EP2280817A2 - Structural aircraft panel made of composite material incorporating protection against high-energy impacts - Google Patents

Structural aircraft panel made of composite material incorporating protection against high-energy impacts

Info

Publication number
EP2280817A2
EP2280817A2 EP09738291A EP09738291A EP2280817A2 EP 2280817 A2 EP2280817 A2 EP 2280817A2 EP 09738291 A EP09738291 A EP 09738291A EP 09738291 A EP09738291 A EP 09738291A EP 2280817 A2 EP2280817 A2 EP 2280817A2
Authority
EP
European Patent Office
Prior art keywords
skin
hyper
elastic material
layer
composite
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Withdrawn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
EP09738291A
Other languages
German (de)
French (fr)
Inventor
Nicolas Pechnik
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.)
Airbus Operations SAS
Original Assignee
Airbus Operations SAS
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 Airbus Operations SAS filed Critical Airbus Operations SAS
Publication of EP2280817A2 publication Critical patent/EP2280817A2/en
Withdrawn legal-status Critical Current

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
    • B29C70/04—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
    • 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
    • B29C65/48—Joining or sealing of preformed parts, e.g. welding of plastics materials; Apparatus therefor using adhesives, i.e. using supplementary joining material; solvent bonding
    • 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
    • B29C65/48—Joining or sealing of preformed parts, e.g. welding of plastics materials; Apparatus therefor using adhesives, i.e. using supplementary joining material; solvent bonding
    • B29C65/4805—Joining or sealing of preformed parts, e.g. welding of plastics materials; Apparatus therefor using adhesives, i.e. using supplementary joining material; solvent bonding characterised by the type of adhesives
    • B29C65/483—Reactive adhesives, e.g. chemically curing adhesives
    • 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
    • B29C66/45—Joining of substantially the whole surface of the articles
    • 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
    • B29C66/73—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
    • B29C66/735—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 extensive physical properties of the parts to be joined
    • B29C66/7352—Thickness, e.g. very thin
    • 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
    • B29K2011/00—Use of rubber derived from chloroprene as moulding material
    • 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
    • B29K2063/00—Use of EP, i.e. epoxy resins or derivatives thereof, as moulding material
    • 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
    • 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
    • B29L2009/00—Layered products
    • 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
    • B29L2031/3091—Bicycles
    • 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
    • B29L2031/721—Vibration dampening equipment, e.g. shock absorbers
    • 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
    • Y10T156/00—Adhesive bonding and miscellaneous chemical manufacture
    • Y10T156/10—Methods of surface bonding and/or assembly therefor
    • 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.]
    • Y10T428/24942—Structurally defined web or sheet [e.g., overall dimension, etc.] including components having same physical characteristic in differing degree
    • Y10T428/2495—Thickness [relative or absolute]

