WO2011066950A1 - Élément de fuselage d'avion - Google Patents

Élément de fuselage d'avion Download PDF

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Publication number
WO2011066950A1
WO2011066950A1 PCT/EP2010/007269 EP2010007269W WO2011066950A1 WO 2011066950 A1 WO2011066950 A1 WO 2011066950A1 EP 2010007269 W EP2010007269 W EP 2010007269W WO 2011066950 A1 WO2011066950 A1 WO 2011066950A1
Authority
WO
WIPO (PCT)
Prior art keywords
opening
fuselage
component
fuselage component
aircraft fuselage
Prior art date
Application number
PCT/EP2010/007269
Other languages
German (de)
English (en)
Inventor
Andreas Knote
Steffen Niemann
Henrik Borgwardt
Robert Kaps
Tobias Ströhlein
Christian Hühne
Original Assignee
Deutsches Zentrum für Luft- und Raumfahrt e.V.
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 Deutsches Zentrum für Luft- und Raumfahrt e.V. filed Critical Deutsches Zentrum für Luft- und Raumfahrt e.V.
Publication of WO2011066950A1 publication Critical patent/WO2011066950A1/fr

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/68—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts by incorporating or moulding on preformed parts, e.g. inserts or layers, e.g. foam blocks
    • B29C70/86—Incorporated in coherent impregnated reinforcing layers, e.g. by winding
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/06—Fibrous reinforcements only
    • B29C70/08—Fibrous reinforcements only comprising combinations of different forms of fibrous reinforcements incorporated in matrix material, forming one or more layers, and with or without non-reinforced layers
    • B29C70/086—Fibrous reinforcements only comprising combinations of different forms of fibrous reinforcements incorporated in matrix material, forming one or more layers, and with or without non-reinforced layers and with one or more layers of pure plastics material, e.g. foam layers
    • 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
    • B64C1/14—Windows; Doors; Hatch covers or access panels; Surrounding frame structures; Canopies; Windscreens accessories therefor, e.g. pressure sensors, water deflectors, hinges, seals, handles, latches, windscreen wipers
    • B64C1/1407—Doors; surrounding frames
    • B64C1/1461—Structures of doors or surrounding frames
    • 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
    • B29K2105/04—Condition, form or state of moulded material or of the material to be shaped cellular or porous
    • B29K2105/046—Condition, form or state of moulded material or of the material to be shaped cellular or porous with closed cells
    • 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
    • B29K2995/00—Properties of moulding materials, reinforcements, fillers, preformed parts or moulds
    • B29K2995/0037—Other properties
    • B29K2995/0089—Impact strength or toughness
    • 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
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C2001/0054—Fuselage structures substantially made from particular materials
    • B64C2001/0072—Fuselage structures substantially made from particular materials from composite materials
    • 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

