressing of the aircraft.
  • the nanostructure comprises nanofibers and/or nanotubes and/or nanowires.
  • the nanostructure is oriented perpendicular against the structural composite part.
  • the nanostructure has an unidirectional orientation.
  • the strength is increased in a specific direction, e.g. corresponding with the prolongation of the radius filler.
  • the composite radius filler has a triangular cross section.
  • the radius filler can be used in a gap often appearing between three structural composite parts being joined together.
  • the composite radius filler is comprised in an aileron.
  • the step of providing the composite radius filler is characterized by arranging a nanostructure primarily in the periphery of the composite radius filler for reinforcement of the interface between the composite radius filler and the structural composite parts.
  • the viscosity of the resin of the radius filler can be controlled by the amount of nanostructure for an optimal handling of the composite radius filler under production. In this way a cost-effective production is achieved at the same time being flexible regarding the actual application.
  • FIG. 1 illustrates a cross section of a radius filler according to prior art being subjected to stress
  • FIG. 2 illustrates a curing tool for the assembly of structural composite parts formed to an aircraft structure comprising radius fillers according to a first embodiment
  • FIG. 3 illustrates an assembly of structural composite parts wherein the radius filler is positioned in a gap formed between said structural composite parts
  • FIG. 4 illustrates a cross section of a radius filler according to the first embodiment
  • FIG. 5 illustrates a perspective view of a radius filler according to a second embodiment
  • FIG. 6 illustrates a cross section of a radius filler according to a third embodiment.
  • FIG. 1 schematically illustrates a cross section of a radius filler 1 of an aircraft structure 2 according to prior art which has been subjected to stress.
  • a triangular gap 4 is filled according to the prior art with the traditional radius filler 1 being comprised of dry fibres in the form of carbon fibres (not shown) for reinforcing the connection of three surfaces 5', 5", 5'" of respective structural composite part T ', 7", 7'".
  • a first rounded corner outer surface 5' of a first semi-cured composite beam 11 ' being connected to a second rounded corner outer surface 5" of a second semi-cured beam 11 " and to an essentially flat inside surface 5'" of a semi-cured shell 11 '".
  • the aircraft structure 2 is after assembly cured and forms an integrated structure.
  • FIG. 2 schematically illustrates a curing tool 19 for the assembly of structural composite parts 7 formed to an aircraft structure 2, here an aileron , comprising radius fillers 3 according to a first embodiment of the present invention.
  • the radius filler 3 is in detail explained in connection with the FIG. 4 below.
  • the semi-cured structural composite parts 7 are assembled, wherein the radius fillers 3 according to the first embodiment are placed in gaps 4 formed between the rounded surfaces 5 of the structural composite parts T ', 7", and positioned in the curing tool 19.
  • Two gaps 4' of the triangular gaps have each three rounded surfaces (gaps 4, 4' are illustrated in the FIGS. 2 and 3 with broken reference arrow lines and are marked for the sake of clarity as they are filled with the radius fillers 3, wherein the definition of a gap can thus be a hollow space filled with a radius filler).
  • the curing tool 19 thus being loaded with structural composite parts 7, each being earlier formed over by a so called ATL-machine (not shown).
  • the curing tool 19 being loaded with the structural composite parts 7 for building the aircraft structure 2.
  • the structural composite parts 7 being assembled are: a nose cap of reinforced resin being bonded to an upper and lower shell inner surface, a structural nose beam of composite being arranged and bonded to the web of an adjacent first structural U-beam 21 , the flanges of which being bonded to the inner surface of the shell 11 '" and bonded to the flange edges of a second structural U-beam, a third structural U-beam having its web bonded to the web of the second structural U-beam, etc.
  • the upper 11 '" and lower shells are bonded in the rear part (not shown) of the aileron.
  • a radius filler 3 can also thus be regarded as a structural composite part, one of which is in more detail shown in Fig. 3.
  • the structural composite parts 7 and the radius fillers 3 plus a bonding interlayer material 18 (see FIG. 3) adapted for connecting the structural composite parts 7 and the radius fillers 3 together -the bonding interlayer material 18 preferably comprises a curable resin and reinforcing nanostructure (such as nanofibres)-, are positioned in the curing tool 19.
  • the curing tool 19 consisting of an upper 20' and lower 20' ' forming tool part including heating elements (not shown) for increasing the temperature of the structural composite parts 7, the radius fillers 3 and the bonding interlayer material 18 for a proper curing of the bonding interlayer material comprising the nanotubes and a proper curing and bonding of the semi-cured structural composite parts 7 to the radius fillers 3.
  • Interior holding-on tools 24 are placed within the nose beam and the U-beams 21 for achieving a pressure from inside. Each interior holding-on tool can be divided into parts by releasing a wedge 26 arranged for keeping the parts together.
  • FIG. 3 is shown schematically the assembly of the structural composite parts 7 regarding one of the radius 3 fillers in FIG. 2, wherein the radius filler 3 is positioned in the radius gap 4 formed by upper rounded corners of two U-beams 21 , the webs of which is closely adhered to each other, and by the underside of the composite shell 11 '".