Definitions

  • the present invention relates to structural panels of laminated composite material used for the construction of aircraft fuselages. More particularly, the invention relates to an aircraft structural panel subjected to high energy impacts and whose particular structure avoids the intrusion of projectiles impacting beyond a defined distance inside the fuselage. It is known that the use of composite materials makes it possible, for equal mechanical performance, to form lighter structures. This is particularly advantageous in the case of aeronautical structures. In many cases, the composite structures of low or medium thickness such as the skin of an aircraft fuselage, the nacelle panels, the gear box panels, do not contain projectiles having a speed and / or high incident energy.
  • wrinkling can lead to delamination (detachment of folds) that is harmful under impact and special features can appear during dimensioning under static loads and fatigue.
  • shielding protection solutions (combination of several absorbent or resistant materials within a laminate) as a secondary structure, as disclosed in international application WO 2006/070014 are disadvantageous from several points of view:
  • US patent application 2007/095982 discloses an aircraft structural panel made of a composite material with fiber reinforcement and capable of withstanding impacts such as collisions with birds.
  • the skin is made of a composite material specially optimized in its composition to resist shocks and not to break during these impacts but to deform and deflect the trajectory of the impacting body.
  • This solution is effective in the event of impact with a projectile such as a bird that behaves like a viscous fluid and whose impact energy is distributed over a large panel area.
  • the solution is not effective against shocks with debris that generally impact a small area.
  • the invention proposes a structural panel made of a laminated composite material and comprising a face exposed to impacts, further comprising a layer made of a hyper-elastic material reported by collage on his other face.
  • a debris impinging the exposed face of this composite panel will see a part of its energy dissipated by the local rupture of the composite skin, the remaining energy being absorbed by the deformation of the layer of material hyper- elastic that holds the debris and pushes it outward. Thanks to the layer of hyper- elastic, the dissipative power of the composite material can be exploited to sound. maximum.
  • the structural panel comprises a fiber reinforced composite skin in the form of continuous carbon fibers in an epoxy matrix.
  • This type of material has optimum structural resistance characteristics with respect to service demands such as static mechanical stresses or fatigue and thus allows significant gains in mass on the primary structure of the aircraft, in comparison with a primary structure. metallic.
  • this material has no significant plastic deformation capacity capable of dissipating the energy of an impact and prevent the penetration of a projectile by its own deformation.
  • the addition of a layer of hyper-elastic material makes it possible to dimension such a panel with respect to service requirements only, the layer of hyper-elastic material ensuring the absence of penetration of the projectile into the fuselage where it would be likely to degrade systems.
  • the structural panels implemented according to this embodiment are particularly suitable for constituting fuselage structures in areas of the aircraft where protection of the systems by the composite primary structure is necessary and where an analysis of the damage tolerance of said primary structure makes it possible to demonstrate the feasibility of aircraft return after damage. Indeed, a structural panel according to the invention dissipates part of the impact by the damage and the fractures of the composite folds.
  • the primary structure panels concerned have a thickness of between 2 mm and 4 mm of carbon composite - epoxy resin for a continuous fiber volume ratio greater than or equal to 50%.
  • the density of such a panel is of the order of 1500 kg / m 3 .
  • the thickness of the layer of hyper-elastic material is equal to or less than the skin thickness of composite material.
  • the typical density of the hyper-elastic materials with rubber behavior is of the order of 1000 Kg / m 3 . So that the protection according to the invention of the internal systems of the aircraft with respect to the penetration into the fuselage of a projectile is obtained for a mass of material used less than the solution of the prior art of dimension the thickness of the composite material so that the impact can not cause it to break.
  • the structural panel is a panel stiffened by profiles reported on said panel by any means known to those skilled in the art such as co-firing, gluing or riveting, the layer of hyper-elastic material is simply reported between the stiffeners.
  • the layer of hyper-elastic material consists of a polychloropene elastomer such as NEOPRENE® distributed by the company Dupont Chemicals.
  • This material has hyperelastic elongation capacities of the order of 500% and is able to withstand the operating conditions which involve, in the zones considered, temperature variations between -55 0 C and +70 0 C in a humid atmosphere, depending on the phases of the flight, but also chemical attacks such as products such as hydraulic oils or fuel.
  • the layer of hyper-elastic material is preferably reported by gluing. Although the direct vulcanization on the relevant face of the structural composite material panel is possible this solution involves modifying the embodiment of the panels and handling heavier panels during assembly, itself made more complex because of the presence of the layer of hyper-elastic material. It is therefore preferable to report the layer of hyper-elastic material after assembly. Said bonding must be strong so that the penetration of the projectile or debris mainly causes a local hyper-elastic deformation of the layer of rubber material and not a "peeling" of this layer by breaking the adhesive along the surface. composite skin interface - layer of hyper-elastic material. The bonding must also withstand the same environmental conditions as the layer of hyper-elastic material.
  • the structural element exposed to the impacts of an aircraft fuselage is manufactured according to the steps of: fabricating skin panels of composite material assembling said panels to constitute the fuselage element to report the layer of hyper-elastic material on the inner faces of the panels exposed to impacts after assembly.
  • a composite structural panel according to the invention makes it possible to exploit damage by rupture of the skin made of stratified composite material which is a major source of dissipation of energy.
  • a laminated composite skin would be dimensioned so as not to break under the effect of the impact because ensuring only the non penetration of the projectile.
  • a landing gear box roof of carbon-epoxy resin composite material which, according to the prior art, would require a thickness of 6.5 mm in order to withstand the impact of a tire debris
  • a thickness of 3.25 mm corresponding to the thickness necessary for the resumption of operating stress, associated with a layer of hyper-elastic material of chloropolymer type of 3 mm , a gain in mass of the order of 20%.
  • Figure 1 shows schematically the different phases (Figures IA, IB, IC) of the impact of a projectile on the outer face of a structural panel 1 according to the invention.
  • Figure 2 shows schematically the structural features of a panel according to the invention comprising stiffeners.
  • FIG. IA a projectile (1) impacting the composite panel according to the invention (2) on the side of the composite material skin (3).
  • This type of projectile (1) may consist of: debris such as debris of tires whose behavior is flexible, that is to say that the projectile is likely to deform elastically under the effect of the impact , and arrives on the structure with an incident energy of the order of 4000 joules.
  • This type of debris generates a so-called local loading of the structure, the surface of the impact being approximately equal to the largest area of undistorted debris
  • Figure 1B when the projectile (1) encounters the composite skin (3), a large part of the incident energy is dissipated by the mechanisms of damage and rupture of the material on a zone affected by the impact (5). These modes of damage (delamination, fiber breakage, vaporization of the resin and others) dissipate a good part of the incident energy and slow down the projectile.
  • Figure IC the projectile passes through the composite skin and meets the layer of hyper-elastic material (4) which thanks to its hyper-elastic deformation capacity absorbs the remaining kinetic energy of the projectile without breaking and then pushes it outside .
  • the thickness of the layer of hyper-elastic material is chosen so that for a projectile of characteristics given the penetration distance of the projectile remains below a threshold (s) thus ensuring the absence of damage to the systems located behind the structural panel.
  • a threshold s
  • the structural panels have stiffeners (6) attached to the inner face of the skin.
  • the layer of hyper-elastic material (4) is simply reported by gluing (7) between said stiffeners after assembly of the structural panels.

Abstract

The invention provides a structural panel consisting of a stratified composite material comprising one face exposed to impacts and further comprising a layer consisting of a hyper-elastic material bonded adhesively to its other face. In accordance with this embodiment, debris striking the exposed face of this composite panel will have some of its energy dissipated by the local rupture of the composite skin, while the rest of the energy is absorbed by the deformation of the layer of hyper-elastic material which captures the debris and expels it again.

