EP1427581A2 - Structure composite et procede de construction correspondant - Google Patents

Structure composite et procede de construction correspondant

Info

Publication number
EP1427581A2
EP1427581A2 EP02763675A EP02763675A EP1427581A2 EP 1427581 A2 EP1427581 A2 EP 1427581A2 EP 02763675 A EP02763675 A EP 02763675A EP 02763675 A EP02763675 A EP 02763675A EP 1427581 A2 EP1427581 A2 EP 1427581A2
Authority
EP
European Patent Office
Prior art keywords
panel
stiffening
skin
skin panel
composite material
Prior art date
Legal status (The legal status 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 status listed.)
Withdrawn
Application number
EP02763675A
Other languages
German (de)
English (en)
Inventor
Elbert Lee Mckague, Jr.
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Lockheed Martin Corp
Original Assignee
Lockheed Corp
Lockheed Martin Corp
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 Lockheed Corp, Lockheed Martin Corp filed Critical Lockheed Corp
Publication of EP1427581A2 publication Critical patent/EP1427581A2/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • 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
    • B29D99/001—Producing wall or panel-like structures, e.g. for hulls, fuselages, or buildings
    • B29D99/0014—Producing wall or panel-like structures, e.g. for hulls, fuselages, or buildings provided with ridges or ribs, e.g. joined ribs
    • 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
    • B29C70/16—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length
    • B29C70/22—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length oriented in at least two directions forming a two-dimensional [2D] structure
    • B29C70/222—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length oriented in at least two directions forming a two-dimensional [2D] structure the structure being shaped to form a three dimensional configuration
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/06—Fibrous reinforcements only
    • B29C70/10—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres
    • B29C70/16—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length
    • B29C70/24—Fibrous reinforcements only characterised by the structure of fibrous reinforcements, e.g. hollow fibres using fibres of substantial or continuous length oriented in at least three directions forming a three-dimensional [3D] structure
    • 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
    • B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core