  • the radius filler 3 is made of a resin filled with reinforcing nanofibres.
  • the U-beams 21 are firstly assembled, the radius filler 3 is thereafter positioned in a v-formed groove (later on constituting walls of the gap 4) formed by said upper rounded corners.
  • the shell 11 '" is positioned over the U-beams' 21 flanges and the radius filler 3.
  • the structural composite parts 7 are cured for forming the aircraft structure 2.
  • the aircraft structure 2 comprises structural composite parts 7 assembled together to form said aircraft structure 2, wherein the composite radius filler 3 is arranged between rounded sections of at least two of the structural composite parts 7 for filling the gap 4 between the structural composite parts 7.
  • the composite radius filler 3 is made structural by a nanostructure 23 arranged within the composite radius filler 3.
  • the method having the steps of providing said nanostructure reinforced composite radius filler 3 by mixing a curable resin and a nanostructure 23, forming separately at least three structural composite parts 7 from pre-fabhcated pre- impregnated fibre plies, laid-up to each other and having different fibre orientations, assembling the separately formed structural composite parts 7 with said nanostructure reinforced composite radius filler 3 between rounded sections 9 of the structural composite parts 7 for filling the gap 4 between the structural composite parts 7, curing the assembled structural composite parts 7 and the nanostructure reinforced composite radius filler 3 together in a curing tool 19, and removing the finished cured aircraft composite integrated structure (aircraft structure 2) from the curing tool 19.
  • the radius filler 3 is schematically illustrated in a cross section according to the first embodiment.
  • the composite radius filler is thus made structural by the nanostructure 23 arranged within the composite radius filler 3.
  • the nanostructure primarily is arranged in the periphery P of the composite radius filler for reinforcement of the interface area 15 between the composite radius filler 3 and the structural composite parts 7.
  • the nanostructure 23 is in the form of nanofibers 22 and is oriented perpendicular against the respective structural composite part 7.
  • the aircraft structure 2 can be reinforced effective by means of a cost- effective feature in addition which is of low weight, meaning that the overall weight of the aircraft can be reduced.
  • the nanostructure 23 located in the periphery P of the radius filler 3 hinders that any cracks will appear in the radius filler 3 near the interface area 15 between the radius filler 3 and the structural composite part 7 due to overloading the aircraft.
  • the aircraft as a whole can have an optimal strength since the composite radius filler 3 will be considered as a structural part of the aircraft.
  • the nano-reinforced composite radius filler 3 will thereby provide the aircraft's structural composite parts a strength in x-, y-, z-direction.
  • the composite radius filler 3 will have a reduced thermal linear expansion and the viscosity can be controlled by the amount of nanostructure 23 for an optimal handling of the composite radius filler 3 under production.
  • the aircraft composite integrated structure as a whole will have an optimal impact strength.
  • FIG. 5 illustrates a perspective view of a prolonged radius filler 3 according to a second embodiment.
  • the nanostructure 23 comprises nanotubes 25 having a unidirectional orientation along the prolongation of the radius filler 3.
  • the nanotubes 25 are located in the periphery P of the composite radius filler 3 for reinforcement of the interface area 15 between the composite radius filler 3 and the structural composite parts 7.
  • the prolongation of the nanotubes is thus parallel with the prolongation of the radius filler, i.e. perpendicular to a plane corresponding with the shown triangular cross section of the radius filler 3.
  • FIG. 5 is shown a radius filler 3 being arranged with a protective tape 30 for protecting the radius filler and to simplify the handling of the radius filler 3 in the production line. Also the viscosity can be controlled by the amount of nanostructure for an optimal handling of the composite radius filler under production.
  • the radius filler 3 comprises a resin, which is of such a viscosity in a semi-cured state, that the radius filler is easy to handle.
  • FIG. 6 schematically illustrates a cross section of a radius filler 3 according to a th ird embodiment.
  • the nanostructure comprises nanowires 27 being oriented unhomogeneous with a major part positioned in the periphery P of the composite radius filler 3 for reinforcement of the interface area 15 between the composite radius filler 3 and the structural composite parts 7.
  • the radius filler 3 in FIG. 6 has a triangular cross-section with three convex outer surfaces 5', 5", 6.
  • the present invention is of course not in any way restricted to the preferred embodiments described above, but many possibilities to modifications, or combinations of the described embodiments, thereof should be apparent to a person with ordinary skill in the art without departing from the basic idea of the invention as defined in the appended claims.
  • the nanostructure can be embodied in epoxy ,polymides, bismaleimides, phenolics, polycyanurate, Cyanatester, PEEK, PPS, Polyester, Vinylester and other curable resins or mixtures thereof.
  • the nanostructure may be of ceramic, carbon and metal or mixtures thereof.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Composite Materials (AREA)
  • Textile Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Compositions Of Macromolecular Compounds (AREA)
  • Moulding By Coating Moulds (AREA)