Description

PANNEAU STRUCTURAL D'AERONEF EN MATERIAU COMPOSITE INCORPORANT UNE PROTECTION CONTRE LES IMPACTS A HAUTE STRUCTURAL AIRCRAFT PANEL IN COMPOSITE MATERIAL INCORPORATING PROTECTION AGAINST IMPACTS AT HIGH
ÉNERGIEENERGY
L'invention présente se rapporte aux panneaux structuraux en matériau composite stratifié utilisés pour la construction de fuselages d'aéronefs. Plus particulièrement l'invention se rapporte à un panneau structural d'aéronef soumis à des impacts à haute énergie et dont la structure particulière permet d'éviter l'intrusion des projectiles impactant au delà d'une distance définie à l'intérieur du fuselage. Il est connu que l'utilisation des matériaux composites permet, à performance mécanique égale, de constituer des structures plus légères. Ceci est particulièrement avantageux dans le cas des structures aéronautiques. Dans de nombreux cas, les structures composites de faible ou moyenne épaisseur telle que la peau d'un fuselage d'aéronef, les panneaux de nacelles, les panneaux de case de train, ne permettent pas de contenir des projectiles présentant une vitesse et /ou une énergie incidente élevées. Contrairement aux matériaux métalliques pouvant dissiper l'énergie par déformation plastique, les matériaux composites possèdent un comportement fragile sous impact qui ne permet pas d'utiliser d'une manière robuste le potentiel absorbant du matériau. Il est alors classiquement recommandé d'épaissir fortement le stratifié pour éviter la rupture dans les zones où la structure doit faire office de protection de systèmes avions vitaux. Néanmoins, dans la plupart des cas, cette solution entraîne d'autres difficultés :The present invention relates to structural panels of laminated composite material used for the construction of aircraft fuselages. More particularly, the invention relates to an aircraft structural panel subjected to high energy impacts and whose particular structure avoids the intrusion of projectiles impacting beyond a defined distance inside the fuselage. It is known that the use of composite materials makes it possible, for equal mechanical performance, to form lighter structures. This is particularly advantageous in the case of aeronautical structures. In many cases, the composite structures of low or medium thickness such as the skin of an aircraft fuselage, the nacelle panels, the gear box panels, do not contain projectiles having a speed and / or high incident energy. Unlike metal materials that can dissipate energy by plastic deformation, composite materials have a fragile behavior under impact that does not make it possible to use the absorbent potential of the material in a robust manner. It is then conventionally recommended to strongly thicken the laminate to avoid breaking in the areas where the structure must serve as protection of vital aircraft systems. Nevertheless, in most cases, this solution causes other difficulties:
Introduction de gradients de rigidité locale : l'arrêt de plis peut entraîner des délaminages (décollements de plis) néfastes sous impact et des particularités peuvent apparaître lors du dimensionnement sous charges statiques et de fatigue.Introduction of local rigidity gradients: wrinkling can lead to delamination (detachment of folds) that is harmful under impact and special features can appear during dimensioning under static loads and fatigue.
Introduction d'une épaisseur trop importante faisant apparaître des phénomènes de traction transverse par réflexion d'ondes, entraînant des délaminages pouvant se propager.Introduction of excessive thickness showing transverse tensile phenomena by reflection of waves, resulting in delamination that can spread.
Fabrication plus complexe : traitement des surépaisseurs entraînant des modifications dans le cycle de production.More complex manufacturing: treatment of overthickness resulting in changes in the production cycle.
Augmentation de la masseIncrease in mass
D'autre part, des solutions de protection par blindage (association de plusieurs matériaux absorbants ou résistants au sein d'un stratifié) comme structure secondaire, telle que le divulgue la demande internationale WO 2006/070014 sont désavantageuses selon plusieurs points de vue :On the other hand, shielding protection solutions (combination of several absorbent or resistant materials within a laminate) as a secondary structure, as disclosed in international application WO 2006/070014 are disadvantageous from several points of view:
- Ajout d'une masse supplémentaire non travaillante vis à vis des sollicitations de services - Durabilité des éléments de blindage dans les conditions environnementales subies par l'aéronef Modification du procédé de fabrication et des procédures de contrôle et de maintenance.- Addition of an additional non-working mass with respect to service requests - Durability of the shielding elements in the environmental conditions undergone by the aircraft Modification of the manufacturing process and the control and maintenance procedures.
Il existe donc un besoin pour un panneau structural d'aéronef incorporant une protection contre les impacts à haute énergie. Or les seuls mécanismes de déformation aptes à dissiper de l'énergie dans les matériaux composites sont des modes d' endommagement .There is therefore a need for an aircraft structural panel incorporating high energy impact protection. However, the only deformation mechanisms capable of dissipating energy in composite materials are modes of damage.
La demande de brevet US 2007/095982 décrit un panneau structural d'aéronef constitué d'un matériau composite à renfort fibreux et apte à résister aux impacts tel que des collisions avec des oiseaux. Dans ce cas la peau est réalisée dans un matériau composite spécialement optimisé dans sa composition pour résister aux chocs et ne pas se rompre lors de ces impacts mais se déformer et dévier la trajectoire du corps impactant. Cette solution est efficace en cas de choc avec un projectile tel qu'un oiseau qui se comporte comme un fluide visqueux et dont l'énergie d'impact est répartie sur une surface importante de panneau. La solution n' est pas efficace vis à vis des chocs avec des débris qui d'une manière générale impactent une surface réduite. Afin de résoudre les inconvénients constatés de l'art antérieur, l'invention propose un panneau structural constitué d'un matériau composite stratifié et comprenant une face exposée à des impacts, comprenant en outre une couche constituée d'un matériau hyper- élastique rapportée par collage sur son autre face. Selon ce mode