Definitions

  • the present invention relates generally to the field of composites construction, and more particularly to a composite structure and method for constructing same.
  • Composite structures are desirable in many industries for many applications. For example, aircraft, space, and land/sea vehicles employ a variety of curved and multiple- contoured surface structures in their fabrication. Composite materials are commonly used for these structures because, among other desirable attributes, composite materials have high strength-to-weight ratios. Even so, composite structures formed from composite materials oftentimes need to be stiffened. Therefore, manufacturers of composite structures are continually searching for better and more economical ways of stiffening composite structures.
  • a method of constructing a composite structure includes positioning a plurality of forming elements on a skin panel formed from a composite material in a predetermined configuration, disposing a stiffening panel formed from an uncured composite material outwardly from the forming elements to create a plurality of contact regions between the skin panel and the stiffening panel, and curing the skin panel and the stiffening panel to bond the skin panel and the stiffening panel together at the contact regions.
  • Embodiments of the invention provide a number of technical advantages. Embodiments of the invention may include all, some, or none of these advantages. In one embodiment, less weight of stiffened composite structures is achieved. This weight reduction is particularly advantageous in aircraft applications. No mechanical fasteners or adhesives are needed to attach the stiffening panel to the composite panel, which saves considerable time and money in constructing stiffened composite structures. However, Z-pins may be used to complement the bonding of the stiffening panel to the composite panel to create crack propagation resistance. An additional technical advantage is that the quality of stiffened composite structures is improved by substantially reducing or eliminating folds, kinks, bumps, or other imperfections when laying-up composite stiffening panels on composite panels.
  • FIGURE 1 is a perspective view of an aircraft having a panel formed from a composite structure constructed according to one embodiment of the present invention
  • FIGURE 2 is a perspective view of the composite structure of FIGURE 1;
  • FIGURES 3 A through 3C are perspective elevational views illustrating one method of constructing a composite structure according to one embodiment of the present invention.
  • FIGURE 3D is a perspective elevational view of a composite structure in a waffle configuration constructed according to one embodiment of the present invention.
  • FIGURES 1 through 3D of the drawings in which like numerals refer to like parts.
  • FIGURE 1 is a perspective view of an aircraft 100 having a panel 102 formed from a composite structure 200 (FIGURE 2) constructed according to one embodiment of the present invention.
  • Aircraft 100 may be any suitable aircraft and panel 102 may be any suitable structural panel on aircraft 100, such as a wing panel, a tail panel, or a fuselage panel.
  • panel 102 may be employed in any suitable aircraft, space, land/sea vehicle, or other machines, devices, or structures formed by composite materials. Oftentimes, structures formed from composite materials need to be stiffened.
  • the following detailed description uses an aircraft application to illustrate one or more embodiments of composite structure 200 manufactured according to the teachings of the present invention.
  • One embodiment of composite structure 200 is illustrated below in conjunction with FIGURE 2.
  • FIGURE 2 is a perspective view of one embodiment of composite structure 200.
  • Composite structure 200 includes a skin panel 202, a plurality of forming elements 204, and a stiffening panel 206.
  • Composite structure 200 may also include a plurality of fasteners 208 coupling skin panel 202 and stiffening panel 206.
  • Skin panel 202 is formed from any suitable composite material, such as graphite epoxy or graphite bismaleimide ("bmi"). Skin panel 202 may have any suitable number of layers and the fibers that are included in skin panel 202 may be unidirectional, bidirectional, or have any suitable orientation. In the illustrated embodiment, skin panel 202 forms a portion of an outer skin of aircraft 100. For example, skin panel 202 may coincide with a portion of the outer skin of a wing, a tail section, or a fuselage section. Accordingly, skin panel 202 may have any suitable shape, dimensions, and thickness. In addition, skin panel 202 may be substantially flat or may have one or more contours to conform to the shape of a particular portion of aircraft 100.
  • bmi graphite epoxy or graphite bismaleimide
  • Forming elements 204 are utilized to impart a predetermined configuration to stiffening panel 206, as described more fully below in conjunction with FIGURES 3A through 3D.
  • the properties and weight of forming elements 204 are selected according to end-use requirements.
  • forming elements 204 are "flyaway" tooling mandrels formed from a lightweight material suitable for permanent incorporation into composite structure 200, such as low-density, closed cell epoxy or bmi foam, or other suitable polymeric foams, such as carbon or graphitic foam, having suitable compression strength.
  • suitable materials may be such materials as balsa wood, honeycomb structures, or open cell foams.
  • forming elements 204 are such that after composite structure 200 is formed, forming elements 204 are removable via any suitable meltout or washout technique.
  • forming elements 204 may be composed of a thin shell of material that is filled with a meltout or washout type of material where the thin shell is suitable for permanent incorporation into composite structure 200.
  • a thin shell could be formed from a thermoformed plastic, such as ABS, and filled with a removable material, such as a water soluble eutectic salt.
  • forming elements 204 may be any suitable size or shape and may form any suitable configuration when positioned on skin panel 202. In the illustrated embodiment, forming elements 204 form a corrugated configuration on skin panel 202; however, other configurations, such as waffle configurations may be formed.
  • Stiffening panel 206 is formed from any suitable composite material, such as graphite epoxy or graphite bmi. To facilitate bonding between stiffening panel 206 and skin panel 202, stiffening panel 206 typically includes a resin that is the same as the resin in skin panel 202; however, other suitable resins may be used. In one embodiment, stiffening panel 206 is formed from a composite material having a plurality of discontinuous fibers so that stiffening panel 206 may be flexible enough to form over forming elements 204, as discussed more fully below in conjunction with FIGURES 3A through 3D.