Abstract

The present invention relates to a composite radius filler for use in an aircraft structure (2), which comprises structural composite parts (7) assembled together to form the aircraft structure (2). The composite radius filler (3) is arranged between rounded sections (5', 5") of the structural composite parts (7) for filling a gap (4) formed between the structural composite parts (7). The composite radius filler (3) is made structural by a nanostructure (23) arranged within the composite radius filler (3) for the reinforcement of the interface between the radius filler (3) and the structural composite part (7). The invention also relates to a method for producing an aircraft structure comprising said nano-reinforced radius filler.

Description

Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler
TECHNICAL FIELD
The present invention relates to a composite radius filler according to the preamble of cla im 1 and a method of producing an aircraft structure according to claim 10.
BACKGROUND ART
When producing an integrated monolithic structure (also called an aircraft structure such as a wing, rudder, flap etc) for an aircraft (such.), structural composite parts (such as semi-cured and pre-formed plastic beams, resin shells etc.) are cured together with a composite radius filler filling a gap between between the structural composite parts being connected. The gap is formed by at least two rounded sections of rounded or curved areas of the respective structural composite part facing each other.
The composite radius fillers are thus arranged between the rounded sections of the structural composite parts (such as composite laminates) for filling the gap between structural composite parts' rounded areas. The radius filler is thus applied in the aircraft structure for preventing deformation and/or weakening of the area near the gap or cavity otherwise present between the structural composite parts of the aircraft structure. The radius filler is often used to fill the gap between outer rounded edges of adjacent structural composite parts being bonded to each other. The structural composite parts will thus, bonded together, constitute a rigid and strong aircraft structure for use in an aircraft, wherein the structural composite parts must be of a satisfactory strength. It is of great importance to create a strong integrated monolithic structure (aircraft structure) for an aircraft, since the overall thickness of the structural composite parts can be made relatively thin and the material being used can thus be saved for achieving a lower total weight of the air craft. This means for example that less fuel has to be consumed in transport of passenger and cargo, thus being more environmental friendly and promote a cost effective transport. The aircraft structure is in this application defined as a specific structure of an aircraft, such as a wing, a fuselage, a rudder, a flap, an aileron, a fin, a tailplane etc. The aircraft structure consists of at least two assembled, and bond together, three-dimensional structural composite parts.
The structural composite parts are formed (e.g. hot drape forming or mechanical forming) earlier into three-dimensional structural composite parts and are thereafter assembled together to form the aircraft structure. The three-dimensional structural composite parts are bonded to each other by means of a bonding interlayer material, i.e. an adhesive. This adhesive is today used for holding together the aircraft structure, i.e. the assembly of the structural composite parts. The adhesive can be a melt-bondable adhesive resin, such as an epoxy. At the same date as the filing date of the present application, the applicant of this application also has filed an application reciting a bonding interlayer material comprising a nanostructure provided for connecting at least two structural composite parts (e.g. a wing beam and wing shell) together for building an aircraft structure (e.g. a wing) of an aircraft. Earlier such beam and shell connections were made by pure resin, screws, rivets etc.
Each structural composite part can be comprised of fibre reinforced plies being laid onto each other before for the forming of a stack. The stack is then formed into said structural composite part and semi-cured for effective handling.
Composite radius fillers of today, which being comprised of traditional resin filler material without any fibres, serve for filling up the gap between the structural composite parts assembled together to form the aircraft structure.
Such a composite radius filler weakens the finished aircraft structure, especially within the interface area between the composite radius filler and the surface of the structural composite parts. The composite radius filler according to prior art thus may tear apart and crack within said interface area due to tear and shear strain. Also, the radius filler may separate from the structural composite part due to heavy loads or stress.
The structural composite part of the aircraft structure is defined in this application as a specific three-dimensional structural composite part being used together with at least another specific three-dimensional structural composite part for building the aircraft structure. The three-dimensional structural composite part can be a beam, a shell, a wing rib, a bulkhead, a nose cone shell, a frame, a web stiffener etc.
The structural composite part can be made of a stack of pre-preg plies (fiber layers impregnated with resin before being placed on a temporary support by means of for example an Automatic Tape Laying-machine). The stack has plies with different fiber directions. The stack is thereafter moved to a forming tool for forming the stack into a three-dimensional shape, or at least a two- dimensional shape. When forming the stack of plies over the forming tool, a force generated from a forming medium (e.g. vacuum bag or rollers) will generate shearing forces onto the stack of plies, wherein the plies will slide against each other. This is achieved for avoiding wrinkles in the finished formed three-dimensional structural composite part. The benefit with the gliding effect or sliding between the plies is several, especially the lack of wrinkles. The finished formed structural composite part is thereafter moved to an assembly and curing tool for assembly and curing together with at least another finished formed three-dimensional structural composite part with the radius fillers arranged in specific gaps or cavities between adjacent structural composite parts.
Several solutions exist today strengthening the radius fillers being arranged between structural composite parts in an aircraft.
US 4 331 495 describes structural composite parts being joined, wherein the gaps created by the rounded surfaces are filled with radius fillers, by sewing the structural composite parts together within the area of said rounded areas and the radius filler for preventing "peel" forces tending to separate the structural composite parts at their interface.
US 5 650 229 describes a preformed radius filler which is adapted to fill a gap between composite laminates of a structural composite part. The radius filler comprises unidirectional fibers and a stabilizer as injected resin.
US 5 827 383 describes a radius filler comprising resin and fibers, wherein the radius filler is further strengthened by means of pins.
US 2003/0183067 describes a fiber reinforced radius filler having a binding agent/tackifier solution.
Known radius fillers suffer from that the peripheral area of the radius filler, which area is adjacent the interface between the radius filler and the structural composite part rounded surface, has a tendency to crack when the structural composite part is under heavy load and/or stress.