de réalisation, un débris venant impacter la face exposée de ce panneau composite verra une partie de son énergie dissipée par la rupture locale de la peau en composite, l'énergie restante étant absorbée par la déformation de la couche de matériau hyper- élastique qui retient le débris et le repousse vers l'extérieur. Grâce à la couche de matériau hyper- élastique, le pouvoir dissipatif du matériau composite peut être exploité à son. maximum.US patent application 2007/095982 discloses an aircraft structural panel made of a composite material with fiber reinforcement and capable of withstanding impacts such as collisions with birds. In this case the skin is made of a composite material specially optimized in its composition to resist shocks and not to break during these impacts but to deform and deflect the trajectory of the impacting body. This solution is effective in the event of impact with a projectile such as a bird that behaves like a viscous fluid and whose impact energy is distributed over a large panel area. The solution is not effective against shocks with debris that generally impact a small area. In order to overcome the disadvantages noted in the prior art, the invention proposes a structural panel made of a laminated composite material and comprising a face exposed to impacts, further comprising a layer made of a hyper-elastic material reported by collage on his other face. According to this embodiment, a debris impinging the exposed face of this composite panel will see a part of its energy dissipated by the local rupture of the composite skin, the remaining energy being absorbed by the deformation of the layer of material hyper- elastic that holds the debris and pushes it outward. Thanks to the layer of hyper- elastic, the dissipative power of the composite material can be exploited to sound. maximum.
La couche de matériau hyper-élastique étant située en zone interne, c'est-à-dire à l'intérieur du fuselage, l'accès et la contrôlabilité de la santé de la structure primaire au cours de la vie de l'avion sont maintenus. D'autre part la couche de matériau hyper- élastique est protégée de l'environnement extérieur et des agressions physico-chimiques telles que l'exposition aux rayonnements, aux intempéries et aux agents chimiques de nettoyage ou de dégivrage etc ... Selon ce mode de réalisation, le panneau structural comprend une peau en composite .à renfort fibreux sous forme de fibres de carbone continues dans une matrice époxyde. Ce type de matériau présente des caractéristiques de tenue structurale optimales vis à vis des sollicitations de services telles les sollicitations mécaniques statiques ou la fatigue et permet ainsi des gains significatifs de masse sur la structure primaire de l'avion, en comparaison d'une structure primaire métallique. Toutefois ce matériau ne présente pas de capacité de déformation plastique significative apte à dissiper l'énergie d'un impact et empêcher la pénétration d'un projectile par sa propre déformation. L'adjonction d'une couche de matériau hyper-élastique permet de dimensionner un tel panneau vis à vis des sollicitations de service uniquement, la couche de matériau hyper-élastique assurant l'absence de pénétration du projectile dans le fuselage où celui- ci serait susceptible de dégrader des systèmes. Les panneaux structuraux mis en œuvre selon ce mode de réalisation sont particulièrement adaptés pour constituer des structures de fuselage dans des zones de l'aéronef où une protection des systèmes par la structure primaire en composite est nécessaire et où une analyse de la tolérance aux dommages de ladite structure primaire permet de démontrer la faisabilité du retour avion après endommagement . En effet, un panneau structural selon l'invention dissipe une partie de l'impact par les endommagements et les multi- ruptures des plis composites.The layer of hyper-elastic material being located in the internal zone, that is to say inside the fuselage, access and controllability of the health of the primary structure during the life of the aircraft are maintained. On the other hand the layer of hyper-elastic material is protected from the external environment and physicochemical aggressions such as exposure to radiation, weather and chemical cleaning or de-icing agents etc ... According to this mode In one embodiment, the structural panel comprises a fiber reinforced composite skin in the form of continuous carbon fibers in an epoxy matrix. This type of material has optimum structural resistance characteristics with respect to service demands such as static mechanical stresses or fatigue and thus allows significant gains in mass on the primary structure of the aircraft, in comparison with a primary structure. metallic. However, this material has no significant plastic deformation capacity capable of dissipating the energy of an impact and prevent the penetration of a projectile by its own deformation. The addition of a layer of hyper-elastic material makes it possible to dimension such a panel with respect to service requirements only, the layer of hyper-elastic material ensuring the absence of penetration of the projectile into the fuselage where it would be likely to degrade systems. The structural panels implemented according to this embodiment are particularly suitable for constituting fuselage structures in areas of the aircraft where protection of the systems by the composite primary structure is necessary and where an analysis of the damage tolerance of said primary structure makes it possible to demonstrate the feasibility of aircraft return after damage. Indeed, a structural panel according to the invention dissipates part of the impact by the damage and the fractures of the composite folds.
