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EP2076430A2 - Aircraft fuselage made from longitudinal panels and method of producing such a fuselage - Google Patents

Aircraft fuselage made from longitudinal panels and method of producing such a fuselage

Info

Publication number
EP2076430A2
EP2076430A2 EP07858512A EP07858512A EP2076430A2 EP 2076430 A2 EP2076430 A2 EP 2076430A2 EP 07858512 A EP07858512 A EP 07858512A EP 07858512 A EP07858512 A EP 07858512A EP 2076430 A2 EP2076430 A2 EP 2076430A2
Authority
EP
European Patent Office
Prior art keywords
panels
longitudinal
panel
aircraft
fuselage
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Withdrawn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
EP07858512A
Other languages
German (de)
French (fr)
Inventor
Guillaume Gallant
Philippe Bernadet
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 EP2076430A2 publication Critical patent/EP2076430A2/en
Withdrawn legal-status Critical Current

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
    • B64C1/068—Fuselage sections
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
    • B64C1/068—Fuselage sections
    • B64C1/0683—Nose cones
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
    • B64C1/068—Fuselage sections
    • B64C1/0685—Tail cones
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C2211/00—Modular constructions of airplanes or helicopters

Definitions

  • the invention relates to an aircraft fuselage whose central section is made by means of longitudinal panels assembled directly with each other, without circumferential junctions. It also relates to a method for producing such a fuselage.
  • the invention has applications in the field of aeronautics and, in particular, in the field of manufacture of aircraft fuselages. State of the art
  • the fuselage of aircraft is generally made from several metal panels assembled with each other. These metal panels are often aluminum panels. They are fixed with each other by means of circumferential and longitudinal junctions. These panels are assembled to form fuselage sections, themselves assembled to form the fuselage. As represented in FIG. 1, an aircraft fuselage comprises several sections, namely:
  • the rear section 2, or rear cone generally comprising the empennage
  • the central section 3 which is the part of the fuselage connecting the front tip 1 with the rear section 2.
  • the central section 3 generally comprises the passenger cabin and baggage compartments, for passenger transport, or cargo bunkers, for a freight transport.
  • the central section is made from several section elements, or parts of sections, assembled with each other to form the central section connecting the front tip with the rear cone of the aircraft.
  • each part of the central section 3, for example parts 3a, 3b, 3c and 3d of the fuselage shown in FIG. 1 is manufactured separately from several metal panels.
  • Figure 2 there is shown an example of two parts of a central section 3 as conventionally realized. Each of these parts 3a and 3b comprises several metal panels assembled with each other.
  • the section portion 3a comprises the panels 41 to 46 and the portion of section 3b comprises the panels 47 to 52.
  • the different panels of the same portion of section are assembled by means of longitudinal junctions.
  • longitudinal junction refers to a type of fastening that involves placing the panels so that two consecutive panels overlap partially and insert fasteners, such as rivets, into the areas where the panels overlap.
  • central section portions are then assembled together.
  • Two consecutive central section portions are assembled by means of circumferential junctions.
  • the term "circumferential junction” is a fastening means for assembling portions of sections along the entire circumference of these parts. Indeed, two parts of sections can not be assembled directly with each other (by overlapping) for reasons of tolerance, because it is impossible to make two parts of section so that they fit perfectly one in the other.
  • to assemble two portions of section is generally placed a shell between the two portions of section.
  • a ferrule is a local skin, internal to the section, which provides reinforcement to the junction zone of the two portions of section. It makes it possible to transfer the efforts from one section to another. This shell is fixed, on both sides, on each part of the section.
  • an additional skin (made from one or more pieces of skin mounted on a frame) is placed at the joint of two portions of section and fixed by fastening elements, such as splints , on each of the section parts.
  • the composite materials have the advantage of being relatively light compared to the metal, which allows to significantly reduce the total mass of an aircraft.
  • the aircraft manufacturers generally seek to achieve the central section of composite material. They then try to reproduce a single skin that would encompass the central section over its entire circumference, that is to say 360 °. In other words, they try to make a central section of a single piece. Since it is difficult to make such a central section, it has been planned to make one-piece portions of section to be assembled with each other to form a central section. Each section portion is therefore a cylinder which must then be assembled with the consecutive section portions. This assembly is made by means of circumferential junctions, as described above, and for which the additional skins are made of composite material.
  • the circumferential junctions are expensive in mass. In addition, they require a relatively high assembly time, especially as they present assembly difficulties because the aerodynamic profile requires a perfect fit of the different portions of the section and an adjustment of the docking of the sub-section. structure, when substructure elements are mounted in the section portions. In addition, in addition to the circumferential junctions, many pieces of interfaces are required to secure the additional skin on the two portion portions. Moreover, given the current division of the central section into different parts of sections, an aircraft fuselage is transported, from one site to another, portion of section per portion of section. In other words, each section part is transported on a specific transport vehicle. Presentation of the invention
  • the purpose of the invention is precisely to overcome the disadvantages of the techniques described above.
  • the invention provides an aircraft fuselage whose central section is made from longitudinal panels assembled directly with each other, that is to say without additional skins.
  • the central section does not require the use of circumferential junctions throughout its circumference.
  • the invention thus provides a gain in the total mass of the fuselage.
  • the assembly of the panels by longitudinal junctions is simpler than by circumferential junctions, which allows a faster implementation.
  • the invention relates to an aircraft fuselage comprising a nose, a rear section and a central section, characterized in that the central section comprises longitudinal panels assembled directly with each other, at least one of these longitudinal panels having a length corresponding to the distance between the front tip and the rear section to connect said front tip with said rear section.
  • the invention may also include one or more of the following features:
  • the longitudinal panels are made of composite materials.
  • a longitudinal panel has a length corresponding to the distance between the front tip and a central wing box of the aircraft.
  • a longitudinal panel has a length corresponding to the distance between a central wing box and the rear section.
  • a longitudinal panel has a length corresponding to a width of a central wing box.
  • a longitudinal panel incorporates a stiffening substructure or a floor structure.
  • a longitudinal panel incorporates skin reinforcements.
  • a longitudinal panel corresponds to a fuselage single curvature zone.
  • a longitudinal panel corresponds to a zone of the fuselage with double curvature.
  • the invention also relates to a method of producing such a fuselage.
  • This method comprises the manufacture of a nose, a rear section and a central section.
  • the manufacture of the central section is characterized by the following steps:
  • the method of the invention may also include one or more of the following features:
  • the longitudinal panels are made of composite materials.
  • the direct assembly of a first panel with a second panel consists of partially superimposing the first and second panels and to fix the two panels by means of fasteners.
  • the invention also relates to an aircraft comprising a fuselage as described above. It also relates to an aircraft comprising a fuselage made according to the method as described above.
  • FIG. 1, already described, represents an example of aircraft fuselage of the prior art composed of several sections.
  • FIG. 1 already described, shows an example of central section portions according to the prior art.
  • FIG. 3 represents an example of a central section according to the invention.
  • FIG. 4 represents another example of a central section according to the invention.
  • FIG. 5 represents an example of a central section with a single curvature and a rear section with double curvature, according to the invention.
  • the invention proposes to make the central section of an aircraft from longitudinal panels, that is to say, long panels each made of a single piece.
  • the dimension of each of these longitudinal panels corresponds, at the most, to the distance between the nose of the aircraft and the rear section of the aircraft and, at a minimum, the distance between the nose of the aircraft and the central wing box or between the rear section of the aircraft and the central box of sail.
  • other panels can be mounted between panels of great length, such as, for example, a panel whose length corresponds to the width of the central wing box.
  • At least one longitudinal panel connects the front tip of the aircraft with the rear section of said aircraft.
  • the longitudinal panels can be assembled by means of longitudinal junctions, that is to say by partially superimposing a panel on another panel and fixing the two panels one on the other by means of elements of fixings passing through the two panels in their overlapping zone or by juxtaposing two panels with an internal doubler (ferrule) to ensure continuity.
  • the panels can be made of composite materials.
  • the composite materials allow the realization of large parts, in one piece. These parts may be panels whose size is chosen according to the aircraft to be built and not according to the technical difficulties of producing said part.
  • the fasteners are known elements, suitable for fixing these materials.
  • FIG. 3 shows an example of a central fuselage section produced according to the invention.
  • the central section 3 comprises five longitudinal panels 31 to 35, assembled directly with each other.
  • two panels 31 and 32 are assembled to form the upper part of the fuselage. These two panels have a length between the front tip and the rear cone of the aircraft.
  • the panels 33, 34 and 35 are each fixed, on the one hand, on the panel 31 and, on the other hand, on the panel 32.
  • the panel 33 has a length corresponding to the distance between the front tip of the aircraft and the box central wing.
  • the panel 34 has a length corresponding to the distance between the rear section of the aircraft and the central wing box.
  • the panel 35 has a length corresponding to the width of the central wing box.
  • the assembly is carried out longitudinally, that is to say along the length of the panels.
  • the panels are fixed to each other along the longitudinal axis XX of the aircraft. Only some panels require a partial circumferential assembly, in addition to the longitudinal assembly.
  • the panel 35 must be fixed not only on the panels 31 and 32 with a longitudinal type assembly, but also on the panels 33 and 34.
  • the assembly of the panel 35 with the panels 33 and 34 is an assembly of partially circumferential or semi-circumferential type. When the assembly is not circumferential over the entire circumference of the fuselage, play is possible between the two panels to be assembled.
  • the panels forming the central section are assembled only by longitudinal junctions. No circumferential junction is necessary except at the front and rear ends of the central section to secure said central section with the nose and the rear cone of the aircraft.
  • the total mass of the central section is thus lightened by the mass corresponding to the circumferential junctions and other interface parts.
  • the length of the various longitudinal panels is adapted to the shape of the central section of the aircraft. It can be adapted also to the mode of transport of these longitudinal panels. Indeed, the longitudinal panels can be transported in a simplified manner, with respect to a central section portion of the prior art, because they can be placed one above the other in the transport vehicle. For example, the panels 31 and 32 of FIG. 3 can be placed one inside the other at the bottom of the vehicle and the panels 34, 33 and 35 placed, one inside the other, above the panels 31 and 32. Indeed, the transport of fuselage elements in the form of longitudinal panels makes it possible to better occupy the load volume made available. We can thus transport several central sections, broken down into long panels, in a volume where we can could only carry one central section, if it was cut into sections.
  • the assembly of the various longitudinal panels by overlapping said panels is simplified, compared to the prior art, since the number of fasteners is decreased by more than half.
  • the longitudinal junctions are more tolerant than the circumferential junctions; indeed, the circumferential assembly of two portions of section implies that the dimensions of these two portions of section are almost identical to allow continuity in the fuselage while this constraint does not exist with a longitudinal assembly.
  • the longitudinal panels may comprise openings and substructures of the aircraft.
  • the panels 31 and 32 comprise openings corresponding to the locations of the windows 5 and the passenger doors 4 and 6 of the aircraft.
  • the panels 33 and 34 comprise openings 7 corresponding to the doors of the landing hatch boxes and the bunker doors.
  • these techniques make it possible to insert one or more elements, made of composite materials or other materials, into a panel made of composite materials.
  • Longitudinal panels can therefore integrate a stiffening substructure of the aircraft, such as rails, frames, door frames, window frames, or even structural primers or complete structures of floors.
  • These different panels can also incorporate skin reinforcements such as those located at the openings or close to the heavily loaded areas of the aircraft.
  • the part of the fuselage containing the central wing box (corresponding to the panel 35 in FIG. 3) is made in the longitudinal panels 31, 32 or in the longitudinal panels. 33, 34.
  • the central section 3 can then comprise only four longitudinal panels. In these examples, only one partial circumferential assembly may be sufficient, see no partial circumferential assembly. All the joints ensuring the assembly of the different panels between them can be longitudinal junctions.
  • the shape of the longitudinal panels and their number may vary according to different criteria, such as the type of aircraft to be built and the transport provided for these panels.
  • the length of the panels may vary to include or not certain elements of the aircraft.
  • the length of the panels 31, 32 and 34, in Figure 3 may vary to include or not the opening corresponding to the rear door 4.
  • Figure 4 there is shown an example of central section realized from five panels and not including the opening of the rear door 4.
  • a panel 37 forms the roof of the central section
  • two panels 36 and 38, symmetrical on either side of the panel 37, comprise the openings of the windows 5
  • a panel 33 has a cargo door 7a
  • a panel 39 has a cargo door 7b and the central box 8.
  • the longitudinal panels are intended to form single curvature fuselage areas.
  • single curvature zone is meant a zone of the fuselage whose radius of curvature is identical over the entire length of the zone.
  • double curvature an area of the fuselage whose radius of curvature differs along the length of the zone.
  • the rear section of an aircraft has a cone shape.
  • This cone zone typically, is a double curvature zone.
  • the central part of the central section which can be cylindrical, bilobed, trilobed, ..., constitutes a single curvature zone. From an industrial point of view, it is simpler to produce single curvature panels, especially composite materials.
  • single curvature panels can be made with a flat surface and shaped at the time of cooking or draper with special automated draping machines because single curvature areas are developable, unlike double curvature areas.
  • double curvature panels must be made using relatively complex machines, such as fiber placement machines to achieve a curvature of the panel that is consistent with the shape of the central section. desired.
  • the manufacture of single curvature panels therefore requires less expensive tools than the manufacture of double curvature panels.
  • the assembly of single curvature panels can be done by sliding the panels on each other while the assembly of double curvature panels requires a specific cavage panels between them, which involves problems of hyperstaticity .
  • the invention proposes an embodiment in which the longitudinal panels are only applied in fuselage zone with simple curvature. It is thus possible to reduce the manufacturing cost of these panels and to reduce the constraints of positioning the panels together.
  • this embodiment proposes to realize the double curvature fuselage area independently of the single curvature central section.
  • the fuselage portion shown in Figure 5 is a double curvature zone. In the center of the central section 3, the radius of curvature is greater than that of the section 9. To avoid the production of longitudinal panels for double-curvature zones, it may be chosen to make the section 9 independently. The section 9 is then considered to form the rear section of the aircraft.
  • the longitudinal panels 36, 37, 38, 33, 35 and 34 are thus made to be all applied only in single curvature fuselage areas.
  • the section 9 can be made in a conventional manner, that is to say in the form of a 360 ° section attached to the rest of the central section by a conventional circumferential junction.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Automatic Assembly (AREA)
  • Moulding By Coating Moulds (AREA)
  • Connection Of Plates (AREA)