It is thus desirable to improve the strength of an integrated monolithic structure, such as an aircraft structure, being comprised of at least three assembled structural composite parts and at least one composite radius filler filling a gap between said at least three assembled structural composite parts.
It is also desirable to achieve an aircraft structure of low weight. Also is it desirable to make more effective the handling of the radius filler in the production of aircraft structures and to provide an effective method of producing a reinforced aircraft structure.
SUMMARY OF THE INVENTION
Th is has been achieved by the composite radius filler defined in the introduction being characterized by the features of the characterizing part of claim 1.
In such way the aircraft composite integrated structure as a whole can have an optimal strength since the composite radius filler by means of the present invention will be considered as a structural part of the aircraft composite integrated structure. The reinforced composite radius filler will thereby provide the aircraft composite structure a strength in x-, y-, z-direction. The composite radius filler will have a reduced thermal linear expansion and the viscosity can be controlled by the amount of nanostructure for an optimal handling of the composite radius filler under production . The aircraft composite integrated structure (aircraft structure) as a whole will thus have an optimal impact strength.
Preferably, the nanostructure primarily is arranged in the periphery of the composite radius filler for reinforcement of an interface area between the composite radius filler and the structural composite parts.
In this way the aircraft structure can be reinforced by means of a small addition of material, meaning that the overall weight of the aircraft can be reduced. The nanostructure located within the peripheral area of the radius filler hinders that cracks will appear in the radius filler near the interface between the radius filler and the structural composite part during heavy stressing of the aircraft.
Suitably, the nanostructure comprises nanofibers and/or nanotubes and/or nanowires.
Preferably, the nanostructure is oriented perpendicular against the structural composite part.
In such way the strength of the structural composite part is further increased within the area of said interface.
Suitably, the nanostructure has an unidirectional orientation.
Thereby, the strength is increased in a specific direction, e.g. corresponding with the prolongation of the radius filler.
Preferably, the composite radius filler has a triangular cross section.
In such way the radius filler can be used in a gap often appearing between three structural composite parts being joined together.
Suitably, the composite radius filler is comprised in an aileron.
In such way an aircraft structure is provided having a relatively small weight, still having a satisfactory strength.
This has also been achieved by the method defined in the introduction being characterized by the steps of claim 10. Preferably, the step of providing the composite radius filler is characterized by arranging a nanostructure primarily in the periphery of the composite radius filler for reinforcement of the interface between the composite radius filler and the structural composite parts.
Suitably, the viscosity of the resin of the radius filler can be controlled by the amount of nanostructure for an optimal handling of the composite radius filler under production. In this way a cost-effective production is achieved at the same time being flexible regarding the actual application.
BRIEF DESCRIPTION OF THE DRAWINGS
The present invention will now be described by way of example with reference to the accompanying schematic drawings, of which:
FIG. 1 illustrates a cross section of a radius filler according to prior art being subjected to stress;
FIG. 2 illustrates a curing tool for the assembly of structural composite parts formed to an aircraft structure comprising radius fillers according to a first embodiment;
FIG. 3 illustrates an assembly of structural composite parts wherein the radius filler is positioned in a gap formed between said structural composite parts;
FIG. 4 illustrates a cross section of a radius filler according to the first embodiment;
FIG. 5 illustrates a perspective view of a radius filler according to a second embodiment; and FIG. 6 illustrates a cross section of a radius filler according to a third embodiment.
DETAILED DESCRIPTION
Hereinafter, embodiments of the present invention will be described in detail with reference to the accompanying drawings, wherein for the sake of clarity and understanding of the invention some details of no importance are deleted from the drawings.
FIG. 1 schematically illustrates a cross section of a radius filler 1 of an aircraft structure 2 according to prior art which has been subjected to stress. A triangular gap 4 is filled according to the prior art with the traditional radius filler 1 being comprised of dry fibres in the form of carbon fibres (not shown) for reinforcing the connection of three surfaces 5', 5", 5'" of respective structural composite part T ', 7", 7'". A first rounded corner outer surface 5' of a first semi-cured composite beam 11 ' being connected to a second rounded corner outer surface 5" of a second semi-cured beam 11 " and to an essentially flat inside surface 5'" of a semi-cured shell 11 '". The aircraft structure 2 is after assembly cured and forms an integrated structure. During its use in an aircraft, intensive stress has (illustrated in FIG. 1 ) caused cracks 13 in an interface area 15 of the radius filler 1 , i.e. in the peripheral area of the radius filler 1 near the interface area 15 between the radius filler 1 and the structural composite part 7". There is thus a need for reinforcing the interface area 15 as it revealing has been shown in experiments that cracks can occur in this specific interface area 15. The applicant of this application has tried to solve this problem by that the laminate plies 17', 17", 17'", 17"" of the composite shell 11 '" being adhered to each other by means of a fibre reinforced epoxy. This does however not hinder the appearance of the cracks, it may even promote the causing of cracks. FIG. 2 schematically illustrates a curing tool 19 for the assembly of structural composite parts 7 formed to an aircraft structure 2, here an aileron , comprising radius fillers 3 according to a first embodiment of the present invention. The radius filler 3 is in detail explained in connection with the FIG. 4 below. The semi-cured structural composite parts 7 are assembled, wherein the radius fillers 3 according to the first embodiment are placed in gaps 4 formed between the rounded surfaces 5 of the structural composite parts T ', 7", and positioned in the curing tool 19. Two gaps 4' of the triangular gaps have each three rounded surfaces (gaps 4, 4' are illustrated in the FIGS. 2 and 3 with broken reference arrow lines and are marked for the sake of clarity as they are filled with the radius fillers 3, wherein the definition of a gap can thus be a hollow space filled with a radius filler).
The curing tool 19 thus being loaded with structural composite parts 7, each being earlier formed over by a so called ATL-machine (not shown). The curing tool 19 being loaded with the structural composite parts 7 for building the aircraft structure 2. The structural composite parts 7 being assembled are: a nose cap of reinforced resin being bonded to an upper and lower shell inner surface, a structural nose beam of composite being arranged and bonded to the web of an adjacent first structural U-beam 21 , the flanges of which being bonded to the inner surface of the shell 11 '" and bonded to the flange edges of a second structural U-beam, a third structural U-beam having its web bonded to the web of the second structural U-beam, etc. The upper 11 '" and lower shells are bonded in the rear part (not shown) of the aileron. A radius filler 3 can also thus be regarded as a structural composite part, one of which is in more detail shown in Fig. 3.