  • stiffening panel 206 is formed from a composite material having chopped fibers that is sprayed on using well-known spraying techniques onto forming elements 204 and skin panel 202, as discussed more fully below.
  • the fibers used in the composite material for stiffening panels 206 may be unidirectional, bi-directional, or have any suitable orientation.
  • Stiffening panel 206 may be formed from any suitable number of layers and may have any suitable size or shape before being placed onto forming elements 204 and skin panel 202.
  • stiffening panel 206 has the same width as skin panel 202 but is longer than skin panel 202 such that when stiffening panel 206 forms over forming elements 204 it ends up being approximately the same length as skin panel 202. This reduces or eliminates any trimming or finishing operations after composite structure 200 is formed.
  • Fasteners 208 are Z pins; however, other suitable fasteners may be used.
  • One purpose of fasteners 208 is for damage resistance. In other words, if a crack starts to develop in composite structure 200 fasteners 208 inhibit crack propagation. As illustrated in FIGURE 2, fasteners 208 couple skin panel 202 and stiffening panel 206 at one or more contact regions 210. Fasteners 208 may be inserted by any suitable process, such as pushing or driving while being vibrated with ultrasonic energy.
  • FIGURES 3 A through 3C are perspective elevational views illustrating one method of constructing composite structure 200.
  • the method starts with skin panel 202, as illustrated in FIGURE 3A.
  • Skin panel 202 which as described above is formed from any suitable composite material having any suitable dimensions, is provided in either a cured or uncured state.
  • Skin panel 202 is placed on any appropriate flat or suitably contoured working table or tool (not explicitly shown).
  • Forming elements 204 are positioned on skin panel 202 in a predetermined configuration.
  • the predetermined configuration is a corrugated configuration; however, other suitable configurations are feasible, such as a waffle configuration (FIGURE 3D).
  • forming elements 204 may have tapered ends.
  • Forming elements 204 may be located on skin panel 202 by any suitable means, such as a laser positioning locator, a hard positioning template, or by measuring from one or more reference points.
  • stiffening panel 206 is applied on the top of skin panel 202 and forming elements 204, as illustrated in FIGURE 3C.
  • FIGURE 3C shows stiffening panel 206 draped over forming elements 204, thereby forming contact regions 210 between stiffening panel 206 and skin panel 202.
  • stiffening panel 206 is typically formed from an uncured composite material.
  • the fibers that are part of the composite material that stiffening panel 206 is formed from are, in one embodiment, discontinuous fibers.
  • the fibers may, in other embodiments, be continuous; however, if continuous then the direction of the fibers needs to be in approximately the same direction as the direction of stretching of stiffening panel 206.
  • stiffening panel 206 An important consideration when applying stiffening panel 206 is that substantially no air gaps exist between the bottom surface of stiffening panel 206 and the top surface of skin panel 202. This insures that adequate bonding takes place between stiffening panel 206 and skin panel 202 at contact regions 210 when the composite material is cured, as described more fully below. As an option, although not required, adhesive may be applied at contact regions 210 if so desired. However, an important technical advantage of one embodiment of the present invention is that stiffening panel 206 and skin panel 202 may be coupled to one another without the use of adhesives and/or mechanical fasteners. This saves considerable time and money in constructing composite structure 200.
  • stiffening panel 206 is sprayed on the surface of skin panel 202 and forming elements 204 using any suitable well-known composite spraying techniques.
  • This embodiment may be particularly suitable for stiffening panels that are formed in a waffle configuration, such as stiffening panel 206 illustrated in FIGURE 3D.
  • the next step in the method is to cure skin panel 202 and stiffening panel 206, whereby skin panel 202 and stiffening panel 206 are bonded together at contact regions 210.
  • Any suitable curing techniques well-known in the art of composites forming may be utilized.
  • composite structure 200 may be enveloped in a vacuum bag, a vacuum pulled, and then placed in an autoclave with suitable temperature and pressure to cure the composite materials.
  • the matrix used in the composite materials may be thermosetting or thermoplastic. If thermoplastic, then higher curing temperatures would be involved.
  • Composite structure 200 is heated and pressurized for a predetermined time so that the composite material may sufficiently cure. Any trimming or finishing of composite structure 200 may then take place, which ends one method of forming composite structure 200.
  • composite structure 200 may be utilized. For example, two mechanical presses may be used to bond stiffening panel 206 to skin panel 202. In this embodiment, composite structure 200 would be heated to a sufficient temperature before the presses are pressed together for a sufficient time to cure the composite materials. In another embodiment, a resin transfer molding ("RTM") process may be used, wherein dried fabric is used for stiffening panel 206 and skin panel 202 and injected with any suitable resin. Other suitable curing processes may be utilized. Although not illustrated in FIGURES 3 A through 3C, another step in the method may be to couple skin panel 202 and stiffening panel 206 with fasteners 208 proximate contact regions 210.
  • RTM resin transfer molding
  • skin panel 202 would need to be in an uncured or partially cured state so that skin panel 202 is deformable enough to facilitate the inserting of fasteners 208.
  • fasteners 208 may be Z-pins, which act to resist any damage during usage of composite structure 200.
  • forming elements 204 would then have to be formed from a material or materials that facilitate their removal.
  • forming elements 204 each may be a rubber bladder filled with sand or eutectic salt with a vacuum pulled on it.
  • Other suitable meltout or washout type of materials may also be used.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Composite Materials (AREA)
  • Textile Engineering (AREA)
  • Architecture (AREA)
  • Civil Engineering (AREA)
  • Structural Engineering (AREA)
  • Moulding By Coating Moulds (AREA)
  • Laminated Bodies (AREA)