Abstract

The invention relates to an aircraft fuselage comprising a nose cone (2), a tail section (2) and a central section (3), the central section comprising longitudinal panels (31 - 39) assembled directly with one another, at least one of these longitudinal panels having a length corresponding to the distance between the nose cone and the tail section in order to connect said nose cone with said tail section.

Description

FUSELAGE D'AERONEF REALISE A PARTIR DE PANNEAUX AIRCRAFT FUSELAGE MADE FROM PANELS
LONGITUDINAUX ET PROCEDE DE REALISATIONLONGITUDINALS AND METHOD FOR PRODUCING THE SAME
D'UN TEL FUSELAGEOF SUCH A FUSELAGE
Domaine de l'inventionField of the invention
L'invention concerne un fuselage d'aéronef dont le tronçon central est réalisé au moyen de panneaux longitudinaux assemblés directement les uns avec les autres, sans jonctions circonférentielles. Elle concerne également un procédé pour réaliser un tel fuselage. L'invention trouve des applications dans le domaine de l'aéronautique et, en particulier, dans le domaine de la fabrication de fuselages d'aéronefs. Etat de la techniqueThe invention relates to an aircraft fuselage whose central section is made by means of longitudinal panels assembled directly with each other, without circumferential junctions. It also relates to a method for producing such a fuselage. The invention has applications in the field of aeronautics and, in particular, in the field of manufacture of aircraft fuselages. State of the art
Le fuselage des aéronefs, notamment les aéronefs dédiés au transport de passagers ou de fret, est généralement réalisé à partir de plusieurs panneaux métalliques assemblés les uns avec les autres. Ces panneaux métalliques sont souvent des panneaux en aluminium. Ils sont fixés les uns avec les autres au moyen de jonctions circonférentielles et longitudinales. Ces panneaux sont assemblés de façon à former des tronçons de fuselage, eux-mêmes assemblés pour former le fuselage. Comme représenté sur la figure 1 , un fuselage d'aéronef comporte plusieurs tronçons, à savoir :The fuselage of aircraft, especially aircraft dedicated to the transport of passengers or cargo, is generally made from several metal panels assembled with each other. These metal panels are often aluminum panels. They are fixed with each other by means of circumferential and longitudinal junctions. These panels are assembled to form fuselage sections, themselves assembled to form the fuselage. As represented in FIG. 1, an aircraft fuselage comprises several sections, namely:
- le tronçon avant 1 , ou pointe avant, qui comporte notamment le poste de pilotage,- the front section 1, or front tip, which includes the cockpit,
- le tronçon arrière 2, ou cône arrière, comportant généralement l'empennage, etthe rear section 2, or rear cone, generally comprising the empennage, and
- le tronçon central 3, qui est la partie du fuselage reliant la pointe avant 1 avec le tronçon arrière 2. Le tronçon central 3 comporte généralement la cabine passagers et les soutes à bagages, pour un transport de passagers, ou les soutes à marchandises, pour un transport de fret. Le tronçon central est réalisé à partir de plusieurs éléments de tronçons, ou parties de tronçons, assemblés les uns avec les autres pour former le tronçon central reliant la pointe avant avec le cône arrière de l'aéronef. De nos jours, chaque partie du tronçon central 3, par exemple les parties 3a, 3b, 3c et 3d du fuselage représenté sur la figure 1 , est fabriquée séparément à partir de plusieurs panneaux métalliques. Sur la figure 2, on a représenté un exemple de deux parties d'un tronçon central 3 tel que réalisé classiquement. Chacune de ces parties 3a et 3b comporte plusieurs panneaux métalliques assemblés les uns avec les autres. Par exemple, la partie de tronçon 3a comporte les panneaux 41 à 46 et la partie de tronçon 3b comporte les panneaux 47 à 52. Les différents panneaux d'une même partie de tronçon sont assemblés au moyen de jonctions longitudinales. On appelle «jonction longitudinale » un type de fixation qui consiste à placer les panneaux de façon à ce que deux panneaux consécutifs se chevauchent partiellement et à insérer des éléments de fixation, tels que des rivets, dans les zones où les panneaux se superposent.- The central section 3, which is the part of the fuselage connecting the front tip 1 with the rear section 2. The central section 3 generally comprises the passenger cabin and baggage compartments, for passenger transport, or cargo bunkers, for a freight transport. The central section is made from several section elements, or parts of sections, assembled with each other to form the central section connecting the front tip with the rear cone of the aircraft. Nowadays, each part of the central section 3, for example parts 3a, 3b, 3c and 3d of the fuselage shown in FIG. 1, is manufactured separately from several metal panels. In Figure 2, there is shown an example of two parts of a central section 3 as conventionally realized. Each of these parts 3a and 3b comprises several metal panels assembled with each other. For example, the section portion 3a comprises the panels 41 to 46 and the portion of section 3b comprises the panels 47 to 52. The different panels of the same portion of section are assembled by means of longitudinal junctions. The term "longitudinal junction" refers to a type of fastening that involves placing the panels so that two consecutive panels overlap partially and insert fasteners, such as rivets, into the areas where the panels overlap.
Les parties de tronçon central sont ensuite assemblées les unes avec les autres. Deux parties de tronçon central consécutives sont assemblées au moyen de jonctions circonférentielles. On appelle «jonction circonférentielle » un moyen de fixation permettant d'assembler des parties de tronçon sur toute la circonférence de ces parties. En effet, deux parties de tronçons ne peuvent être assemblées directement l'une avec l'autre (par chevauchement) pour des raisons de tolérance, car il est impossible de réaliser deux parties de tronçon de façon à ce qu'elles s'emboîtent parfaitement l'une dans l'autre. Aussi, pour assembler deux parties de tronçon, on place généralement une virole entre les deux parties de tronçon. Une virole est une peau locale, interne au tronçon, qui assure un renfort à la zone de jonction des deux parties de tronçon. Elle permet de transférer les efforts d'un tronçon à l'autre. Cette virole est fixée, de part et d'autre, sur chaque partie de tronçon. En d'autres termes, une peau supplémentaire (réalisée à partir d'une seule ou de plusieurs pièces de peau montées sur un cadre) est placée à la jointure de deux parties de tronçon et fixée par des éléments de fixation, tels que des éclisses, sur chacune des parties de tronçon.The central section portions are then assembled together. Two consecutive central section portions are assembled by means of circumferential junctions. The term "circumferential junction" is a fastening means for assembling portions of sections along the entire circumference of these parts. Indeed, two parts of sections can not be assembled directly with each other (by overlapping) for reasons of tolerance, because it is impossible to make two parts of section so that they fit perfectly one in the other. Also, to assemble two portions of section, is generally placed a shell between the two portions of section. A ferrule is a local skin, internal to the section, which provides reinforcement to the junction zone of the two portions of section. It makes it possible to transfer the efforts from one section to another. This shell is fixed, on both sides, on each part of the section. In other words, an additional skin (made from one or more pieces of skin mounted on a frame) is placed at the joint of two portions of section and fixed by fastening elements, such as splints , on each of the section parts.
L'assemblage de différentes parties du tronçon central nécessite donc l'ajout de peaux supplémentaires et de différents éléments de fixation de ces peaux supplémentaires sur les panneaux métalliques. Ces peaux et ces éléments de fixation, généralement métalliques, sont autant des pièces qui augmentent la masse de l'aéronef.The assembly of different parts of the central section therefore requires the addition of additional skins and different fastening elements of these additional skins on the metal panels. These skins and these fasteners, generally metallic, are as many pieces that increase the mass of the aircraft.
De plus, la réalisation d'un tronçon central à partir d'autant de panneaux assemblés un par un est relativement longue à mettre en œuvre. L'assemblage de ces nombreux panneaux constitue donc un facteur important dans la durée de fabrication d'un fuselage d'aéronef.In addition, the realization of a central section from as many panels assembled one by one is relatively long to implement. The assembly of these many panels is therefore an important factor in the manufacturing time of an aircraft fuselage.
Avec l'arrivée des matériaux composites dans le domaine de l'aéronautique, les constructeurs d'aéronefs cherchent à réaliser un maximum d'éléments de fuselage en matériaux composites. En effet, les matériaux composites ont l'avantage d'être relativement léger par rapport au métal, ce qui permet d'alléger de façon significative la masse totale d'un aéronef. Pour cela, les constructeurs aéronautiques cherchent généralement à réaliser le tronçon central en matériau composite. Ils cherchent alors à reproduire une peau unique qui engloberait le tronçon central sur toute sa circonférence, soit sur 360° Autrement dit, ils ch erchent à réaliser un tronçon central d'une seule pièce. Comme il est difficile de réaliser un tel tronçon central, il a été prévu de réaliser des parties de tronçon d'une seule pièce à assembler les unes avec les autres pour former un tronçon central. Chaque partie de tronçon est donc un cylindre qui doit ensuite être assemblé avec les parties de tronçon consécutives. Cet assemblage est réalisé au moyen de jonctions circonférentielles, telles que décrites précédemment, et pour lesquelles les peaux supplémentaires sont en matériau composite.With the advent of composite materials in the field of aeronautics, aircraft manufacturers are seeking to achieve a maximum of fuselage elements made of composite materials. Indeed, the composite materials have the advantage of being relatively light compared to the metal, which allows to significantly reduce the total mass of an aircraft. For this, the aircraft manufacturers generally seek to achieve the central section of composite material. They then try to reproduce a single skin that would encompass the central section over its entire circumference, that is to say 360 °. In other words, they try to make a central section of a single piece. Since it is difficult to make such a central section, it has been planned to make one-piece portions of section to be assembled with each other to form a central section. Each section portion is therefore a cylinder which must then be assembled with the consecutive section portions. This assembly is made by means of circumferential junctions, as described above, and for which the additional skins are made of composite material.