The structural composite parts 7 and the radius fillers 3 plus a bonding interlayer material 18 (see FIG. 3) adapted for connecting the structural composite parts 7 and the radius fillers 3 together -the bonding interlayer material 18 preferably comprises a curable resin and reinforcing nanostructure (such as nanofibres)-, are positioned in the curing tool 19. The curing tool 19 consisting of an upper 20' and lower 20' ' forming tool part including heating elements (not shown) for increasing the temperature of the structural composite parts 7, the radius fillers 3 and the bonding interlayer material 18 for a proper curing of the bonding interlayer material comprising the nanotubes and a proper curing and bonding of the semi-cured structural composite parts 7 to the radius fillers 3. Interior holding-on tools 24 are placed within the nose beam and the U-beams 21 for achieving a pressure from inside. Each interior holding-on tool can be divided into parts by releasing a wedge 26 arranged for keeping the parts together.
Thereafter, the aircraft structure 2 is cured and removed from the curing tool 19. In FIG. 3 is shown schematically the assembly of the structural composite parts 7 regarding one of the radius 3 fillers in FIG. 2, wherein the radius filler 3 is positioned in the radius gap 4 formed by upper rounded corners of two U-beams 21 , the webs of which is closely adhered to each other, and by the underside of the composite shell 11 '". The radius filler 3 is made of a resin filled with reinforcing nanofibres. The U-beams 21 are firstly assembled, the radius filler 3 is thereafter positioned in a v-formed groove (later on constituting walls of the gap 4) formed by said upper rounded corners. Finally, the shell 11 '" is positioned over the U-beams' 21 flanges and the radius filler 3. When all parts are in place, the structural composite parts 7 are cured for forming the aircraft structure 2.
Thus is a method of producing an aircraft composite integrated structure achieved, wherein the aircraft structure 2 comprises structural composite parts 7 assembled together to form said aircraft structure 2, wherein the composite radius filler 3 is arranged between rounded sections of at least two of the structural composite parts 7 for filling the gap 4 between the structural composite parts 7. The composite radius filler 3 is made structural by a nanostructure 23 arranged within the composite radius filler 3. The method having the steps of providing said nanostructure reinforced composite radius filler 3 by mixing a curable resin and a nanostructure 23, forming separately at least three structural composite parts 7 from pre-fabhcated pre- impregnated fibre plies, laid-up to each other and having different fibre orientations, assembling the separately formed structural composite parts 7 with said nanostructure reinforced composite radius filler 3 between rounded sections 9 of the structural composite parts 7 for filling the gap 4 between the structural composite parts 7, curing the assembled structural composite parts 7 and the nanostructure reinforced composite radius filler 3 together in a curing tool 19, and removing the finished cured aircraft composite integrated structure (aircraft structure 2) from the curing tool 19.
In FIG. 4 the radius filler 3 is schematically illustrated in a cross section according to the first embodiment. The composite radius filler is thus made structural by the nanostructure 23 arranged within the composite radius filler 3. According to this embodiment the nanostructure primarily is arranged in the periphery P of the composite radius filler for reinforcement of the interface area 15 between the composite radius filler 3 and the structural composite parts 7. The nanostructure 23 is in the form of nanofibers 22 and is oriented perpendicular against the respective structural composite part 7. In this way the aircraft structure 2 can be reinforced effective by means of a cost- effective feature in addition which is of low weight, meaning that the overall weight of the aircraft can be reduced. The nanostructure 23 located in the periphery P of the radius filler 3 hinders that any cracks will appear in the radius filler 3 near the interface area 15 between the radius filler 3 and the structural composite part 7 due to overloading the aircraft.
In this way the aircraft as a whole can have an optimal strength since the composite radius filler 3 will be considered as a structural part of the aircraft. The nano-reinforced composite radius filler 3 will thereby provide the aircraft's structural composite parts a strength in x-, y-, z-direction. The composite radius filler 3 will have a reduced thermal linear expansion and the viscosity can be controlled by the amount of nanostructure 23 for an optimal handling of the composite radius filler 3 under production. The aircraft composite integrated structure as a whole will have an optimal impact strength.
FIG. 5 illustrates a perspective view of a prolonged radius filler 3 according to a second embodiment. The nanostructure 23 comprises nanotubes 25 having a unidirectional orientation along the prolongation of the radius filler 3. The nanotubes 25 are located in the periphery P of the composite radius filler 3 for reinforcement of the interface area 15 between the composite radius filler 3 and the structural composite parts 7. The prolongation of the nanotubes is thus parallel with the prolongation of the radius filler, i.e. perpendicular to a plane corresponding with the shown triangular cross section of the radius filler 3.
In FIG. 5 is shown a radius filler 3 being arranged with a protective tape 30 for protecting the radius filler and to simplify the handling of the radius filler 3 in the production line. Also the viscosity can be controlled by the amount of nanostructure for an optimal handling of the composite radius filler under production. The radius filler 3 comprises a resin, which is of such a viscosity in a semi-cured state, that the radius filler is easy to handle.
FIG. 6 schematically illustrates a cross section of a radius filler 3 according to a th ird embodiment. The nanostructure comprises nanowires 27 being oriented unhomogeneous with a major part positioned in the periphery P of the composite radius filler 3 for reinforcement of the interface area 15 between the composite radius filler 3 and the structural composite parts 7.
The nanowires 27 located within the periphery P of the radius filler 3 hinders that any cracks appear in the radius filler 3 near the interface area 15, which is a transition zone between the radius filler 3 and the structural composite part 7, during heavy stressing of the aircraft. The radius filler 3 in FIG. 6 has a triangular cross-section with three convex outer surfaces 5', 5", 6.
The present invention is of course not in any way restricted to the preferred embodiments described above, but many possibilities to modifications, or combinations of the described embodiments, thereof should be apparent to a person with ordinary skill in the art without departing from the basic idea of the invention as defined in the appended claims. The nanostructure can be embodied in epoxy ,polymides, bismaleimides, phenolics, polycyanurate, Cyanatester, PEEK, PPS, Polyester, Vinylester and other curable resins or mixtures thereof. The nanostructure may be of ceramic, carbon and metal or mixtures thereof.

