EP3283377A1 - Dispositif d'habillage de module d'avion - Google Patents

Dispositif d'habillage de module d'avion

Info

Publication number
EP3283377A1
EP3283377A1 EP16718615.4A EP16718615A EP3283377A1 EP 3283377 A1 EP3283377 A1 EP 3283377A1 EP 16718615 A EP16718615 A EP 16718615A EP 3283377 A1 EP3283377 A1 EP 3283377A1
Authority
EP
European Patent Office
Prior art keywords
panel
aircraft module
aircraft
connecting element
module
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Withdrawn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
EP16718615.4A
Other languages
German (de)
English (en)
Inventor
Eugen Wohlmann
Dennis Gieswein
Adam Wojsiat
Marius HAGEMANN
Philip GRÄSER
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.)
Recaro Aircraft Seating GmbH and Co KG
Original Assignee
Recaro Aircraft Seating GmbH and Co KG
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 Recaro Aircraft Seating GmbH and Co KG filed Critical Recaro Aircraft Seating GmbH and Co KG
Publication of EP3283377A1 publication Critical patent/EP3283377A1/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D11/00—Passenger or crew accommodation; Flight-deck installations not otherwise provided for
    • B64D11/06—Arrangements of seats, or adaptations or details specially adapted for aircraft seats
    • B64D11/0606—Arrangements of seats, or adaptations or details specially adapted for aircraft seats with privacy shells, screens, separators or the like
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D11/00—Passenger or crew accommodation; Flight-deck installations not otherwise provided for
    • B64D11/06—Arrangements of seats, or adaptations or details specially adapted for aircraft seats
    • B64D11/0602—Seat modules, i.e. seat systems including furniture separate from the seat itself
    • B64D11/0604—Seat modules, i.e. seat systems including furniture separate from the seat itself including a bed, e.g. cocoon type passenger seat modules
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D11/00—Passenger or crew accommodation; Flight-deck installations not otherwise provided for
    • B64D11/06—Arrangements of seats, or adaptations or details specially adapted for aircraft seats
    • B64D11/0646—Seats characterised by special features of stationary arms, foot or head rests
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16B—DEVICES FOR FASTENING OR SECURING CONSTRUCTIONAL ELEMENTS OR MACHINE PARTS TOGETHER, e.g. NAILS, BOLTS, CIRCLIPS, CLAMPS, CLIPS OR WEDGES; JOINTS OR JOINTING
    • F16B2/00—Friction-grip releasable fastenings
    • F16B2/20—Clips, i.e. with gripping action effected solely by the inherent resistance to deformation of the material of the fastening
    • F16B2/22—Clips, i.e. with gripping action effected solely by the inherent resistance to deformation of the material of the fastening of resilient material, e.g. rubbery material
    • F16B2/24—Clips, i.e. with gripping action effected solely by the inherent resistance to deformation of the material of the fastening of resilient material, e.g. rubbery material of metal
    • F16B2/241—Clips, i.e. with gripping action effected solely by the inherent resistance to deformation of the material of the fastening of resilient material, e.g. rubbery material of metal of sheet metal
    • F16B2/245—Clips, i.e. with gripping action effected solely by the inherent resistance to deformation of the material of the fastening of resilient material, e.g. rubbery material of metal of sheet metal external, i.e. with contracting action
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16B—DEVICES FOR FASTENING OR SECURING CONSTRUCTIONAL ELEMENTS OR MACHINE PARTS TOGETHER, e.g. NAILS, BOLTS, CIRCLIPS, CLAMPS, CLIPS OR WEDGES; JOINTS OR JOINTING
    • F16B21/00—Means for preventing relative axial movement of a pin, spigot, shaft or the like and a member surrounding it; Stud-and-socket releasable fastenings
    • F16B21/06—Releasable fastening devices with snap-action
    • F16B21/07—Releasable fastening devices with snap-action in which the socket has a resilient part
    • F16B21/073—Releasable fastening devices with snap-action in which the socket has a resilient part the socket having a resilient part on its inside
    • F16B21/075—Releasable fastening devices with snap-action in which the socket has a resilient part the socket having a resilient part on its inside the socket having resilient parts on its inside and outside