Un exemple d'une telle réalisation de fuselage en matériau composite est décrit dans la demande de brevet PCT WO 2006/001860.An example of such a fuselage embodiment of composite material is described in PCT patent application WO 2006/001860.
Or, comme expliqué précédemment, les jonctions circonférentielles sont coûteuses en masse. En outre, elles nécessitent un temps d'assemblage relativement élevé, d'autant plus qu'elles présentent des difficultés d'assemblage car le profil aérodynamique nécessite un ajustement parfait des différentes parties de tronçon et un ajustement de l'accostage de la sous-structure, lorsque des éléments de sous-structure sont montés dans les parties de tronçon. En outre, en plus des jonctions circonférentielles, de nombreuses pièces d'interfaces sont nécessaires pour fixer la peau supplémentaire sur les deux parties de tronçon. Par ailleurs, compte tenu du découpage actuel du tronçon central en différentes parties de tronçons, un fuselage d'aéronef est transporté, d'un site à un autre, partie de tronçon par partie de tronçon. Autrement dit, chaque partie de tronçon fait l'objet d'un transport particulier sur un véhicule de transport approprié. Exposé de l'inventionHowever, as explained above, the circumferential junctions are expensive in mass. In addition, they require a relatively high assembly time, especially as they present assembly difficulties because the aerodynamic profile requires a perfect fit of the different portions of the section and an adjustment of the docking of the sub-section. structure, when substructure elements are mounted in the section portions. In addition, in addition to the circumferential junctions, many pieces of interfaces are required to secure the additional skin on the two portion portions. Moreover, given the current division of the central section into different parts of sections, an aircraft fuselage is transported, from one site to another, portion of section per portion of section. In other words, each section part is transported on a specific transport vehicle. Presentation of the invention
L'invention a justement pour but de remédier aux inconvénients des techniques exposées précédemment. A cette fin, l'invention propose un fuselage d'aéronef dont le tronçon central est réalisé à partir de panneaux longitudinaux assemblés directement les uns avec les autres, c'est-à-dire sans peaux supplémentaires. Ainsi, le tronçon central ne nécessite pas l'utilisation de jonctions circonférentielles sur toute sa circonférence.The purpose of the invention is precisely to overcome the disadvantages of the techniques described above. To this end, the invention provides an aircraft fuselage whose central section is made from longitudinal panels assembled directly with each other, that is to say without additional skins. Thus, the central section does not require the use of circumferential junctions throughout its circumference.
L'invention offre ainsi un gain de la masse totale du fuselage. En outre, l'assemblage des panneaux par jonctions longitudinales est plus simple que par jonctions circonférentielles, ce qui permet une mise en œuvre plus rapide.The invention thus provides a gain in the total mass of the fuselage. In addition, the assembly of the panels by longitudinal junctions is simpler than by circumferential junctions, which allows a faster implementation.
De façon plus précise, l'invention concerne un fuselage d'aéronef comportant une pointe avant, un tronçon arrière et un tronçon central, caractérisé en ce que le tronçon central comporte des panneaux longitudinaux assemblés directement les uns avec les autres, au moins un de ces panneaux longitudinaux ayant une longueur correspondant à la distance entre la pointe avant et le tronçon arrière pour relier ladite pointe avant avec ledit tronçon arrière.More specifically, the invention relates to an aircraft fuselage comprising a nose, a rear section and a central section, characterized in that the central section comprises longitudinal panels assembled directly with each other, at least one of these longitudinal panels having a length corresponding to the distance between the front tip and the rear section to connect said front tip with said rear section.
L'invention peut comporter également une ou plusieurs des caractéristiques suivantes :The invention may also include one or more of the following features:
- les panneaux longitudinaux sont réalisés en matériaux composites.- The longitudinal panels are made of composite materials.
- l'assemblage direct d'un premier panneau avec un second panneau comporte une superposition partielle du premier et du second panneaux et des éléments de fixation traversant lesdits panneaux. - un panneau longitudinal a une longueur correspondant à la distance entre la pointe avant et un caisson central de voilure de l'aéronef.- The direct assembly of a first panel with a second panel comprises a partial superposition of the first and second panels and fasteners passing through said panels. - A longitudinal panel has a length corresponding to the distance between the front tip and a central wing box of the aircraft.
- un panneau longitudinal a une longueur correspondant à la distance entre un caisson central de voilure et le tronçon arrière.- A longitudinal panel has a length corresponding to the distance between a central wing box and the rear section.
- un panneau longitudinal a une longueur correspondant à une largeur d'un caisson central de voilure.- A longitudinal panel has a length corresponding to a width of a central wing box.
- un panneau longitudinal intègre une sous-structure raidissante ou une structure de plancher.- A longitudinal panel incorporates a stiffening substructure or a floor structure.
- un panneau longitudinal intègre des renforts de peau.- A longitudinal panel incorporates skin reinforcements.
- un panneau longitudinal correspond à une zone du fuselage à simple courbure. - un panneau longitudinal correspond à une zone du fuselage à double courbure.- A longitudinal panel corresponds to a fuselage single curvature zone. a longitudinal panel corresponds to a zone of the fuselage with double curvature.
L'invention concerne également un procédé de réalisation d'un tel fuselage. Ce procédé comporte la fabrication d'une pointe avant, d'un tronçon arrière et d'un tronçon central. La fabrication du tronçon central se caractérise par les étapes suivantes :The invention also relates to a method of producing such a fuselage. This method comprises the manufacture of a nose, a rear section and a central section. The manufacture of the central section is characterized by the following steps:
- réalisation de panneaux longitudinaux, au moins un de ces panneaux longitudinaux ayant une longueur correspondant à la distance entre la pointe avant et le tronçon arrière, et - assemblage de ces panneaux longitudinaux directement les uns avec les autres.- Making longitudinal panels, at least one of these longitudinal panels having a length corresponding to the distance between the front tip and the rear section, and - assembling these longitudinal panels directly with each other.
Le procédé de l'invention peut comporter également une ou plusieurs des caractéristiques suivantes :The method of the invention may also include one or more of the following features:
- les panneaux longitudinaux sont réalisés en matériaux composites. - l'assemblage direct d'un premier panneau avec un second panneau consiste à superposer partiellement le premier et le second panneaux et à fixer les deux panneaux au moyen d'éléments de fixation.- The longitudinal panels are made of composite materials. - The direct assembly of a first panel with a second panel consists of partially superimposing the first and second panels and to fix the two panels by means of fasteners.
L'invention concerne également un aéronef comportant un fuselage tel que décrit précédemment. Elle concerne aussi un aéronef comportant un fuselage réalisé suivant le procédé tel que décrit précédemment.The invention also relates to an aircraft comprising a fuselage as described above. It also relates to an aircraft comprising a fuselage made according to the method as described above.
Brève description des dessinsBrief description of the drawings
La figure 1 , déjà décrite, représente un exemple de fuselage d'aéronef de l'art antérieur composé de plusieurs tronçons.FIG. 1, already described, represents an example of aircraft fuselage of the prior art composed of several sections.
La figure 2, déjà décrite, représente un exemple de parties de tronçon central selon l'art antérieur.Figure 2, already described, shows an example of central section portions according to the prior art.
La figure 3 représente un exemple de tronçon central selon l'invention.FIG. 3 represents an example of a central section according to the invention.
La figure 4 représente un autre exemple de tronçon central selon l'invention.FIG. 4 represents another example of a central section according to the invention.
La figure 5 représente un exemple de tronçon central à simple courbure et tronçon arrière à double courbure, selon l'invention.FIG. 5 represents an example of a central section with a single curvature and a rear section with double curvature, according to the invention.
Description détaillée de modes de réalisation de l'inventionDETAILED DESCRIPTION OF EMBODIMENTS OF THE INVENTION