Claims

1. A composite radius filler for use in an aircraft structure (2), which comprises structural composite parts (7) assembled together to form the aircraft structure (2), the composite radius filler (3) is arranged between rounded sections (5', 5") of the structural composite parts (7) for filling a gap (4) formed between the structural composite parts (7), characterized by that the composite radius filler (3) is made structural by a nanostructure (23) arranged within the composite radius filler (3).
2. The composite radius filler according to claim 1 , wherein the nanostructure (23) primarily is arranged in the periphery (P) of the composite radius filler (3) for reinforcement of an interface area (15) between the composite radius filler (3) and the structural composite parts (7).
3. The composite radius filler according to claim 1 or 2, wherein the nanostructure comprises nanofibers (22).
4. The composite radius filler according to claim 1 or 2, wherein the nanostructure comprises nanotubes (25).
5. The composite radius filler according to claim 1 or 2, wherein the nanostructure comprises nanowires (27).
6. The composite radius filler according to any of the preceding claims, wherein the nanostructure (23) is oriented perpendicular against the structural composite part (7).
7. The composite radius filler according to any of claims 1 -5, wherein the nanostructure (23) has an unidirectional orientation.
8. The composite radius filler according to any of the preceding claims, wherein the composite radius filler (3) has a triangular cross section.
9. An aileron comprising the composite radius filler (3) according to any one of the preceding claims.
10. A method of producing an aircraft structure comprising structural composite parts (7) assembled together to form said aircraft structure (2), wherein a composite radius filler (3) is arranged between rounded sections (5', 5") of the structural composite parts (7) for filling a gap (4) formed between the structural composite parts (7), and the composite radius filler (3) is made structural by a nanostructure (23) arranged within the composite radius filler (3), the method is characterized by the steps of:
-providing said nanostructure reinforced composite radius filler (3); -forming separately at least three structural composite parts (7); -assembling the separately formed structural composite parts (7) with said nanostructure reinforced composite radius filler (3) between the rounded sections (5', 5") of the structural composite parts (7) for filling the gap (4) between the structural composite parts (7); -curing the assembled structural composite parts (7) and the nanostructure reinforced composite radius filler (3) in a curing tool (19); and -removing the finished cured aircraft structure (2) from the curing tool
(19).
11. The method according to claim 10, wherein the composite radius filler (3) is made structural by arranging a nanostructure (23) primarily in the periphery (P) of the composite radius filler (3) for reinforcement of the interface area (15) between the composite radius filler (3) and the structural composite parts (7).
PCT/SE2009/050716 2009-06-11 2009-06-11 Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler WO2010144007A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US13/377,054 US20120074265A1 (en) 2009-06-11 2009-06-11 Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler
CA2765138A CA2765138A1 (en) 2009-06-11 2009-06-11 Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler
PCT/SE2009/050716 WO2010144007A1 (en) 2009-06-11 2009-06-11 Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler
BRPI0924562A BRPI0924562A2 (en) 2009-06-11 2009-06-11 nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler.
EP09845906.8A EP2440392B1 (en) 2009-06-11 2009-06-11 Method of producing an aircraft structure comprising a nano-reinforced radius filler

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/SE2009/050716 WO2010144007A1 (en) 2009-06-11 2009-06-11 Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler

Publications (1)

Publication Number Publication Date
WO2010144007A1 true WO2010144007A1 (en) 2010-12-16