Les panneaux de structure primaire concernés présentent une épaisseur comprise entre 2 mm et 4 mm de composite carbone - résine époxyde pour un taux volumique de fibres continues supérieur ou égal à 50%. La masse volumique d'un tel panneau est de l'ordre de 1500 kg/m3. L'épaisseur de la couche de matériau hyper- élastique est égale ou inférieure à l'épaisseur de peau en matériau composite. La masse volumique typique des matériaux hyper-élastique à comportement caoutchouctique est de l'ordre de 1000 Kg/m3. De sorte que la protection selon l'invention des systèmes internes de l'aéronef vis à vis de la pénétration dans le fuselage d'un projectile est obtenue pour une masse de matériau mis en œuvre inférieure à la solution de l'art antérieur consistant à dimensionner l'épaisseur de matériau composite de telle sorte que l'impact ne puisse provoquer sa rupture. Dans le cas fréquent où le panneau structural est un panneau raidi par des profilés rapportés sur ledit panneau par tous moyens connus de l'homme du métier tels que co-cuisson, collage ou rivetage, la couche de matériau hyper-élastique est simplement rapportée entre les raidisseurs.The primary structure panels concerned have a thickness of between 2 mm and 4 mm of carbon composite - epoxy resin for a continuous fiber volume ratio greater than or equal to 50%. The density of such a panel is of the order of 1500 kg / m 3 . The thickness of the layer of hyper-elastic material is equal to or less than the skin thickness of composite material. The typical density of the hyper-elastic materials with rubber behavior is of the order of 1000 Kg / m 3 . So that the protection according to the invention of the internal systems of the aircraft with respect to the penetration into the fuselage of a projectile is obtained for a mass of material used less than the solution of the prior art of dimension the thickness of the composite material so that the impact can not cause it to break. In the frequent case where the structural panel is a panel stiffened by profiles reported on said panel by any means known to those skilled in the art such as co-firing, gluing or riveting, the layer of hyper-elastic material is simply reported between the stiffeners.
Selon un mode de réalisation avantageux la couche de matériau hyper-élastique est constituée d'un élastomère polychloropène tel que le NEOPRENE® distribué par la société Dupont Chemicals. Ce matériau présente des capacités d'allongement hyper-élastique de l'ordre de 500% et est apte à résister aux conditions de fonctionnement qui impliquent, dans les zones considérées, des variations de température comprise entre -550C et +700C sous une ambiance humide, selon les phases du vol, mais aussi des agressions chimiques telles par des produits tels que des huiles hydrauliques ou du carburant.According to an advantageous embodiment the layer of hyper-elastic material consists of a polychloropene elastomer such as NEOPRENE® distributed by the company Dupont Chemicals. This material has hyperelastic elongation capacities of the order of 500% and is able to withstand the operating conditions which involve, in the zones considered, temperature variations between -55 0 C and +70 0 C in a humid atmosphere, depending on the phases of the flight, but also chemical attacks such as products such as hydraulic oils or fuel.
La couche de matériau hyper-élastique est rapportée préférentiellement par collage. Bien que la vulcanisation directe sur la face concernée du panneau structural en matériau composite soit possible cette solution implique de modifier le mode de réalisation des panneaux et de manipuler des panneaux plus lourds durant l'assemblage, lui même rendu plus complexe du fait de la présence de la couche de matériau hyper- élastique. Il est donc préférable de rapporter la couche de matériau hyper-élastique après assemblage. Ledit collage doit être fort de manière à ce que la pénétration du projectile ou des débris provoque majoritairement une déformation hyper-élastique locale de la couche de matériau caoutchoutique et non un « pelage » de cette couche par la rupture de la colle le long de l'interface peau composite - couche de matériau hyper-élastique. Le collage doit en outre résister aux mêmes conditions environnementales que la couche de matériau hyper-élastique. Une colle de type époxyde répond à ces exigences . Selon un procédé de mise en œuvre avantageux de l'invention, l'élément structural exposé aux impacts d'un fuselage d'aéronef est fabriqué selon les étapes consistant à : fabriquer des panneaux de peau en matériau composite assembler lesdits panneaux afin de constituer l'élément de fuselage rapporter la couche de matériau hyper- élastique sur les faces internes des panneaux exposés aux impacts après l'assemblage.The layer of hyper-elastic material is preferably reported by gluing. Although the direct vulcanization on the relevant face of the structural composite material panel is possible this solution involves modifying the embodiment of the panels and handling heavier panels during assembly, itself made more complex because of the presence of the layer of hyper-elastic material. It is therefore preferable to report the layer of hyper-elastic material after assembly. Said bonding must be strong so that the penetration of the projectile or debris mainly causes a local hyper-elastic deformation of the layer of rubber material and not a "peeling" of this layer by breaking the adhesive along the surface. composite skin interface - layer of hyper-elastic material. The bonding must also withstand the same environmental conditions as the layer of hyper-elastic material. An epoxy type of glue meets these requirements. According to an advantageous implementation method of the invention, the structural element exposed to the impacts of an aircraft fuselage is manufactured according to the steps of: fabricating skin panels of composite material assembling said panels to constitute the fuselage element to report the layer of hyper-elastic material on the inner faces of the panels exposed to impacts after assembly.
Lorsque le panneau structural est de type raidi ce procédé de fabrication intègre en outre une étape de montage et de fixation des raidisseurs sur la peau après la première étape. Ainsi la mise en œuvre de la couche de protection est locale et n'entre pas dans le processus de fabrication de la structure primaire, les procédures d'assemblage ne sont pas modifiées. Avantageusement un panneau composite structural selon l'invention permet d'exploiter l' endommagement par rupture de la peau en matériau composite stratifié qui est une source majeure de dissipation de l'énergie. Selon l'art antérieur une telle peau en composite stratifiée serait dimensionnée pour ne pas rompre sous l'effet de l'impact car assurant seule la non pénétration du projectile. A titre d' exemple non limitatif, un toit de case de train d'atterrissage en matériau composite carbone- résine époxyde, qui, selon l'art antérieur, nécessiterait une épaisseur de 6,5 mm afin de résister à l'impact d'un débris de pneumatique, peut être réalisé selon l'invention avec une épaisseur de 3,25 mm, correspondant à l'épaisseur nécessaire à la reprise des sollicitations de service, associée à une couche de matériau hyper-élastique de type chloropolymère de 3 mm, soit un gain de masse de l'ordre de 20%.When the