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16B—DEVICES FOR FASTENING OR SECURING CONSTRUCTIONAL ELEMENTS OR MACHINE PARTS TOGETHER, e.g. NAILS, BOLTS, CIRCLIPS, CLAMPS, CLIPS OR WEDGES; JOINTS OR JOINTING
    • F16B5/00—Joining sheets or plates, e.g. panels, to one another or to strips or bars parallel to them
    • F16B5/06—Joining sheets or plates, e.g. panels, to one another or to strips or bars parallel to them by means of clamps or clips
    • F16B5/0607—Joining sheets or plates, e.g. panels, to one another or to strips or bars parallel to them by means of clamps or clips joining sheets or plates to each other
    • F16B5/0621—Joining sheets or plates, e.g. panels, to one another or to strips or bars parallel to them by means of clamps or clips joining sheets or plates to each other in parallel relationship
    • F16B5/065—Joining sheets or plates, e.g. panels, to one another or to strips or bars parallel to them by means of clamps or clips joining sheets or plates to each other in parallel relationship the plates being one on top of the other and distanced from each other, e.g. by using protrusions to keep contact and distance
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16B—DEVICES FOR FASTENING OR SECURING CONSTRUCTIONAL ELEMENTS OR MACHINE PARTS TOGETHER, e.g. NAILS, BOLTS, CIRCLIPS, CLAMPS, CLIPS OR WEDGES; JOINTS OR JOINTING
    • F16B5/00—Joining sheets or plates, e.g. panels, to one another or to strips or bars parallel to them
    • F16B5/12—Fastening strips or bars to sheets or plates, e.g. rubber strips, decorative strips for motor vehicles, by means of clips
    • F16B5/123—Auxiliary fasteners specially designed for this purpose
    • F16B5/125—Auxiliary fasteners specially designed for this purpose one of the auxiliary fasteners is comprising wire or sheet material or is made thereof

Definitions

  • the invention relates to an aircraft module covering device according to the preamble of patent claim 1.
  • the invention relates to an aircraft module covering device with at least one aircraft module and at least one panel, which is provided to cover the aircraft module and is fixedly connected to the aircraft module in order to cover the aircraft module in an assembled state at least partially. It is proposed that the panel be provided via at least one
  • an "aircraft module” is to be understood as meaning, in particular, a component within an aircraft cabin, in particular a component of an aircraft passenger seat unit.
  • the aircraft module is designed as another, with a panel to be covered component of a passenger seat unit.
  • a panel is preferably designed as a planar component, but in principle it is also conceivable that the panel is designed as a narrow, thin-walled strip or a thin-walled profile element.
  • a panel is preferably designed as a cover element and is intended to cover an aircraft module in order to generate an appealing appearance
  • the panel can also be provided to protect the aircraft module to be covered from damage
  • Airplane module is ", it should be understood in particular that the panel is attached captive to the aircraft module, by applying a disassembly force, however, can also be separated again from the aircraft module.
  • the panel at least partially covers the aircraft module should be understood in particular to mean that the panel covers at least part of the aircraft module, preferably a large part of the aircraft module, and in a particularly advantageous embodiment the complete aircraft module.
  • an element is to be understood which is intended, by a form and / or adhesion, preferably by a frictional connection, ie in particular a frictional connection, a solid connection with a correspondingly designed element, at least up to a certain disassembly force in particular a connecting element to produce.
  • a “underside of the panel” is to be understood in particular as meaning a side of the panel which faces an aircraft module which is covered by the panel in an assembled state
  • "Visible side of the panel” is to be understood in particular a side of the panel, which is directed in an assembled state to the outside and can be seen by a passenger of the aircraft cabin.
  • an advantageous aircraft module covering device can be provided, the aircraft modules of which can be covered easily and quickly with a panel, wherein the - -
  • panels are easy to install and can advantageously be easily separated again from the corresponding aircraft module.