Definitions

  • the invention relates to an aircraft fuselage component with an outer wall with reinforcing elements and with an opening in the outer wall.
  • Aircraft door environment structures are conventionally constructed by making a framework of frames with main and secondary frames and side rails to which the skin is attached.
  • the structure of the fuselage is defined by large cutouts, such as e.g. noticeably weakened for the door opening. Static and dynamic loads acting on the door environment structure must therefore be redirected around the large area using a specially designed and enhanced environment structure.
  • Internal pressurization of an aircraft fuselage during flight caused by the differential pressure between the aircraft interior side and the aircraft exterior side, is initiated at discrete points from the door into the aircraft fuselage. The forces occurring at these discrete points provide for localized load overshoot, which must be suitably transmitted through the door environment structure into the fuselage.
  • the door environment structure thus supports the large door cutout and introduces the loads from the door into the fuselage.
  • the door environment structure is made in differential construction.
  • the functions of the individual components, in particular the main and secondary frames, the door side rails, the intercostal and the fuselage skin are clearly separated.
  • the differential design requires a high assembly cost to assemble the items. This leads to a large amount of time and high costs. Furthermore, created by the joining of many items a long tolerance chain, so that the production of the individual parts and the
  • CONFIRMATION COPY Assembly must be done with high precision. Furthermore, in the area of the door environment structure, to the left and right of the door cut-out, the rib of the fuselage is continued in the radial direction, and intercostals are used at the level of the door's load introduction points. This gives the door environment structure its typical ladder shape.
  • a disadvantage of the differential design is also the high cost of repair and replacement of damaged items in case of failure.
  • DE 10 2004 009 020 A1 describes intercosts for frames in aircraft, with which a load can be derived from a first frame to a second frame and / or to the outer skin of the aircraft. With the help of these separate intercostal a targeted power transmission, especially in the field of door openings, possible.
  • the disadvantage of conventional aircraft fuselage components is the high assembly costs due to the number of individual parts and the necessary compliance with tolerances.
  • damaged items are very difficult to replace.
  • the optical detectability of damage is difficult or impossible with fiber composite materials. This increases the risk of discovering a damaged structure only after many flight cycles during extensive maintenance.
  • Another problem is that repairs, especially on the fuselage skin in the door area are very expensive, since the outer skin undergoes very high shear stress at this large section.
  • the object of the present invention is therefore to provide an improved aircraft fuselage component with outer wall, reinforcing elements and an opening in the outer wall.
  • the object is achieved with the aircraft fuselage component of the aforementioned type in that the aircraft fuselage component is formed as an integral fiber composite element, in which adjacent to the opening and the outer wall, a reinforcing fiber webbing for forming the reinforcing elements is arranged, wherein the reinforcing fiber web is integrally connected to the outer wall and wherein the fiber orientation of the reinforcing fiber scrim is adapted to divert loads around the aperture and divert loads applied to load introduction points at the aperture into the side surface regions of the aircraft fuselage.
  • an aircraft fuselage component as an integral fiber composite element, in which the reinforcing elements are not designed, as usual, discretely from individual components each having an assigned function. Rather, the pronounced anisotropy of fiber composites is exploited, which can accommodate significantly higher tensile loads in the fiber longitudinal direction, as transverse to it.
  • This makes it possible to form the aircraft fuselage as an integral fiber composite element with integrated reinforcing fiber, in which the fiber orientation is adapted to optimize the environmental structure of the opening.
  • the fiber orientation of the reinforcing fiber fabric makes it possible to divert the cutting loads which are introduced from the aircraft fuselage and, for example, a large opening into the environmental structure of the opening, around the large cutout, ie the opening.
  • the opening is preferably provided for receiving an aircraft door.
  • aircraft are to be understood as meaning any aircraft, but in particular, commercial aircraft and helicopters, in which the aircraft fuselage component can be advantageously used.
  • the reinforcing fiber fabric is covered with a sacrificial layer for detecting damage, for example by impact stress.
  • the sacrificial layer may e.g. a polymer foam with a closed-pored surface.
  • the polymer foam absorbs impact stress to a certain extent. In case of overuse, the pores collapse and the damage becomes visible on the surface by a duckling.