structural panel is stiff type this manufacturing method also incorporates a step of mounting and fixing the stiffeners on the skin after the first step. Thus the implementation of the protective layer is local and does not enter the manufacturing process of the primary structure, assembly procedures are not changed. Advantageously, a composite structural panel according to the invention makes it possible to exploit damage by rupture of the skin made of stratified composite material which is a major source of dissipation of energy. According to the prior art, such a laminated composite skin would be dimensioned so as not to break under the effect of the impact because ensuring only the non penetration of the projectile. By way of non-limiting example, a landing gear box roof of carbon-epoxy resin composite material, which, according to the prior art, would require a thickness of 6.5 mm in order to withstand the impact of a tire debris may be produced according to the invention with a thickness of 3.25 mm, corresponding to the thickness necessary for the resumption of operating stress, associated with a layer of hyper-elastic material of chloropolymer type of 3 mm , a gain in mass of the order of 20%.
La figure 1 représente schématiquement les différentes phases (figures IA, IB, IC) de l'impact d'un projectile sur la face externe d'un panneau structural selon 1' invention. La figure 2 représente schématiquement les caractéristiques structurelles d' un panneau selon l'invention comportant des raidisseurs.Figure 1 shows schematically the different phases (Figures IA, IB, IC) of the impact of a projectile on the outer face of a structural panel 1 according to the invention. Figure 2 shows schematically the structural features of a panel according to the invention comprising stiffeners.
Figure IA, un projectile (1) venant impacter le panneau composite selon l'invention (2) du côté de la peau en matériau composite (3) . Ce type de projectile (1) peut être constitué par : des débris tels que des débris de pneumatiques dont le comportement est souple, c'est-à-dire que le projectile est susceptible de se déformer élastiquement sous l'effet de l'impact, et arrive sur la structure avec une énergie incidente de l'ordre de 4000 joules. Ce type de débris engendre une sollicitation dite locale de la structure, la surface de l'impact étant approximativement égale à la plus grande surface du débris non déforméFigure IA, a projectile (1) impacting the composite panel according to the invention (2) on the side of the composite material skin (3). This type of projectile (1) may consist of: debris such as debris of tires whose behavior is flexible, that is to say that the projectile is likely to deform elastically under the effect of the impact , and arrives on the structure with an incident energy of the order of 4000 joules. This type of debris generates a so-called local loading of the structure, the surface of the impact being approximately equal to the largest area of undistorted debris
- de petits débris émanant des moteurs. Ce type de débris présente un comportement très rigide et engendrent une sollicitation très locale de la structure, la zone d' impact étant approximativement égale à la surface de contact entre le projectile et le panneau. L'énergie incidente est comprise entre 1000 et 4000 joules- small debris emanating from the engines. This type of debris exhibits a very rigid behavior and generates a very local stress on the structure, the impact zone being approximately equal to the contact surface between the projectile and the panel. The incident energy is between 1000 and 4000 joules
- un oiseau, le comportement de ce projectile est assimilable à celui d'un fluide visqueux. Celui-ci va, au moment de l'impact, « s'écouler » sur la structure, engendrant ainsi un effort réparti. L'énergie incidente est de l'ordre de 30000 joules- a bird, the behavior of this projectile is comparable to that of a viscous fluid. It will, at the moment of impact, "flow" on the structure, thus generating a distributed effort. The incident energy is of the order of 30000 joules
Figure IB, lorsque le projectile (1) rencontre la peau composite (3), une part importante de l'énergie incidente est dissipée par les mécanismes d' endommageπient et de rupture du matériau sur une zone affectée par l'impact (5) . Ces modes d' endommagement (délaminage, rupture de fibres, vaporisation de la résine et autres) dissipent une bonne partie de l'énergie incidente et ralentissent le projectile. Figure IC, le projectile traverse la peau composite et rencontre la couche de matériau hyper-élastique (4) qui grâce à sa capacité de déformation hyper-élastique absorbe l'énergie cinétique restante du projectile sans rompre puis repousse celui-ci à l'extérieur. L'épaisseur de la couche de matériau hyper-élastique est choisie de telle sorte que pour un projectile de caractéristiques données la distance de pénétration du projectile reste inférieure à un seuil (s) garantissant ainsi l'absence d' endommagement des systèmes situés derrière le panneau structural. Figure 2, dans la majorité des cas les panneaux structuraux comportent des raidisseurs (6) rapportés sur la face interne de la peau. Dans ce cas la couche de matériau hyper-élastique (4) est simplement rapportée par collage (7) entre lesdits raidisseurs après assemblage des panneaux structuraux. Figure 1B, when the projectile (1) encounters the composite skin (3), a large part of the incident energy is dissipated by the mechanisms of damage and rupture of the material on a zone affected by the impact (5). These modes of damage (delamination, fiber breakage, vaporization of the resin and others) dissipate a good part of the incident energy and slow down the projectile. Figure IC, the projectile passes through the composite skin and meets the layer of hyper-elastic material (4) which thanks to its hyper-elastic deformation capacity absorbs the remaining kinetic energy of the projectile without breaking and then pushes it outside . The thickness of the layer of hyper-elastic material is chosen so that for a projectile of characteristics given the penetration distance of the projectile remains below a threshold (s) thus ensuring the absence of damage to the systems located behind the structural panel. Figure 2, in most cases the structural panels have stiffeners (6) attached to the inner face of the skin. In this case the layer of hyper-elastic material (4) is simply reported by gluing (7) between said stiffeners after assembly of the structural panels.