L'invention propose de réaliser le tronçon central d'un aéronef à partir de panneaux longitudinaux, c'est-à-dire de panneaux de grande longueur réalisés chacun d'une seule pièce. De préférence, la dimension de chacun de ces panneaux longitudinaux correspond, au maximum, à la distance entre la pointe avant de l'aéronef et le tronçon arrière de l'aéronef et, au minimum, à la distance entre la pointe avant de l'aéronef et le caisson central de voilure ou entre le tronçon arrière de l'aéronef et le caisson central de voilure. Il est bien entendu que d'autres panneaux peuvent être montés entre des panneaux de grande longueur, comme par exemple, un panneau dont la longueur correspond à la largueur du caisson central de voilure. Au moins un panneau longitudinal relie la pointe avant de l'aéronef avec le tronçon arrière dudit aéronef.The invention proposes to make the central section of an aircraft from longitudinal panels, that is to say, long panels each made of a single piece. Preferably, the dimension of each of these longitudinal panels corresponds, at the most, to the distance between the nose of the aircraft and the rear section of the aircraft and, at a minimum, the distance between the nose of the aircraft and the central wing box or between the rear section of the aircraft and the central box of sail. It is understood that other panels can be mounted between panels of great length, such as, for example, a panel whose length corresponds to the width of the central wing box. At least one longitudinal panel connects the front tip of the aircraft with the rear section of said aircraft.
Le fait de réaliser un tronçon central de fuselage à partir de plusieurs panneaux longitudinaux permet un assemblage plus aisé des panneaux entre eux. En effet, les panneaux longitudinaux peuvent être assemblés au moyen de jonctions longitudinales, c'est-à-dire en superposant partiellement un panneau sur un autre panneau et en fixant les deux panneaux l'un sur l'autre au moyen d'éléments de fixations traversant les deux panneaux dans leur zone de superposition ou bien en juxtaposant deux panneaux avec un doubleur interne (virole) pour assurer la continuité.The fact of making a central fuselage section from several longitudinal panels allows an easier assembly of the panels between them. Indeed, the longitudinal panels can be assembled by means of longitudinal junctions, that is to say by partially superimposing a panel on another panel and fixing the two panels one on the other by means of elements of fixings passing through the two panels in their overlapping zone or by juxtaposing two panels with an internal doubler (ferrule) to ensure continuity.
L'assemblage de deux panneaux longitudinaux se fait, ainsi, directement, sans nécessiter de renfort localement entre les deux panneaux. Selon l'invention, les panneaux peuvent être réalisés en matériaux composites. En effet, les matériaux composites permettent la réalisation de pièces de grandes dimensions, d'un seul tenant. Ces pièces peuvent être des panneaux dont la dimension est choisie en fonction de l'aéronef à construire et non en fonction des difficultés techniques de réalisation de ladite pièce. Lorsque les panneaux sont en matériaux composites, les éléments de fixation sont des éléments connus, adaptés à la fixation de ces matériaux.The assembly of two longitudinal panels is, thus, directly, without requiring reinforcement locally between the two panels. According to the invention, the panels can be made of composite materials. Indeed, the composite materials allow the realization of large parts, in one piece. These parts may be panels whose size is chosen according to the aircraft to be built and not according to the technical difficulties of producing said part. When the panels are made of composite materials, the fasteners are known elements, suitable for fixing these materials.
Sur la figure 3, on a représenté un exemple de tronçon central de fuselage réalisé selon l'invention. Dans cet exemple, le tronçon central 3 comporte cinq panneaux longitudinaux 31 à 35, assemblés directement les uns avec les autres. En particulier, deux panneaux 31 et 32 sont assemblés pour former la partie supérieure du fuselage. Ces deux panneaux ont une longueur comprise entre la pointe avant et le cône arrière de l'aéronef. Les panneaux 33, 34 et 35 sont fixés chacun, d'une part, sur le panneau 31 et, d'autre part, sur le panneau 32. Le panneau 33 a une longueur correspondant à la distance entre la pointe avant de l'aéronef et le caisson central de voilure. Le panneau 34 a une longueur correspondant à la distance entre le tronçon arrière de l'aéronef et le caisson central de voilure. Le panneau 35 a une longueur correspondant à la largeur du caisson central de voilure. Dans l'invention, l'assemblage est réalisé longitudinalement, c'est-à- dire suivant la longueur des panneaux. Les panneaux sont fixés les uns aux autres suivant l'axe longitudinal XX de l'aéronef. Seuls certains panneaux nécessitent un assemblage circonférentiel partiel, en plus de l'assemblage longitudinal. Par exemple, le panneau 35 doit être fixé, non seulement sur les panneaux 31 et 32 avec un assemblage de type longitudinal, mais également sur les panneaux 33 et 34. L'assemblage du panneau 35 avec les panneaux 33 et 34 est un assemblage de type partiellement circonférentiel ou semi- circonférentiel. Lorsque l'assemblage n'est pas circonférentiel sur la totalité de la circonférence du fuselage, un jeu est possible entre les deux panneaux à assembler.FIG. 3 shows an example of a central fuselage section produced according to the invention. In this example, the central section 3 comprises five longitudinal panels 31 to 35, assembled directly with each other. In particular, two panels 31 and 32 are assembled to form the upper part of the fuselage. These two panels have a length between the front tip and the rear cone of the aircraft. The panels 33, 34 and 35 are each fixed, on the one hand, on the panel 31 and, on the other hand, on the panel 32. The panel 33 has a length corresponding to the distance between the front tip of the aircraft and the box central wing. The panel 34 has a length corresponding to the distance between the rear section of the aircraft and the central wing box. The panel 35 has a length corresponding to the width of the central wing box. In the invention, the assembly is carried out longitudinally, that is to say along the length of the panels. The panels are fixed to each other along the longitudinal axis XX of the aircraft. Only some panels require a partial circumferential assembly, in addition to the longitudinal assembly. For example, the panel 35 must be fixed not only on the panels 31 and 32 with a longitudinal type assembly, but also on the panels 33 and 34. The assembly of the panel 35 with the panels 33 and 34 is an assembly of partially circumferential or semi-circumferential type. When the assembly is not circumferential over the entire circumference of the fuselage, play is possible between the two panels to be assembled.
On comprend alors que les panneaux formant le tronçon central sont assemblés uniquement par des jonctions longitudinales. Aucune jonction circonférentielle n'est nécessaire, excepté aux extrémités avant et arrière du tronçon central pour fixer ledit tronçon central avec la pointe avant et le cône arrière de l'aéronef. La masse totale du tronçon central est donc allégée de la masse correspondant aux jonctions circonférentielles et autres pièces d'interface.It is understood that the panels forming the central section are assembled only by longitudinal junctions. No circumferential junction is necessary except at the front and rear ends of the central section to secure said central section with the nose and the rear cone of the aircraft. The total mass of the central section is thus lightened by the mass corresponding to the circumferential junctions and other interface parts.
La longueur des différents panneaux longitudinaux est adaptée à la forme du tronçon central de l'aéronef. Elle peut être adaptée également au mode de transport de ces panneaux longitudinaux. En effet, les panneaux longitudinaux peuvent être transportés de façon simplifiée, par rapport à une partie de tronçon central de l'art antérieur, car ils peuvent être placés les uns au-dessus des autres dans le véhicule de transport. Par exemple, les panneaux 31 et 32 de la figure 3 peuvent être placés l'un dans l'autre au fond du véhicule et les panneaux 34, 33 et 35 placés, les uns dans les autres, au- dessus des panneaux 31 et 32. En effet, le transport d'éléments de fuselage sous forme de panneaux longitudinaux permet de mieux occuper le volume de chargement mis à disposition. On peut ainsi transporter plusieurs tronçons centraux, décomposés en panneaux longs, dans un volume où l'on ne pourrait transporter qu'un seul tronçon central, si ce dernier était découpé en tronçons.The length of the various longitudinal panels is adapted to the shape of the central section of the aircraft. It can be adapted also to the mode of transport of these longitudinal panels. Indeed, the longitudinal panels can be transported in a simplified manner, with respect to a central section portion of the prior art, because they can be placed one above the other in the transport vehicle. For example, the panels 31 and 32 of FIG. 3 can be placed one inside the other at the bottom of the vehicle and the panels 34, 33 and 35 placed, one inside the other, above the panels 31 and 32. Indeed, the transport of fuselage elements in the form of longitudinal panels makes it possible to better occupy the load volume made available. We can thus transport several central sections, broken down into long panels, in a volume where we can could only carry one central section, if it was cut into sections.
En outre, l'assemblage des différents panneaux longitudinaux par chevauchement desdits panneaux est simplifié, par rapport à l'art antérieur, car le nombre de pièces de fixation est diminué de plus de la moitié. De plus, comme expliqué précédemment, les jonctions longitudinales sont plus tolérantes que les jonctions circonférentielles ; en effet, l'assemblage circonférentiel de deux parties de tronçon implique que les dimensions de ces deux parties de tronçon soient quasi identiques pour permettre une continuité dans le fuselage alors que cette contrainte n'existe pas avec un assemblage longitudinal.In addition, the assembly of the various longitudinal panels by overlapping said panels is simplified, compared to the prior art, since the number of fasteners is decreased by more than half. In addition, as explained above, the longitudinal junctions are more tolerant than the circumferential junctions; indeed, the circumferential assembly of two portions of section implies that the dimensions of these two portions of section are almost identical to allow continuity in the fuselage while this constraint does not exist with a longitudinal assembly.