  • a similar principle can also be achieved with a honeycomb core structure with a thin cover layer as the sacrificial layer on the outer skin.
  • the aircraft fuselage component has a separate supply element that is detachable from the outer wall and the reinforcing fiber web formed integrally with the outer wall. Not all To disassemble supply lines for electrical, air, water and hydraulics of the aircraft around the opening during an exchange of the fuselage component, these supply lines are mounted on this separate carrier element decoupled from the aircraft fuselage component. When the aircraft fuselage component is replaced, the carrier element only has to be detached from it and remains in the aircraft. For this purpose, it is releasably connected to only a few defined and easily dismantled bearings with the aircraft fuselage.
  • This separate detachable support element for supply lines thus has a significant advantage over conventional door environment structures in which the supply lines are nested housed in the door environment structure and removable only with great effort.
  • the main direction of extension of the fibers of the reinforcing fiber fabric is preferably adapted to the direction of the load acting on the aircraft fuselage, such that the fibers redistribute the flows of load distributed from the opening into the peripheral regions of the aircraft fuselage and the flows from the peripheral regions of the aircraft fuselage and loads acting on the outer wall the opening around in other edge regions of the fuselage component derives.
  • a first laminate of a reinforcing fiber fabric with a fiber orientation in the main direction of the fibers is provided such that reinforcing fibers extend from the side edges of the aircraft body component diagonally past the opening to the upper edges and / or lower edges of the aircraft fuselage.
  • This first layer of reinforcing fiber fabric thus provides for the load redirection of loads on the side edges of the aircraft fuselage and on the outer skin around the opening.
  • a second structure which is produced integrally with the first laminate, of a reinforcing fiber scrim having a fiber orientation in the main extension direction of the fibers is provided such that reinforcing fibers extend from the region adjacent to the opening from the opening extend away in the direction of the side edges, wherein a plurality of reinforcing fiber bundles are arranged at a distance from each other.
  • This second structure of the reinforcing fiber fabric serves to dissipate it onto the opening, e.g. loads acting on a door in the opening into the aircraft fuselage component and for as even as possible a planar discharge of this load from the opening to the fuselage in the vicinity of the aircraft fuselage component.
  • the derivation of the load from a load introduction point at the opening to the side edges of the aircraft fuselage component preferably does not take place simply by means of a reinforcing fiber layer directed outwardly from the load introduction point. Rather, it is advantageous if the reinforcing fiber bundles extend with part of their length along an edge region of the opening.
  • the reinforcing fiber bundles of the second layer of the reinforcing fiber scrim should preferably completely enclose the opening altogether so as not only to dissipate loads from certain load transfer points, but to pick up the loads around the circumference of the opening and dissipate outwardly therefrom over comb-like reinforcing fiber bundles.
  • the opening is advantageously reinforced by the reinforcing fiber bundles extending along edge regions of the opening in this way.
  • Figure 1 sketch of an aircraft with openings in the fuselage
  • FIG. 2 shows a sketch of an aircraft fuselage component with a door opening and integrally formed reinforcing fiber layers
  • FIG. 3 shows a three-dimensional view of a section of an aircraft fuselage with an aircraft fuselage component inserted therein for a door opening.
  • FIG. 1 shows a sketch of an aircraft 1 in whose fuselage 2 openings 3 for doors and windows are provided.
  • FIG. 2 shows a sketch of an aircraft fuselage component 4 according to the invention for use in the region of such an opening 9, in particular a door opening.
  • the aircraft fuselage component 4 is designed as an integral fiber composite element and has an outer wall 5 (outer skin). Adjacent to this outer wall 5, a first laminate 6 of a reinforcing fiber scrim 7 is integrally formed therewith.
  • This first layer 6 of the reinforcing fiber fabric 7 essentially serves to dissipate loads from the side edges 8a, 8b including the upper edge 8c and lower edge 8d around the opening 9 into the aircraft fuselage as evenly as possible and to guide them out of them.
  • the fiber orientation of the fibers of this reinforcing fiber scrim 7 of the laminate 6 is optimized such that its main extension direction is directed diagonally past the opening 9 from the side edges 8a, 8b to the upper edges 8c and lower edges 8d of the aircraft fuselage component 4, respectively.
  • Diagonal reinforcing fiber bundles can furthermore be provided at the four corners, with the result that reinforcing fibers flow from one side edge 8a or 8b into the outer area of the adjacent upper edge 8c or lower edge 8d or Reinforcing fibers from the central region of the upper edge 8c and lower edge 8d in the outer region of the respective upper edge 8c and lower edge 8d are guided past the opening 9.