Abstract

Selon un mode de réalisation de l'invention, un procédé permettant de construire une structure composite consiste à positionner une pluralité d'éléments de formation sur un panneau de revêtement constitué d'un matériau composite dans une configuration prédéterminée, à disposer un panneau de renfort constitué d'un matériau composite non durci à l'extérieur des éléments de formation afin de créer une pluralité de régions de contact entre le panneau de revêtement et le panneau de renfort, enfin à durcir le panneau de revêtement et le panneau de renfort afin de les coller l'un à l'autre au niveau des régions de contact.
EP02763675A 2001-09-20 2002-09-19 Structure composite et procede de construction correspondant Withdrawn EP1427581A2 (fr)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US09/961,084 US20030051434A1 (en) 2001-09-20 2001-09-20 Composite structure and method for constructing same
US961084 2001-09-20
PCT/US2002/029912 WO2003024700A2 (fr) 2001-09-20 2002-09-19 Structure composite et procede de construction correspondant

Publications (1)

Publication Number Publication Date
EP1427581A2 true EP1427581A2 (fr) 2004-06-16

Family

ID=25504033

Family Applications (1)

Application Number Title Priority Date Filing Date
EP02763675A Withdrawn EP1427581A2 (fr) 2001-09-20 2002-09-19 Structure composite et procede de construction correspondant

Country Status (3)

Country Link
US (1) US20030051434A1 (fr)
EP (1) EP1427581A2 (fr)
WO (1) WO2003024700A2 (fr)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7374715B2 (en) * 2002-05-22 2008-05-20 Northrop Grumman Corporation Co-cured resin transfer molding manufacturing method
US7419627B2 (en) * 2002-09-13 2008-09-02 Northrop Grumman Corporation Co-cured vacuum-assisted resin transfer molding manufacturing method
DE102011120636A1 (de) 2011-12-09 2013-06-13 Airbus Operations Gmbh Faserverbundbauteilanordnung mit mindestens zwei plattenförmigen Faserverbundbauteilen sowie Verfahren zur Herstellung derselben

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Also Published As

Publication number Publication date
US20030051434A1 (en) 2003-03-20
WO2003024700A3 (fr) 2003-07-10
WO2003024700A2 (fr) 2003-03-27

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