Selon l'invention, les panneaux longitudinaux peuvent comporter des ouvertures et des sous-structures de l'aéronef. Dans l'exemple de la figure 3, les panneaux 31 et 32 comportent des ouvertures correspondant aux emplacements des hublots 5 et des portes passagers 4 et 6 de l'aéronef. Les panneaux 33 et 34 comportent des ouvertures 7 correspondant aux portes des cases de trappes d'atterrissage et aux portes des soutes. En effet, les techniques connues de réalisation de pièces en matériau composite permettent de fabriquer une pièce, notamment un panneau, dans laquelle des ouvertures de dimensions prédéterminées sont réalisées.According to the invention, the longitudinal panels may comprise openings and substructures of the aircraft. In the example of Figure 3, the panels 31 and 32 comprise openings corresponding to the locations of the windows 5 and the passenger doors 4 and 6 of the aircraft. The panels 33 and 34 comprise openings 7 corresponding to the doors of the landing hatch boxes and the bunker doors. Indeed, known techniques for producing composite material parts make it possible to manufacture a part, in particular a panel, in which openings of predetermined dimensions are made.
En outre, ces techniques permettent d'insérer un ou plusieurs éléments, en matériaux composites ou autres matériaux, dans un panneau en matériaux composites. Les panneaux longitudinaux peuvent donc intégrer une sous-structure raidissante de l'aéronef, telle que des lisses, des cadres, des encadrements de portes, des encadrements des hublots, voir même des amorces de structures ou des structures complètes de planchers. Ces différents panneaux peuvent également intégrer des renforts de peau tels que ceux situés au niveau des ouvertures ou proches des zones fortement chargées de l'aéronef. Dans d'autres exemples de tronçon central de fuselage réalisé selon l'invention, la partie du fuselage contenant le caisson central de voilure (correspondant au panneau 35 sur la figure 3) est réalisée dans les panneaux longitudinaux 31 , 32 ou dans les panneaux longitudinaux 33, 34. Le tronçon central 3 peut alors comporter uniquement quatre panneaux longitudinaux. Dans ces exemples, un seul assemblage circonférentiel partiel peut être suffisant, voir aucun assemblage circonférentiel partiel. Toutes les jonctions assurant l'assemblage des différents panneaux entre eux peuvent être des jonctions longitudinales.In addition, these techniques make it possible to insert one or more elements, made of composite materials or other materials, into a panel made of composite materials. Longitudinal panels can therefore integrate a stiffening substructure of the aircraft, such as rails, frames, door frames, window frames, or even structural primers or complete structures of floors. These different panels can also incorporate skin reinforcements such as those located at the openings or close to the heavily loaded areas of the aircraft. In other examples of central fuselage section produced according to the invention, the part of the fuselage containing the central wing box (corresponding to the panel 35 in FIG. 3) is made in the longitudinal panels 31, 32 or in the longitudinal panels. 33, 34. The central section 3 can then comprise only four longitudinal panels. In these examples, only one partial circumferential assembly may be sufficient, see no partial circumferential assembly. All the joints ensuring the assembly of the different panels between them can be longitudinal junctions.
On comprend donc, à la lecture de ce qui précède, que la forme des panneaux longitudinaux et leur nombre peuvent varier en fonction de différents critères, tels que le type d'aéronef à construire et le transport prévu pour ces panneaux. En particulier, la longueur des panneaux peut varier de façon à inclure ou non certains éléments de l'aéronef. Par exemple, la longueur des panneaux 31 , 32 et 34, sur la figure 3, peut varier de façon à inclure ou non l'ouverture correspondant à la porte arrière 4. Sur la figure 4, on a représenté un exemple de tronçon central réalisé à partie de cinq panneaux et n'incluant pas l'ouverture de la porte arrière 4. Dans cet exemple, un panneau 37 forme le toit du tronçon central, deux panneaux 36 et 38, symétriques de part et d'autre du panneau 37, comportent les ouvertures des hublots 5, un panneau 33 comporte une porte de soute 7a et un panneau 39 comporte une porte de soute 7b ainsi que le caisson central 8.It is therefore clear from reading the above that the shape of the longitudinal panels and their number may vary according to different criteria, such as the type of aircraft to be built and the transport provided for these panels. In particular, the length of the panels may vary to include or not certain elements of the aircraft. For example, the length of the panels 31, 32 and 34, in Figure 3, may vary to include or not the opening corresponding to the rear door 4. In Figure 4, there is shown an example of central section realized from five panels and not including the opening of the rear door 4. In this example, a panel 37 forms the roof of the central section, two panels 36 and 38, symmetrical on either side of the panel 37, comprise the openings of the windows 5, a panel 33 has a cargo door 7a and a panel 39 has a cargo door 7b and the central box 8.
Dans cet exemple de la figure 4, les panneaux longitudinaux sont destinés à former des zones de fuselage à simple courbure. On entend, par zone à simple courbure, une zone du fuselage dont le rayon de courbure est identique sur toute la longueur de la zone. Au contraire, on comprend par double courbure, une zone du fuselage dont le rayon de courbure diffère sur la longueur de la zone. Par exemple, le tronçon arrière d'un aéronef a une forme en cône. Cette zone en cône, typiquement, est une zone à double courbure. Au contraire, la partie centrale du tronçon central qui peut être cylindrique, bilobée, trilobée, ..., constitue une zone à simple courbure. D'un point vu industriel, il est plus simple de réaliser des panneaux à simple courbure, notamment en matériaux composites. En effet, les panneaux à simple courbure peuvent être réalisés avec une surface plane puis mis en forme au moment de la cuisson ou draper avec des machines de drapage automatisé spéciales car les zones à simple courbure sont développables, contrairement aux zones à double courbure. Au contraire, les panneaux à double courbure doivent être réalisés au moyen de machines relativement complexes, telles que des machines à placement de fibres pour obtenir une courbure du panneau évolutive, conforme à la forme du tronçon central souhaitée. La fabrication de panneaux à simple courbure nécessite donc un outillage moins coûteux que la fabrication de panneaux à double courbure. En outre, l'assemblage des panneaux à simple courbure peut être fait en glissant les panneaux les uns sur les autres alors que l'assemblage des panneaux à double courbure nécessite un cavage spécifique des panneaux entre eux, ce qui implique des problèmes d'hyperstaticité.In this example of Figure 4, the longitudinal panels are intended to form single curvature fuselage areas. By single curvature zone is meant a zone of the fuselage whose radius of curvature is identical over the entire length of the zone. On the contrary, it is understood by double curvature, an area of the fuselage whose radius of curvature differs along the length of the zone. For example, the rear section of an aircraft has a cone shape. This cone zone, typically, is a double curvature zone. On the contrary, the central part of the central section which can be cylindrical, bilobed, trilobed, ..., constitutes a single curvature zone. From an industrial point of view, it is simpler to produce single curvature panels, especially composite materials. Indeed, single curvature panels can be made with a flat surface and shaped at the time of cooking or draper with special automated draping machines because single curvature areas are developable, unlike double curvature areas. In contrast, double curvature panels must be made using relatively complex machines, such as fiber placement machines to achieve a curvature of the panel that is consistent with the shape of the central section. desired. The manufacture of single curvature panels therefore requires less expensive tools than the manufacture of double curvature panels. In addition, the assembly of single curvature panels can be done by sliding the panels on each other while the assembly of double curvature panels requires a specific cavage panels between them, which involves problems of hyperstaticity .
Pour ces raisons, l'invention propose un mode de réalisation dans lequel les panneaux longitudinaux ne sont appliqués qu'en zone de fuselage à simple courbure. On peut ainsi réduire le coût de fabrication de ces panneaux et réduire les contraintes de positionnement des panneaux entre eux. Comme montré dans l'exemple de la figure 5, ce mode de réalisation propose de réaliser la zone du fuselage à double courbure de façon indépendante du tronçon central à simple courbure. La partie de fuselage représentée sur la figure 5 est une zone à double courbure. Au centre du tronçon central 3, le rayon de courbure est supérieur à celui du tronçon 9. Pour éviter la réalisation de panneaux longitudinaux pour des zones à double courbure, il peut être choisi de réaliser le tronçon 9 de façon indépendante. Le tronçon 9 est alors considéré comme formant le tronçon arrière de l'aéronef. Les panneaux longitudinaux 36, 37, 38, 33, 35 et 34 sont ainsi réalisés de façon à être tous appliqués uniquement dans des zones de fuselage à simple courbure. Dans cet exemple, le tronçon 9 peut être réalisée de façon classique, c'est-à-dire sous la forme d'un tronçon à 360° fixé sur le reste du tronçon central par une jonction circonférentielle classique. For these reasons, the invention proposes an embodiment in which the longitudinal panels are only applied in fuselage zone with simple curvature. It is thus possible to reduce the manufacturing cost of these panels and to reduce the constraints of positioning the panels together. As shown in the example of Figure 5, this embodiment proposes to realize the double curvature fuselage area independently of the single curvature central section. The fuselage portion shown in Figure 5 is a double curvature zone. In the center of the central section 3, the radius of curvature is greater than that of the section 9. To avoid the production of longitudinal panels for double-curvature zones, it may be chosen to make the section 9 independently. The section 9 is then considered to form the rear section of the aircraft. The longitudinal panels 36, 37, 38, 33, 35 and 34 are thus made to be all applied only in single curvature fuselage areas. In this example, the section 9 can be made in a conventional manner, that is to say in the form of a 360 ° section attached to the rest of the central section by a conventional circumferential junction.