  • a second structure 10 with reinforcing fiber webs 7 is provided to guide load application points at the opening 9 away from the opening 9 into the aircraft fuselage component 4 and to distribute it in an areally possible manner.
  • the main direction of extension of the fibers of this second structure 10 of the reinforcing fiber fabric 7 is optimized such that reinforcing fibers extend from the region adjacent to the opening 9 from the opening 9 in the direction of the side edges 8a, 8b or possibly the upper edges 8c and lower edges 8d as side edges ,
  • a plurality of reinforcing fiber bundles are arranged at a distance from one another, so that the reinforcing fiber bundles extend in a finger-like manner away from the opening 9.
  • These fingers of the reinforcing fiber bundles are spaced apart from each other and extend to each other so that a comb-like structure is formed, is derived by the load surface of the edge region of the opening 9 in the aircraft fuselage component 4. From there, the load can then be discharged relatively evenly over the side edges 8a, 8b to the body regions adjacent to the aircraft fuselage component 4.
  • the area of the illustrated door environment structure is thus considered as an arbitrary structure to be optimized.
  • the opening 9 such as the door opening
  • around the cutting loads are considered, which act from the rest of the fuselage structure on the illustrated environment structure.
  • the loads from the door to the door environment structure are considered.
  • shear loads are considered as cutting loads.
  • the reinforcing fiber scrims 7 now exploit the pronounced anisotropy of fiber composites, which can absorb significantly higher tensile loads in the fiber longitudinal direction than transversely thereto.
  • the aircraft fuselage component 4 is thus designed as an optimized surface structure with ideal fiber orientation.
  • the cutting loads introduced from the fuselage and the door into the door environment structure may be along the ideal fiber orientation the load paths are passed around the large opening 9. It can be seen that the aircraft fuselage component is not bound to the geometric boundary conditions of existing frames and intercostals, so that the weight of the aircraft fuselage component can be reduced compared to conventional door opening structures.
  • the exact shape of the optimized structure depends on the type of aircraft and the structural design of a door or tailgate in the opening 9 and the hull design and must be optimally designed accordingly for the prevailing loads. It is important to exploit the optimized fiber orientation so that the tensile loads are optimally absorbed in the fiber longitudinal direction and passed around the opening 9.
  • the fully integral aircraft fuselage component 4 can be manufactured either in resin injection process in one shot. It is also conceivable, however, for the one or more layers of reinforcing fiber ply 7 to be glued to the outer skin or riveted to the outer skin. In this case, the reinforcing fiber scrims would be made as sub-modules which are then integrally bonded to the outer skin.
  • FIG. 3 shows a representation of a section of an aircraft fuselage 2 with the usual frames 1 1 and stringers 12. It is clear that the aircraft fuselage component 4 shown in FIG. 2 is inserted into a section of the fuselage 2. In this case, via the ribs 1 1 and stringer 12 and optionally via the outer wall 5 introduced loads on the reinforcing fiber scrim 7 around the opening 9 around. This load redirection is preferably such that the hull 2 is still sufficiently strong despite the opening 9 and the aircraft fuselage component 4 used therein. The weakening of the hull 2 caused by the opening 9 is compensated by the reinforcing fiber webs 7, which are formed integrally with the outer wall 5 and adapted to the opening 9 and the load flow.
  • FIG. 3 shows the view of a half-shell of the fuselage 2 from the inside on the inside of the fuselage component 4.
  • Support elements 13 for supply lines (not shown) can be seen, which are configured as separate components for the fuselage component 4.
  • These carrier elements 13 are connected to as few bearings as possible with the aircraft fuselage component 4 and carry Supply lines for eg electrical, air, water and hydraulics of the aircraft 1.
  • an optional sacrificial layer may be provided on the outer side of the fuselage in the vicinity of the opening 9 or the door, which protects the aircraft fuselage component 4 from damage and damages e.g. optically visible. Impact damage on the outer wall 5 and the sacrificial layer arranged on the outer side of the outer wall 5 are visually visible through deformation of the sacrificial layer due to the plastic deformation of the sacrificial layer. Such a damage can then be easily detected and lead to the replacement of the aircraft fuselage component or only the sacrificial layer 4.
  • the sacrificial layer may e.g. a polymer foam with a closed-pored surface. The polymer foam absorbs impact stress to a certain extent.
  • the sacrificial layer can also be achieved with the aid of a honeycomb core structure with a thin cover layer as the sacrificial layer on the outer skin.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Composite Materials (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Laminated Bodies (AREA)
  • Moulding By Coating Moulds (AREA)