Family

ID=43309072

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/SE2009/050716 WO2010144007A1 (en) 2009-06-11 2009-06-11 Nano-reinforced radius filler for an aircraft structure and a method of producing an aircraft structure comprising such filler

Country Status (5)

Country Link
US (1) US20120074265A1 (en)
EP (1) EP2440392B1 (en)
BR (1) BRPI0924562A2 (en)
CA (1) CA2765138A1 (en)
WO (1) WO2010144007A1 (en)

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2013035538A (en) * 2011-08-10 2013-02-21 Boeing Co:The Vertical laminate noodle for high capacity pull-off for composite stringer
US20130048187A1 (en) * 2011-08-31 2013-02-28 Airbus Operations Limited Curable assembly and a filler component
JP2013059947A (en) * 2011-09-14 2013-04-04 Toyota Industries Corp Skin and rib structure
EP2610165A1 (en) * 2011-12-28 2013-07-03 Eurocopter Deutschland GmbH Composite gusset filler and method of manufacture of said composite gusset filler
EP2764987A1 (en) * 2013-02-07 2014-08-13 The Boeing Company Method and system of making composite structures having gap fillers with chopped fiber material
WO2016089340A3 (en) * 2013-10-03 2016-08-04 The Boeing Company Pre-formed thermoplastic filler for thermoset structure
EP3141377A1 (en) * 2015-07-29 2017-03-15 The Boeing Company Noodle comprised of a matrix of discontinuous filaments
CN107150785A (en) * 2016-03-04 2017-09-12 波音公司 Conductive fillet filling system and method
GB2588603A (en) * 2019-10-28 2021-05-05 Airbus Operations Ltd Noodle

Families Citing this family (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9327470B1 (en) * 2012-12-05 2016-05-03 The Boeing Company Variable-radius laminated radius filler and system and method for manufacturing same
US9475256B2 (en) * 2014-07-10 2016-10-25 The Boeing Company Composite filler
US9463864B1 (en) 2015-07-28 2016-10-11 The Boeing Company Radius filler containing vertical ply stacks and thin plies
US10195779B2 (en) * 2015-07-29 2019-02-05 The Boeing Company Systems and methods for making composite structures
US10046525B2 (en) 2016-09-30 2018-08-14 The Boeing Company Advanced variable radius laminated composite radius filler
JP6978957B2 (en) * 2018-02-13 2021-12-08 三菱重工業株式会社 Assembly manufacturing method, reinforcement members, and assembly
US11180238B2 (en) * 2018-11-19 2021-11-23 The Boeing Company Shear ties for aircraft wing
CN115256996A (en) * 2022-07-29 2022-11-01 中国航发沈阳发动机研究所 Layering method for cavity area of composite material blade

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4331495A (en) * 1978-01-19 1982-05-25 Rockwell International Corporation Method of fabricating a reinforced composite structure
EP0287103A2 (en) * 1987-04-15 1988-10-19 Boeing Company The Method and apparatus for forming cmposite radius fillers
US5650229A (en) * 1994-12-13 1997-07-22 Dow-United Technologies Composite Products Inc. Shaped unidirectional fiber preforms

Family Cites Families (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4848159A (en) * 1987-05-22 1989-07-18 The Boeing Company Ultrasonic inspection probe for laminated structures
US5876540A (en) * 1996-05-31 1999-03-02 The Boeing Company Joining composites using Z-pinned precured strips
WO2001062495A2 (en) * 2000-02-25 2001-08-30 The Boeing Company Laminated composite radius filler
US6871822B2 (en) * 2003-04-18 2005-03-29 The Boeing Company Apparatus and methods of attaching panels to support structures
US7312274B2 (en) * 2003-11-24 2007-12-25 General Electric Company Composition and method for use with ceramic matrix composite T-sections
US8691137B2 (en) * 2009-03-04 2014-04-08 The Boeing Company Method of molding partus using a tool sleeve for mold die
US8491745B2 (en) * 2007-02-03 2013-07-23 The Boeing Company Method and material efficient tooling for continuous compression molding
US7807005B2 (en) * 2006-02-02 2010-10-05 The Boeing Company Fabrication process for thermoplastic composite parts
US8388795B2 (en) * 2007-05-17 2013-03-05 The Boeing Company Nanotube-enhanced interlayers for composite structures
US20080302912A1 (en) * 2007-06-08 2008-12-11 The Boeing Company Bladderless Mold Line Conformal Hat Stringer
ES2443718T3 (en) * 2007-10-25 2014-02-20 Saab Ab Method of forming a T-shaped composite structure comprising a radio fill and radio fill
US8540833B2 (en) * 2008-05-16 2013-09-24 The Boeing Company Reinforced stiffeners and method for making the same
US8771575B2 (en) * 2011-07-27 2014-07-08 Dieffenbacher GmbH Maschinen- und Anlagenbau Methods and systems for forming reinforced composite articles having variable thickness corners
US8758879B2 (en) * 2012-06-24 2014-06-24 The Boeing Company Composite hat stiffener, composite hat-stiffened pressure webs, and methods of making the same

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4331495A (en) * 1978-01-19 1982-05-25 Rockwell International Corporation Method of fabricating a reinforced composite structure
EP0287103A2 (en) * 1987-04-15 1988-10-19 Boeing Company The Method and apparatus for forming cmposite radius fillers
US5650229A (en) * 1994-12-13 1997-07-22 Dow-United Technologies Composite Products Inc. Shaped unidirectional fiber preforms