Claims

REVENDICATIONS
1 - Fuselage d'aéronef comportant une pointe avant (2), un tronçon arrière (2) et un tronçon central (3), caractérisé en ce que le tronçon central comporte des panneaux longitudinaux (31 - 39) réalisés en matériaux composites et assemblés directement les uns avec les autres au moyen de jonctions longitudinales, au moins un de ces panneaux longitudinaux ayant une longueur correspondant à la distance entre la pointe avant et le tronçon arrière pour relier ladite pointe avant avec ledit tronçon arrière.1 - Aircraft fuselage comprising a nose (2), a rear section (2) and a central section (3), characterized in that the central section comprises longitudinal panels (31-39) made of composite materials and assembled directly with each other by means of longitudinal junctions, at least one of these longitudinal panels having a length corresponding to the distance between the front tip and the rear section to connect said front tip with said rear section.
2 - Fuselage d'aéronef selon la revendication 1 , caractérisé en ce que l'assemblage direct d'un premier panneau avec un second panneau comporte une superposition partielle du premier et du second panneaux et des éléments de fixation traversant lesdits panneaux. 3 - Fuselage d'aéronef selon la revendication 1 , caractérisé en ce que l'assemblage direct d'un premier panneau avec un second panneau comporte une juxtaposition de deux panneaux avec un doubleur interne.2 - Aircraft fuselage according to claim 1, characterized in that the direct assembly of a first panel with a second panel comprises a partial superposition of the first and second panels and fasteners passing through said panels. 3 - Aircraft fuselage according to claim 1, characterized in that the direct assembly of a first panel with a second panel comprises a juxtaposition of two panels with an inner doubler.
4 - Fuselage d'aéronef selon l'une quelconque des revendications 1 à4 - Aircraft fuselage according to any one of claims 1 to
3, caractérisé en ce qu'un panneau longitudinal a une longueur correspondant à la distance entre la pointe avant (1 ) et un caisson central de voilure (8) de l'aéronef.3, characterized in that a longitudinal panel has a length corresponding to the distance between the front tip (1) and a central wing box (8) of the aircraft.
5 - Fuselage d'aéronef selon l'une quelconque des revendications 1 à5 - Aircraft fuselage according to any one of claims 1 to
4, caractérisé en ce qu'un panneau longitudinal a une longueur correspondant à la distance entre un caisson central de voilure (8) et le tronçon arrière (2).4, characterized in that a longitudinal panel has a length corresponding to the distance between a central wing box (8) and the rear section (2).
6 - Fuselage d'aéronef selon l'une quelconque des revendications 1 à6 - Aircraft fuselage according to any one of claims 1 to
5, caractérisé en ce qu'un panneau longitudinal a une longueur correspondant à une largeur d'un caisson central de voilure (8).5, characterized in that a longitudinal panel has a length corresponding to a width of a central wing box (8).
7 - Fuselage d'aéronef selon l'une quelconque des revendications 1 à 6, caractérisé en ce qu'un panneau longitudinal intègre une sous-structure raidissante ou une structure de plancher.7 - Aircraft fuselage according to any one of claims 1 to 6, characterized in that a longitudinal panel incorporates a stiffening substructure or a floor structure.
8 - Fuselage d'aéronef selon l'une quelconque des revendications 1 à 7, caractérisé en ce qu'un panneau longitudinal intègre des renforts de peau. 9 - Fuselage d'aéronef selon l'une quelconque des revendications 1 à 8, caractérisé en ce qu'un panneau longitudinal correspond à une zone du fuselage à simple courbure.8 - Aircraft fuselage according to any one of claims 1 to 7, characterized in that a longitudinal panel incorporates skin reinforcements. 9 - Aircraft fuselage according to any one of claims 1 to 8, characterized in that a longitudinal panel corresponds to a fuselage single curvature zone.
10 - Fuselage d'aéronef selon l'une quelconque des revendications 1 à 9, caractérisé en ce qu'un panneau longitudinal correspond à une zone du fuselage à double courbure.10 - Aircraft fuselage according to any one of claims 1 to 9, characterized in that a longitudinal panel corresponds to a fuselage zone double curvature.
1 1 - Procédé de réalisation d'un fuselage d'aéronef comportant la fabrication d'une pointe avant (1 ), d'un tronçon arrière (2) et d'un tronçon central (3) de l'aéronef, caractérisé en ce que la fabrication du tronçon central comporte les étapes suivantes :1 1 - Method for producing an aircraft fuselage comprising the manufacture of a front nose (1), a rear section (2) and a central section (3) of the aircraft, characterized in that the manufacture of the central section comprises the following steps:
- réalisation de panneaux longitudinaux (31 - 39 ) en matériaux composites, au moins un de ces panneaux longitudinaux ayant une longueur correspondant à la distance entre la pointe avant et le tronçon arrière, et - assemblage de ces panneaux longitudinaux directement les uns avec les autres au moyen de jonctions logitudinales.- Making longitudinal panels (31-39) of composite materials, at least one of these longitudinal panels having a length corresponding to the distance between the front tip and the rear section, and - assembling these longitudinal panels directly with each other by means of logitudinal junctions.
12 - Procédé selon la revendication 1 1 , caractérisé en ce que l'assemblage direct d'un premier panneau avec un second panneau consiste à superposer partiellement le premier et le second panneaux et à fixer les deux panneaux au moyen d'éléments de fixation.12 - Process according to claim 1 1, characterized in that the direct assembly of a first panel with a second panel consists of partially superimposing the first and second panels and to fix the two panels by means of fasteners.
13 - Procédé selon la revendication 1 1 , caractérisé en ce que l'assemblage direct d'un premier panneau avec un second panneau consiste à juxtaposer deux panneaux avec un doubleur interne.13 - Process according to claim 1 1, characterized in that the direct assembly of a first panel with a second panel is to juxtapose two panels with an internal doubler.
14 - Aéronef comportant un fuselage selon l'une quelconque des revendications 1 à 10.14 - Aircraft comprising a fuselage according to any one of claims 1 to 10.
15 - Aéronef comportant un fuselage réalisé avec le procédé selon l'une quelconque des revendications 1 1 à 13. 15 - Aircraft comprising a fuselage made with the method according to any one of claims 1 1 to 13.
EP07858512A 2006-10-10 2007-10-04 Aircraft fuselage made from longitudinal panels and method of producing such a fuselage Withdrawn EP2076430A2 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0654185A FR2906785B1 (en) 2006-10-10 2006-10-10 AIRCRAFT FUSELAGE MADE FROM LONGITUDINAL PANELS AND METHOD FOR PRODUCING SUCH A FUSELAGE
PCT/FR2007/052081 WO2008043940A2 (en) 2006-10-10 2007-10-04 Aircraft fuselage made from longitudinal panels and method of producing such a fuselage