Claims

REVENDICATIONS
1. Panneau structural (2) apte à constituer une structure résistant aux sollicitations statiques et de fatigue de service qui lui sont imposées caractérisé en ce qu' il comprend une peau composite à renfort fibreux sous forme de fibres continues, une face de ladite peau étant exposée à des impacts (3), et une couche constituée d'un matériau hyper- élastique (3) d'épaisseur égale ou inférieure à l'épaisseur de ladite peau et rapportée par collage sur son autre face.1. structural panel (2) capable of constituting a structure resistant to static stresses and service fatigue imposed on it characterized in that it comprises a composite skin with fibrous reinforcement in the form of continuous fibers, a face of said skin being exposed to impacts (3), and a layer made of a hyper-elastic material (3) of thickness equal to or less than the thickness of said skin and reported by gluing on its other side.
2. Panneau selon la revendication 1 caractérisé en ce que la peau est constituée d'un composite à renfort fibreux sous forme de fibres de carbone continues dans une matrice époxyde. 2. Panel according to claim 1 characterized in that the skin is made of a fibrous reinforcement composite in the form of continuous carbon fibers in an epoxy matrix.
3. Panneau selon la revendication 2 caractérisé en ce que son épaisseur de peau est comprise entre 2 et 4 mm.3. Panel according to claim 2 characterized in that its skin thickness is between 2 and 4 mm.
4. Panneau selon la revendication 3 caractérisé en ce qu'il comprend des raidisseurs (6) et que la couche de matériau hyper-élastique (4) est collée sur la peau entre lesdits raidisseurs.4. Panel according to claim 3 characterized in that it comprises stiffeners (6) and that the layer of hyper-elastic material (4) is bonded to the skin between said stiffeners.
5. Panneau selon l'une quelconque des revendications précédentes caractérisé en ce que la couche de matériau hyper-élastique (4) est constituée d'un élastomère polychloroprène.5. Panel according to any one of the preceding claims characterized in that the layer of hyper-elastic material (4) consists of a polychloroprene elastomer.
6. Panneau selon l'une quelconque des revendications précédentes caractérisé en ce que le collage est réalisé par une colle époxyde (7) . 6. Panel according to any one of the preceding claims, characterized in that the bonding is performed by an epoxy adhesive (7).
7. Procédé de fabrication d' un élément structural exposé aux impacts (2) d'un fuselage d'aéronef selon la revendication 1 comprenant les étapes consistant à : - fabriquer des panneaux de peau en matériau composite - assembler lesdits panneaux afin de constituer l'élément de fuselage caractérisé en ce qu' il comprend une étape consistant à rapporter une couche de matériau hyper- élastique (3) sur les faces internes des panneaux exposés aux impacts après l'assemblage.7. A method of manufacturing a structural element exposed to the impacts (2) of an aircraft fuselage according to claim 1 comprising the steps of: - manufacturing skin panels of composite material - assembling said panels to form the fuselage element characterized in that it comprises a step of reporting a layer of hyper-elastic material (3) on the inner faces of the panels exposed to impacts after assembly.
8. Procédé selon la revendication 7 caractérisé en ce qu' il comprend en outre une étape consistant à rapporter des raidisseurs (6) par co-cuisson ou par collage sur au moins un des panneaux de peau avant 1' assemblage et que la couche de matériau hyper- élastique (3) est rapportée entre les raidisseurs après l'assemblage des panneaux. 8. A method according to claim 7, characterized in that it further comprises a step of bringing stiffeners (6) by co-cooking or gluing on at least one of the skin panels before assembly and that the hyper-elastic material (3) is reported between the stiffeners after assembly of the panels.
EP09738291A 2008-03-28 2009-03-26 Structural aircraft panel made of composite material incorporating protection against high-energy impacts Withdrawn EP2280817A2 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0801686A FR2929169A1 (en) 2008-03-28 2008-03-28 STRUCTURAL AIRCRAFT PANEL IN COMPOSITE MATERIAL INCORPORATING PROTECTION AGAINST HIGH ENERGY IMPACTS
PCT/FR2009/000333 WO2009133257A2 (en) 2008-03-28 2009-03-26 Structural aircraft panel made of composite material incorporating protection against high-energy impacts

Publications (1)

Publication Number Publication Date
EP2280817A2 true EP2280817A2 (en) 2011-02-09

Family

ID=40091844

Family Applications (1)

Application Number Title Priority Date Filing Date
EP09738291A Withdrawn EP2280817A2 (en) 2008-03-28 2009-03-26 Structural aircraft panel made of composite material incorporating protection against high-energy impacts

Country Status (4)

Country Link
US (1) US8906493B2 (en)
EP (1) EP2280817A2 (en)
FR (1) FR2929169A1 (en)
WO (1) WO2009133257A2 (en)

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11014690B2 (en) 2014-10-08 2021-05-25 Textron Innovations, Inc. Adjustable interior mockup
US10894616B2 (en) * 2014-10-08 2021-01-19 Textron Innovations, Inc. Adjustable interior mockup
RU2637001C1 (en) * 2016-12-14 2017-11-29 Федеральное государственное унитарное предприятие "Центральный аэрогидродинамический институт имени профессора Н.Е. Жуковского" (ФГУП "ЦАГИ") Hybrid composite panel for airframe
US11338903B2 (en) * 2019-10-21 2022-05-24 The Boeing Company Integrally stiffened bonded panel with vented pockets and methods of manufacture
US11554561B2 (en) 2019-10-21 2023-01-17 The Boeing Company Integrally stiffened bonded panel with machined pockets and methods of manufacture
US11724823B2 (en) * 2019-10-21 2023-08-15 The Boeing Company Integrally stiffened bonded panel with machined recesses and improved methods of manufacture
US11685509B2 (en) 2019-10-21 2023-06-27 The Boeing Company Integrally stiffened bonded panel with vented pockets and methods of manufacture