  • connection element is formed on the underside of the panel as a survey.
  • connection element is formed as a survey
  • the connection element extends away from the underside in a direction away from the visible side of the panel, wherein the survey that forms the connection element, at least has a double wall thickness of the panel, preferably at least 5 mm in size and, in a particularly advantageous embodiment, at least 10 mm, whereby the at least one connecting element can be designed to be particularly simple and advantageous.
  • the panel is intended to be detachably connected without damage to the aircraft module.
  • the term "damage-free releasable” is to be understood in particular as meaning that the panel can be dismounted from the aircraft module without damaging the elements involved in the connection between the panel and the aircraft module and without damaging the panel or the aircraft module from a state mounted captively on the aircraft module
  • the panel can advantageously be dismantled easily and quickly and, if appropriate, later be re-mounted on the aircraft module
  • the at least one attachment element be formed on the underside of the panel as a flat pin a survey be understood that in the extension direction orthogonal to the underside of the panel has a preferably rectangular cross-section, wherein first side edges are formed substantially longer than second 39kan- ten different connection elements of a P aneels different
  • a cross-section, ie a thickness and a width of a connection element, as well as the length of the connection element can be influenced by structural conditions as well as by materials of the elements to be connected. It is also conceivable that a connecting element formed as a pin, for example, has a waveform or a variable width over its length. In principle, it is also conceivable that a connection element additionally has a form-locking element, such as a latching element. - -
  • the at least one connection element can be formed particularly advantageously.
  • the at least one attachment element is materially connected to the panel.
  • the term "cohesively connected” is to be understood as meaning, in particular, a material bond, in particular by an adhesive bond, a welded connection or by another material connection appearing meaningful to a person skilled in the art
  • the at least one connecting element can be connected particularly easily to the panel.
  • the at least one attachment element is connected via an adhesive bond with the panel.
  • An “adhesive connection” should be understood to mean, in particular, a connection of at least two elements by means of adhesion forces, wherein the elements are connected to one another by an adhesive, whereby a particularly simple and secure connection of the at least one attachment element to the panel can be achieved.
  • the panel is formed as a plastic thermoformed component.
  • the panel can be formed in a particularly simple and advantageous manner.
  • the aircraft module covering device has at least one connecting element, which is provided to connect the at least one connecting element of the panel in at least one operating state captive with the aircraft module.
  • a “connecting element” should be understood to mean, in particular, an element which is intended to produce a captive connection between the connecting element of the panel and the connecting element by means of a force and / or positive locking connection, the connecting element preferably being designed as a clamping element is intended to exert a clamping force, wherein the connecting element is in particular provided to produce a frictional connection between the connecting element and the connecting element of the panel by the clamping force, in order to connect the connecting element to the connecting element.
  • the connecting element is designed as a differently designed force and / or positive locking element. It is conceivable that the connecting element comprises, for example, a friction element, a locking element or other elements which are provided for a non-positive and / or positive connection.
  • the term "captive bonding" is to be understood in particular as meaning that in a normal operating state in which no forces act on the panel which are greater than a normal load, the panel is securely connected to the aircraft module and can not be unintentionally released therefrom.
  • the at least one connecting element for the captive connection of the connecting element of the panel to the aircraft module is designed separately from the "connecting element is designed separately"
  • the connecting element is formed as an independent element and in particular is formed separately from the panel and / or the aircraft module.
  • a connection between the connecting element and the panel and / or the aircraft module is formed in a connection by a force and / or positive connection.
  • the connecting element can be formed particularly advantageous.
  • the aircraft module has at least one recess in which engages the at least one separately formed connecting element in an assembled state.
  • a particularly simple receptacle for the connecting element can be provided by means of which the connecting element can be connected to the aircraft module.
  • a "barb element" is to be understood here as meaning, in particular, an element which is directed counter to a mounting direction of a connection element which is connected to the connection element and which is provided in particular for a deflection movement To generate restoring force, which acts against the Anbin- element, which is connected to the connecting element, and thus increases a frictional force between the connecting element and the connecting element.