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EP (1) EP2076430A2 (en)
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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2322342A1 (en) 2009-11-17 2011-05-18 Airbus Operations (S.A.S.) Machine for wrapping cylindrical composite parts
CN102126549A (en) * 2010-01-20 2011-07-20 苏章仁 Clockwise-course asymmetric simulation structure of flight carrier shell with circular cross section

Families Citing this family (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2901536B1 (en) 2006-05-23 2009-01-30 Airbus France Sas BEAM FOR PRESSURIZED FLOOR OF AIRCRAFT
FR2913412B1 (en) * 2007-03-05 2012-12-07 Airbus France CONTAINER FOR AERIAL FREIGHT TRANSPORT AND AIRCRAFT FUSELAGE FOR FREIGHT TRANSPORT.
US20080296432A1 (en) * 2007-06-04 2008-12-04 Kernkamp Industries Corporation Method of extending length of aircraft to increase interior space
DE102007028918A1 (en) 2007-06-22 2009-01-02 Airbus Deutschland Gmbh Method for producing a fuselage cell of an aircraft and device
US8752293B2 (en) * 2007-12-07 2014-06-17 The Boeing Company Method of fabricating structures using composite modules and structures made thereby
DE102008010197B4 (en) * 2008-02-20 2012-03-22 Airbus Operations Gmbh Method for connecting two fuselage sections with the creation of a transverse joint and transverse joint connection
FR2933377B1 (en) 2008-07-01 2011-04-15 Airbus France PLANE WITH FRONT LANDING TRAIN
CN102481971B (en) * 2009-10-08 2014-12-31 三菱重工业株式会社 Composite material structure, as well as aircraft wing and fuselage provided therewith
EP2671792A4 (en) * 2011-02-04 2018-01-03 MITSUBISHI HEAVY INDUSTRIES, Ltd. Composite material structure, and aircraft wing and fuselage provided therewith
JP5808111B2 (en) * 2011-02-04 2015-11-10 三菱重工業株式会社 Composite structure for aircraft, aircraft main wing and aircraft fuselage provided with the same
WO2014074085A1 (en) * 2012-11-06 2014-05-15 Hewlett-Packard Development Company, L.P. Near field communication system
WO2022169443A1 (en) * 2021-02-03 2022-08-11 Zsm Holdings Llc Devices and methods for extending aircraft cargo bays with replaceable tailcones
WO2023160824A1 (en) * 2022-02-28 2023-08-31 Ipr Conversions (Switzerland) Ltd Devices, systems, and methods for manufacturing, altering, and converting a fuselage suitable for a freighter aircraft