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3160549A (en) * 1960-12-29 1964-12-08 Minnesota Mining & Mfg Vibration damping structures
US3525703A (en) * 1966-08-22 1970-08-25 Bridgestone Tire Co Ltd Method of adhering a synthetic fibre to a rubber,an adhesive solution and a laminate obtained by improving a method of adhering a synthetic fibre to a rubber
US3472730A (en) * 1967-12-28 1969-10-14 Minnesota Mining & Mfg Heat-curable filament-reinforced resinous sheeting and laminating process using same
US3616185A (en) * 1970-06-17 1971-10-26 Samuel Goldberg Glass fiber articles with a protective layer of polyurethane and method of preparation
AU565384B2 (en) * 1982-09-13 1987-09-17 James Hardie Building Products Pty. Limited Lining polymers with elastomers
US5028056A (en) * 1986-11-24 1991-07-02 The Gates Rubber Company Fiber composite sealing element
JP4318381B2 (en) * 2000-04-27 2009-08-19 本田技研工業株式会社 Manufacturing method of fuselage structure made of fiber reinforced composite material, and fuselage structure manufactured thereby
DE60217542T2 (en) * 2001-11-07 2007-10-25 Toray Industries, Inc. EPOXY RESIN COMPOSITIONS FOR FIBER-REINFORCED COMPOSITE MATERIALS, METHOD FOR PRODUCING MATERIALS AND FIBER-REINFORCED COMPOSITE MATERIALS
ES2279663B1 (en) * 2004-12-30 2008-08-01 Airbus España, S.L. AIRCRAFT WITH REAR FUSELAGE PROTECTION SHIELD.
ITVE20050020A1 (en) * 2005-04-07 2006-10-08 Isolgomma Srl MULTILAYER INSULATING ACOUSTIC PRODUCT FOR BUILDING, METHOD FOR ITS PRODUCTION AND PLANT TO IMPLEMENT THE METHOD
US9359061B2 (en) * 2005-10-31 2016-06-07 The Boeing Company Compliant stiffener for aircraft fuselage
DE102006008423A1 (en) 2006-02-23 2007-08-30 Wilo Ag Motorized centrifugal pump for pumping substances has a stack of contacts for a stator on an electric motor extrusion- coated with plastic fitted with cooling channels

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
None *
See also references of WO2009133257A2 *

Also Published As

Publication number Publication date
WO2009133257A3 (en) 2010-01-14
FR2929169A1 (en) 2009-10-02
US20120040159A1 (en) 2012-02-16
US8906493B2 (en) 2014-12-09
WO2009133257A2 (en) 2009-11-05

Similar Documents

Publication Publication Date Title
EP2280817A2 (en) Structural aircraft panel made of composite material incorporating protection against high-energy impacts
US7738236B2 (en) Light weight system for lightning protection of nonconductive aircraft panels
EP2938537B1 (en) Energy absorbtion device for structural element of an aircraft
CA2895547C (en) Energy absorption device for aircraft structural element
FR2924687A1 (en) AIRCRAFT AND AIRCRAFT SYSTEM COMPRISING SUCH A SYSTEM
FR2859496A1 (en) REINFORCED FAN BLADE AND METHOD OF MANUFACTURING SUCH BLADE
WO2009080990A1 (en) Anti-lightning system and aircraft comprising such a system
EP2143637B1 (en) Method for mounting a de-icing mat and a metal shield on a structure
EP2268482B1 (en) Damping skin for protecting composite parts
FR3075759A1 (en) ANTERIOR PLATFORM PART OF A PROPELLANT AIRCRAFT COMPRISING A SHOCKING ELEMENT
US10611478B2 (en) Structural panels for exposed surfaces
FR2869565A1 (en) PROTECTING A COMPOSITE STRUCTURE WITH IMPACTS
EP2596255A1 (en) Aircraft comprising a link rod one part of which is made of composite
WO2010136494A1 (en) Device for temporarily connecting and pyrotechnically separating two assemblies
EP0741856A1 (en) Anti-fragmentation covering for an armoured vehicle
EP2454154B1 (en) Leading edge structure, in particular for the air intake of the nacelle of an aircraft engine
US20130284339A1 (en) Ballistic Protection Systems and Methods
FR2918036A1 (en) Leading edge forming device for e.g. wing of aircraft, has rigid elements, external and internal walls forming metallic lattice, where rigid elements are extended between walls and connected directly/indirectly to walls
EP2382088B1 (en) Structure made of composite material protected against the effects of lightning
FR2939406A1 (en) Soundproofing trim panel for helicopter type rotorcraft, has core that is melamine self-extinguishing foam in which inserts are arranged, and skins provided with carbon cloth layers respectively
EP2560879B1 (en) Optimized thermal protection material
FR2632393A1 (en) Composite armour, particularly for protection against projectiles with hollow charges
FR2798189A1 (en) Armor for protection against bullets and shrapnel has multilayer steel shell adhered to layer of fibre-reinforced resin
FR2898105A1 (en) Cargo aircraft crew and equipment protecting shield, has plates opposing effects of ballistic missile and presenting face in opposite to surface to be protected and another face oriented towards aerodynamic flow
FR3103137A1 (en) Laminated glazing with improved resistance to relatively punctual types of impact

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 20101022

AK Designated contracting states

Kind code of ref document: A2

Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO SE SI SK TR

AX Request for extension of the european patent

Extension state: AL BA RS

DAX Request for extension of the european patent (deleted)
17Q First examination report despatched

Effective date: 20170208

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION IS DEEMED TO BE WITHDRAWN

18D Application deemed to be withdrawn

Effective date: 20170620