Abstract

Élément de fuselage d'avion (4) comportant une paroi externe (5), des éléments de renforcement et une ouverture (3,9) ménagée dans la paroi externe (5). Selon l'invention, cet élément de fuselage d'avion (4) est un élément intégral composite renforcé par des fibres sur lequel une grille non tissée de fibres de renforcement (7) est disposée de manière contiguë à l'ouverture (3,9) et à la paroi externe (5) pour former les éléments de renforcement, la grille non tissée de fibres de renforcement (7) faisant partie intégrante de la paroi externe (5), et l'orientation des fibres de la grille non tissée de fibres de renforcement (7) étant adaptée pour dévier des charges autour de l'ouverture (3,9) et pour évacuer des charges agissant sur des points d'induction de charges au niveau de l'ouverture (3,9) vers les zones superficielles latérales de l'élément de fuselage d'avion (4).
PCT/EP2010/007269 2009-12-03 2010-12-01 Élément de fuselage d'avion WO2011066950A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102009056533.7 2009-12-03
DE102009056533.7A DE102009056533B4 (de) 2009-12-03 2009-12-03 Flugzeugrumpfbauteil

Publications (1)

Publication Number Publication Date
WO2011066950A1 true WO2011066950A1 (fr) 2011-06-09

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ID=43821833

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2010/007269 WO2011066950A1 (fr) 2009-12-03 2010-12-01 Élément de fuselage d'avion

Country Status (2)

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DE (1) DE102009056533B4 (fr)
WO (1) WO2011066950A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10179438B2 (en) 2016-09-23 2019-01-15 Bell Helicopter Textron Inc. Method and assembly for manufacturing door skin and wall with doorway
US10450048B2 (en) 2015-11-09 2019-10-22 Airbus Operations Gmbh Aircraft fuselage structure

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102009057018B4 (de) * 2009-12-04 2016-09-15 Airbus Defence and Space GmbH Flugzeugrumpfstruktur
DE102017126052A1 (de) 2017-11-08 2019-05-09 Airbus Operations Gmbh Versteifungsanordnung für eine Öffnung in einer Flugzeugstruktur
CN110155304B (zh) * 2019-01-25 2021-11-12 北京机电工程研究所 抗横向冲击大开口舱段结构及具有其的飞行器
DE102020201779B4 (de) * 2020-02-13 2021-12-02 Premium Aerotec Gmbh Türsystem für ein Luftfahrzeug und Luftfahrzeug

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102004009020A1 (de) 2004-02-25 2005-09-15 Airbus Deutschland Gmbh Zwischenkostal für Flugzeuge
FR2869871A1 (fr) * 2004-05-04 2005-11-11 Airbus France Sas Dispositif de protection et revelation de choc pour chant de structure composite a plis superposes
US20060060705A1 (en) * 2004-09-23 2006-03-23 Stulc Jeffrey F Splice joints for composite aircraft fuselages and other structures
US20080169380A1 (en) * 2007-01-12 2008-07-17 The Nordam Group, Inc. Composite aircraft window frame
DE102007021076A1 (de) 2007-05-04 2008-11-06 Airbus Deutschland Gmbh Interkostal für ein Luft- oder Raumfahrzeug
WO2009098088A2 (fr) * 2008-02-08 2009-08-13 Airbus Deutschland Gmbh Procédé de fabrication d'un composant composite renforcé par fibres, composant composite renforcé par fibres, et partie de fuselage composite renforcée par fibres d'un avion

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102004009020A1 (de) 2004-02-25 2005-09-15 Airbus Deutschland Gmbh Zwischenkostal für Flugzeuge
FR2869871A1 (fr) * 2004-05-04 2005-11-11 Airbus France Sas Dispositif de protection et revelation de choc pour chant de structure composite a plis superposes
US20060060705A1 (en) * 2004-09-23 2006-03-23 Stulc Jeffrey F Splice joints for composite aircraft fuselages and other structures
US20080169380A1 (en) * 2007-01-12 2008-07-17 The Nordam Group, Inc. Composite aircraft window frame
DE102007021076A1 (de) 2007-05-04 2008-11-06 Airbus Deutschland Gmbh Interkostal für ein Luft- oder Raumfahrzeug
WO2009098088A2 (fr) * 2008-02-08 2009-08-13 Airbus Deutschland Gmbh Procédé de fabrication d'un composant composite renforcé par fibres, composant composite renforcé par fibres, et partie de fuselage composite renforcée par fibres d'un avion

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10450048B2 (en) 2015-11-09 2019-10-22 Airbus Operations Gmbh Aircraft fuselage structure
US10179438B2 (en) 2016-09-23 2019-01-15 Bell Helicopter Textron Inc. Method and assembly for manufacturing door skin and wall with doorway

Also Published As

Publication number Publication date
DE102009056533A1 (de) 2011-06-09
DE102009056533B4 (de) 2015-02-26

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