Family Cites Families (112)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1555409A (en) * 1920-12-27 1925-09-29 Curtiss Aeroplane & Motor Co Airplane wing
US1438685A (en) * 1921-03-21 1922-12-12 Belcher Aerial Mfg Company Internal brace construction for aeroplanes
US1541976A (en) * 1922-04-10 1925-06-16 Albin K Longren Fuselage
US1619372A (en) * 1922-11-03 1927-03-01 Rohrbach Adolf Wing construction for metal airplanes
US1545129A (en) * 1923-09-27 1925-07-07 Jr John F Cook Aircraft structure
GB246598A (en) * 1924-11-06 1926-02-04 Hugh Oswald Short Improvements in the construction of thin tubular bodies
US1799889A (en) * 1927-02-03 1931-04-07 Firm Dornier Metallbauten G M Flying-machine structure
US1829922A (en) * 1927-03-31 1931-11-03 Cairns Dev Company Wing for aircraft
US1775386A (en) * 1928-05-02 1930-09-09 Blondin Joseph Aircraft supporting and controlling surface
US1814556A (en) * 1929-04-15 1931-07-14 Jr John Brown Jewett Airplane construction and method therefor
US1765189A (en) * 1929-08-12 1930-06-17 Jr William Arthur Woolgar Aeroplane-wing construction
US1790144A (en) * 1929-09-05 1931-01-27 Augustine C Haller Wing construction for aeroplanes
US1810762A (en) * 1930-08-15 1931-06-16 Daniel B Gish Aeroplane
US2006468A (en) * 1931-10-16 1935-07-02 Albin K Longren Airplane fuselage
US1988079A (en) * 1932-05-27 1935-01-15 Curtiss Wright Airplane Compan Airplane wing
US2171434A (en) * 1934-04-14 1939-08-29 Budd Edward G Mfg Co Beam-type housing structure
US2081490A (en) * 1934-09-24 1937-05-25 Edward A Marski Airplane
DE687456C (en) * 1934-11-17 1940-06-20 Focke Wulf Flugzeugbau G M B H Shell fuselage for aircraft
GB494936A (en) * 1937-09-16 1938-11-03 Arado Flugzeugwerke Ges Mit Be Improvements in or relating to fuselages for aircraft
DE676459C (en) * 1937-09-24 1939-06-12 Dornier Werke Gmbh Fuselage
CH212555A (en) * 1938-05-17 1940-11-30 Messerschmitt Boelkow Blohm Carrier composed of several parts, especially for aircraft components.
US2382357A (en) * 1943-03-06 1945-08-14 Budd Edward G Mfg Co Metallic skin-covered structure
US2416245A (en) * 1943-06-28 1947-02-18 Budd Co Aircraft fuselage and wing construction
US2382817A (en) * 1943-08-24 1945-08-14 Ernest H Reiss Protective means for aircraft
US2430643A (en) * 1944-06-10 1947-11-11 Vidal Corp Wing structure for gliders
US2540482A (en) * 1946-10-01 1951-02-06 Hervey Foundation Inc Wooden structure and method
US2791386A (en) * 1953-10-19 1957-05-07 Lockheed Aircraft Corp Truss core
GB830072A (en) * 1955-04-21 1960-03-09 Nat Res Dev Improvements in aircraft pressure cabins
US2902597A (en) * 1956-09-12 1959-09-01 Harold W Leidy Air-borne packaging unit for electronic equipment
US2973073A (en) * 1957-11-29 1961-02-28 Northrop Corp Roller-bolster cargo handling system
FR1333200A (en) * 1962-06-14 1963-07-26 Aviation Louis Breguet Sa Improvements in the construction of aerodyne fuselages
US3256670A (en) * 1962-07-23 1966-06-21 Tersigni Ennio Prefabricated steel joist adapted for the reinforcement of floors
US3249327A (en) * 1963-12-02 1966-05-03 Jr Russell W Smith Airfoil structure
US3361401A (en) * 1966-01-03 1968-01-02 Evans Prod Co Floor structures convertible to a shipping support
US3440976A (en) * 1967-03-02 1969-04-29 Valyi Emery I Inflatable pallets
US3436038A (en) * 1967-04-11 1969-04-01 Aerojet General Co Structures
US3640491A (en) * 1970-04-15 1972-02-08 Robert R Harrison Airplane construction frame
US3727870A (en) * 1971-05-10 1973-04-17 R Bass Resilient support
US4198018A (en) * 1978-03-13 1980-04-15 The Boeing Company Blended wing-fuselage frame made of fiber reinforced resin composites
US4200946A (en) * 1978-11-16 1980-05-06 Westland Aircraft Limited Load-supporting structures
DE3141869C2 (en) * 1981-10-22 1984-02-02 Messerschmitt-Bölkow-Blohm GmbH, 8000 München Floor construction for the upper hold of an aircraft
US4674712A (en) * 1985-01-22 1987-06-23 The Boeing Company Double-lobe fuselage composite airplane
US5088661A (en) * 1986-12-31 1992-02-18 The Boeing Company Aircraft
JPH0615825Y2 (en) * 1987-07-17 1994-04-27 大日本塗料株式会社 FRP ship
US4776534A (en) * 1987-08-10 1988-10-11 Bamford Robert M Controlled torsional resistance warpable beam
US5171510A (en) * 1988-06-08 1992-12-15 Aerospatiale Societe Nationale Industrielle Method of producing a frame made of a composite material, especially for the fuselage of an aircraft
US4909655A (en) * 1989-02-06 1990-03-20 Grumman Aerospace Corporation Interleaved tab assembly for connecting structural members
FR2649665B1 (en) * 1989-07-12 1991-11-08 Airbus Ind LARGE CAPACITY FUSELAGE FOR AIRCRAFT
US5000400A (en) * 1989-12-15 1991-03-19 The Boeing Company Four-post wing-stowed main landing gear for high speed aircraft
US5553437A (en) * 1990-05-03 1996-09-10 Navon; Ram Structural beam
US5330092A (en) * 1991-12-17 1994-07-19 The Boeing Company Multiple density sandwich structures and method of fabrication
JP2507946Y2 (en) * 1992-02-25 1996-08-21 池田物産株式会社 Seat slide rail structure
JPH05286493A (en) * 1992-04-08 1993-11-02 Honda Motor Co Ltd Fuselage structure of aircraft
US5560102A (en) * 1992-10-13 1996-10-01 The Boeing Company Panel and fuselage assembly
DE4313592C2 (en) * 1993-04-26 2000-02-17 Daimler Chrysler Aerospace Wide-body aircraft
US5496001A (en) * 1994-01-28 1996-03-05 Skow; Andrew T-38 aircraft modified with an F-5 wing
US5866272A (en) * 1996-01-11 1999-02-02 The Boeing Company Titanium-polymer hybrid laminates
US6170157B1 (en) * 1996-03-22 2001-01-09 The Boeing Company Determinant spar assembly
US5954111A (en) * 1997-01-22 1999-09-21 Ochoa; Carlos M. Overhead door track structure
US6250361B1 (en) * 1997-01-22 2001-06-26 Icom Engineering Incorporated Stiffeners for sectional overhead doors
DE59806899D1 (en) * 1997-02-05 2003-02-20 Anatoli J Vassiliev Airplane for passenger and / or freight transportation
US6105902A (en) * 1997-07-15 2000-08-22 Mcdonnell Douglas Corporation Aircraft fuselage and method of forming same
DE19841799A1 (en) * 1997-12-19 1999-07-08 Grob Burkhart Dr H C Dipl Ing Assembly of fibrous composite aircraft from prefabricated sections
DE19844035C1 (en) * 1998-09-25 1999-11-25 Daimler Chrysler Aerospace Shell component for an aircraft, and method for its production
FR2793210B1 (en) * 1999-05-05 2001-09-21 Aerospatiale Airbus STRUCTURE FOR ATTACHING A LANDING GEAR TO AN AIRCRAFT FUSELAGE
GB2351750B (en) * 1999-07-07 2003-07-30 Johnson Limited Mabey Lattice panel bridges
JP4316057B2 (en) * 1999-08-04 2009-08-19 富士重工業株式会社 Composite wing
GB9922512D0 (en) * 1999-09-24 1999-11-24 British Aerospace An aircraft fuselage having a rear-end opening for cargo despatch
JP4425422B2 (en) * 2000-04-14 2010-03-03 本田技研工業株式会社 Method for producing composite material structure and composite material structure produced thereby
GB0010308D0 (en) * 2000-04-27 2000-06-14 Mabey & Johnson Ltd Lattice panel structures
JP2003034296A (en) * 2001-07-23 2003-02-04 Mitsubishi Heavy Ind Ltd Window frame part structure of sailing machine body and method for assembling window frame part
US6834833B2 (en) * 2001-10-02 2004-12-28 The Boeing Company Twin aisle small airplane
US6554225B1 (en) * 2002-06-14 2003-04-29 The Boeing Company Commercial aircraft low cost, lightweight floor design
JP3938762B2 (en) * 2003-05-30 2007-06-27 川崎重工業株式会社 Plate-like structure, reinforcing material, and method for manufacturing plate-like structure
US6926235B2 (en) * 2003-06-20 2005-08-09 The Boeing Company Runway-independent omni-role modularity enhancement (ROME) vehicle
FR2863673B1 (en) * 2003-12-10 2006-03-10 Airbus France DEVICE FOR ASSEMBLING BY SEPARATING MULTIPLE ELEMENTS AND METHOD FOR ASSEMBLING PANELS, IN PARTICULAR AIRCRAFT FUSELAGE USING THE SAME
US7234667B1 (en) * 2003-12-11 2007-06-26 Talmage Jr Robert N Modular aerospace plane
FR2866315B1 (en) * 2004-02-13 2007-04-20 Airbus France AIRCRAFT CAR SEAT RAIL AND METHOD OF MANUFACTURING SUCH A RAIL
FR2867151B1 (en) * 2004-03-04 2007-05-25 Airbus France FUSELAGE LONGERON FOR AN AIRCRAFT AND CENTRAL BOX EQUIPPED WITH SUCH A LONGERON
FR2868039B1 (en) * 2004-03-25 2007-06-08 Airbus France Sas DEVICE FOR REINFORCING A HOLLOW STRUCTURE, IN PARTICULAR OF A CAST STRUCTURE FOR AN AIRCRAFT AND HOLLOW STRUCTURE COMPRISING SUCH A DEVICE
US7527222B2 (en) * 2004-04-06 2009-05-05 The Boeing Company Composite barrel sections for aircraft fuselages and other structures, and methods and systems for manufacturing such barrel sections
DE102004033068B4 (en) * 2004-07-08 2008-09-25 Airbus Deutschland Gmbh Airliner with a main deck and a lower deck
US7708325B2 (en) * 2004-09-27 2010-05-04 L-3 Communications Integrated Systems L.P. Systems and methods for rotation of objects
US7261257B2 (en) * 2004-11-23 2007-08-28 Helou Jr Elie Cargo aircraft
FR2884801B1 (en) * 2005-04-22 2008-11-14 Eurocopter France NOSE AUXILLARY LIGHT, CARRIER STRUCTURE AND AIRCRAFT WITH A ROTARY SAIL
FR2887849B1 (en) * 2005-06-30 2008-11-21 Airbus France Sas DEVICE FOR FASTENING A SEAT RAIL ON A FLOOR, AND AIRCRAFT EQUIPPED WITH SUCH A FIXING DEVICE
FR2894225B1 (en) * 2005-12-07 2008-01-11 Airbus France Sa Sa CLOSING WALL FOR VENTRAL AIRCRAFT AND AIRCRAFT FITTING WITH VENTRAL CARENAGE
DE102006003138A1 (en) * 2006-01-24 2007-08-02 Airbus Deutschland Gmbh Emergency supply device for use in aeroplane, has back pressure turbine that is surrounded concentrically by jacket which forms flow channel and energy transducer is coupled directly to back pressure turbine
GB0603539D0 (en) * 2006-02-22 2006-04-05 Airbus Uk Ltd Control surface failsafe drop link
FR2900125B1 (en) * 2006-04-20 2009-01-16 Airbus France Sas FLOOR OF AIRCRAFT, USE OF SUCH FLOOR AND AIRCRAFT TRONCON PROVIDED WITH SUCH FLOOR
DE202006007845U1 (en) * 2006-05-16 2006-09-07 Allsafe Jungfalk Gmbh & Co. Kg Mounting system for armor or ballast plates in helicopters comprises plate with rail on its underside which fits into slot in base plate
FR2901536B1 (en) * 2006-05-23 2009-01-30 Airbus France Sas BEAM FOR PRESSURIZED FLOOR OF AIRCRAFT
FR2903656B1 (en) * 2006-07-12 2008-09-26 Airbus France Sas BOX FOR STORING THE FRONT TRAIN OF AN AIRCRAFT.
US8016234B2 (en) * 2006-09-12 2011-09-13 Airbus Deutschland Gmbh Airframe structure of an aircraft or spacecraft
US7546979B1 (en) * 2006-09-15 2009-06-16 The Boeing Company Trapezoidal panel pin joint allowing free deflection between fuselage and wing
US7775478B2 (en) * 2006-09-29 2010-08-17 The Boeing Company Floor beam assembly, system, and associated method
US20100187352A1 (en) * 2007-02-23 2010-07-29 Mr. Michael Yavilevich Multi deck aircraft
FR2913412B1 (en) * 2007-03-05 2012-12-07 Airbus France CONTAINER FOR AERIAL FREIGHT TRANSPORT AND AIRCRAFT FUSELAGE FOR FREIGHT TRANSPORT.
US7891608B2 (en) * 2007-05-03 2011-02-22 The Boeing Company Space frame fuselage structure and related methods
FR2917369B1 (en) * 2007-06-15 2009-08-07 Airbus France Sas LANDING TRAIN BOX WITH REDUCED SIZE
FR2921899B1 (en) * 2007-10-04 2011-04-15 Airbus France METHOD FOR LOCALLY REINFORCING COMPOSITE MATERIAL ELEMENT AND CENTRAL BOAT BOILER FOR REINFORCED AIRCRAFT
US7887009B2 (en) * 2007-12-05 2011-02-15 The Boeing Company Methods and systems for attaching aircraft wings to fuselages
US7967251B2 (en) * 2008-03-18 2011-06-28 The Boeing Company Truss network for aircraft floor attachment
US8157206B2 (en) * 2008-06-11 2012-04-17 Icon Aircraft, Inc. Two-motion wing-fold mechanism with independent load path
US20100012773A1 (en) * 2008-07-16 2010-01-21 Im Sunstar Aircrft Structure
US8312906B2 (en) * 2008-08-13 2012-11-20 Airbus Operations Gmbh Method and device for positioning stringers on an aircraft skin, in particular on a wing skin
FR2937099B1 (en) * 2008-10-14 2010-12-03 Airbus France FIXING SYSTEM, IN PARTICULAR FOR FIXING SEATS IN A VEHICLE AND, MORE PARTICULARLY, IN AN AIRCRAFT
FR2939405B1 (en) * 2008-12-09 2010-11-26 Airbus France AIRCRAFT FUSELAGE TRUNK
US20100163669A1 (en) * 2008-12-29 2010-07-01 Im Sunstar Super aircraft structrue
FR2947524B1 (en) * 2009-07-02 2011-12-30 Airbus Operations Sas METHOD FOR MANUFACTURING AN AIRCRAFT COMPRISING A FLOOR
FR2947525B1 (en) * 2009-07-02 2011-09-02 Airbus Operations Sas AIRCRAFT FLOOR WITH OPTIMIZED SIZE
US20120193470A1 (en) * 2011-02-01 2012-08-02 Kosheleff Patrick A Wide-body supersonic airliner

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of WO2008043940A2 *

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2322342A1 (en) 2009-11-17 2011-05-18 Airbus Operations (S.A.S.) Machine for wrapping cylindrical composite parts
CN102126549A (en) * 2010-01-20 2011-07-20 苏章仁 Clockwise-course asymmetric simulation structure of flight carrier shell with circular cross section

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RU2489312C2 (en) 2013-08-10
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JP2010505700A (en) 2010-02-25
CN101516723B (en) 2013-06-05
FR2906785B1 (en) 2009-12-04
FR2906785A1 (en) 2008-04-11
WO2008043940A2 (en) 2008-04-17
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BRPI0717735A2 (en) 2013-10-22
RU2009117613A (en) 2010-11-20

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