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See also references of EP2440392A4 *

Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2013035538A (en) * 2011-08-10 2013-02-21 Boeing Co:The Vertical laminate noodle for high capacity pull-off for composite stringer
US20130048187A1 (en) * 2011-08-31 2013-02-28 Airbus Operations Limited Curable assembly and a filler component
US9079365B2 (en) * 2011-08-31 2015-07-14 Airbus Operations Limited Curable assembly and a filler component
JP2013059947A (en) * 2011-09-14 2013-04-04 Toyota Industries Corp Skin and rib structure
EP2610165A1 (en) * 2011-12-28 2013-07-03 Eurocopter Deutschland GmbH Composite gusset filler and method of manufacture of said composite gusset filler
EP2610166A3 (en) * 2011-12-28 2013-09-04 EUROCOPTER DEUTSCHLAND GmbH Composite gusset filler and method of manufacture of said composite gusset filler
RU2646400C2 (en) * 2013-02-07 2018-03-05 Зе Боинг Компани Method and system of manufacturing of the composite structures with gap fillers of the material based on chopped fiber
EP2764987A1 (en) * 2013-02-07 2014-08-13 The Boeing Company Method and system of making composite structures having gap fillers with chopped fiber material
CN103979099A (en) * 2013-02-07 2014-08-13 波音公司 Method and system of making composite structures having gap fillers with chopped fiber material
US9463880B2 (en) 2013-02-07 2016-10-11 The Boeing Company Method and system of making composite structures having gap fillers with chopped fiber material
RU2646400C9 (en) * 2013-02-07 2018-06-22 Зе Боинг Компани Method and system of manufacturing of the composite structures with gap fillers on chopped fiber
CN103979099B (en) * 2013-02-07 2017-04-12 波音公司 Method and system of making composite structures having gap fillers with chopped fiber material
US10144203B2 (en) 2013-10-03 2018-12-04 The Boeing Company Pre-formed thermoplastic filler for thermoset structure
WO2016089340A3 (en) * 2013-10-03 2016-08-04 The Boeing Company Pre-formed thermoplastic filler for thermoset structure
US10875256B2 (en) 2013-10-03 2020-12-29 The Boeing Company Method of making joint for structure
EP3141377A1 (en) * 2015-07-29 2017-03-15 The Boeing Company Noodle comprised of a matrix of discontinuous filaments
US11427687B2 (en) 2015-07-29 2022-08-30 The Boeing Company Noodle comprised of a matrix of discontinuous filaments
CN107150785A (en) * 2016-03-04 2017-09-12 波音公司 Conductive fillet filling system and method
CN107150785B (en) * 2016-03-04 2021-07-06 波音公司 Conductive fillet packing system and method
GB2588603A (en) * 2019-10-28 2021-05-05 Airbus Operations Ltd Noodle

Also Published As

Publication number Publication date
EP2440392B1 (en) 2017-10-25
EP2440392A1 (en) 2012-04-18
CA2765138A1 (en) 2010-12-16
EP2440392A4 (en) 2015-05-20
BRPI0924562A2 (en) 2016-03-01
US20120074265A1 (en) 2012-03-29

Similar Documents

Publication Publication Date Title
EP2440392B1 (en) Method of producing an aircraft structure comprising a nano-reinforced radius filler
KR102047554B1 (en) Composite hat stiffener, composite hat-stiffened pressure webs, and methods of making the same
US9114588B2 (en) Skin-stiffener transition assembly, method of manufacture and application of said skin-stiffener transition assembly
EP2727711B1 (en) Composite radius fillers and methods of forming the same
US9291151B2 (en) Wind turbine blade and its producing method
KR102197337B1 (en) Composite structure having a stabilizing element
US9359060B2 (en) Laminated composite radius filler with geometric shaped filler element and method of forming the same
JP5308533B2 (en) Composite structure, aircraft main wing and aircraft fuselage provided with the same
CN107813920B (en) Open channel stiffener
US20120148789A1 (en) Aircraft structure with structural parts connected by nanostructure and a method for making said aircraft structure
JP2014151648A (en) Method and system of making composite structures having gap fillers with chopped fiber material
US20030190455A1 (en) Textile joint reinforcement and associated method
US10005267B1 (en) Formation of complex composite structures using laminate templates
WO2020003608A1 (en) Composite material, method for manufacturing composite material, and method for curing composite material

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 09845906

Country of ref document: EP

Kind code of ref document: A1

WWE Wipo information: entry into national phase

Ref document number: 9449/DELNP/2011

Country of ref document: IN

WWE Wipo information: entry into national phase

Ref document number: 13377054

Country of ref document: US

REEP Request for entry into the european phase

Ref document number: 2009845906

Country of ref document: EP

WWE Wipo information: entry into national phase

Ref document number: 2765138

Country of ref document: CA

Ref document number: 2009845906

Country of ref document: EP

NENP Non-entry into the national phase

Ref country code: DE

REG Reference to national code

Ref country code: BR

Ref legal event code: B01A

Ref document number: PI0924562

Country of ref document: BR

ENP Entry into the national phase

Ref document number: PI0924562

Country of ref document: BR

Kind code of ref document: A2

Effective date: 20111208