  • the connecting element has another, preferably a basic shape similar to a U-shape.
  • the connecting element can be formed particularly advantageously for connecting the panel to the aircraft module by means of the connecting element.
  • the connecting element is intended to be coupled with an aircraft module from a metal sheet.
  • aircraft modules which are formed from sheet metal can be covered with a panel in a particularly advantageous manner via the connection to the connecting element.
  • the connecting element is intended to be coupled to an aircraft module made of a fiber composite material.
  • aircraft modules which are formed from a fiber composite material, can be covered with a panel in a particularly advantageous manner via the connection to the connecting element.
  • connection element is always the same regardless of a material of an aircraft module to be covered.
  • a panel can be formed independently of a material of the aircraft module to be covered.
  • the aircraft module cladding device according to the invention should not be limited to the application and embodiment described above.
  • the aircraft module casing device according to the invention may have a number deviating from a number of individual elements, components and units mentioned herein.
  • FIG. 1 shows a schematic representation of an aircraft engine paneling device according to the invention with a passenger seat module
  • FIG. 2 shows a schematic detail view of a panel of the aircraft module lining device
  • FIG. 3 shows a schematic detail view of an aircraft module and a connecting element of the aircraft module covering device
  • FIG. 4 is a schematic sectional view through a connection of the aircraft module with the panel by a connecting element
  • FIG. 5 is a schematic representation of another aircraft module made of a fiber composite material
  • Fig. 6 is a schematic detail view of the aircraft module of a fiber composite material and a connecting element and
  • FIG. 7 is a schematic sectional view through a connection of the aircraft module made of the fiber composite material with the panel by a
  • FIGS. 1 to 7 show an aircraft module trim device according to the invention.
  • the aircraft module fairing device is part of an aircraft cabin.
  • the aircraft cabin comprises a passenger seat module 14.
  • the aircraft module lining device is designed as part of the passenger seat module 14.
  • the passenger seat module 14 includes a passenger seat 16 that is adapted to allow a passenger to sit on top of it.
  • the passenger seat module 14 forms a seating area for a passenger, in which, in addition to the passenger seat 16, further elements provided for the passenger, such as, for example, stowage and / or storage possibilities, are arranged.
  • the aircraft module covering device comprises a first aircraft module 10.
  • the aircraft module 10 is formed as part of the passenger seat module 14.
  • the aircraft module 10 forms part of a console, which is arranged next to the passenger seat 16 and is provided to provide a shelf on which a passenger sitting on the passenger seat 16 can deposit objects.
  • the aircraft module 10 is designed as a structural component of the passenger seat module 14.
  • the aircraft module 10 forms part of a - -
  • the aircraft module 10 is formed from a metal sheet.
  • the aircraft module 10 is designed as a bent sheet metal part.
  • the aircraft module is embodied as another component of a passenger seat module, for example as a part of the passenger seat 16, such as part of a backrest or another component of the passenger seat 16 which appears expedient to the person skilled in the art, or as
  • the aircraft module is designed as a part of an aircraft cabin, such as a wall in the area of an aircraft toilet or galley.
  • the aircraft module forms at least part of a luggage compartment of the aircraft cabin.
  • the aircraft module covering device comprises a panel 12.
  • the panel 12 is designed as a thin, flat component.
  • the panel 12 is formed as a plastic thermoforming component.
  • the panel 12 has a shape which is designed corresponding to a shape of the aircraft module 10.
  • the panel 12 is intended to completely cover the aircraft module 10 in an assembled state. In principle, it is also conceivable that the panel 12 is merely intended to partially cover the aircraft module 10, wherein a remaining part of the aircraft module 10 may remain uncovered or is covered by another panel.
  • the panel 12 has a visible side 26 facing away from the aircraft module 10 covered by the panel 12.
  • the visible side 26 faces an interior of the aircraft cabin and thus forms the surface that is perceived by passengers and can also be touched. On its visible side 26, the panel 12 accordingly has a surface that is aesthetically pleasing and damage-resistant.
  • the panel 12 is fixed and captively connected to the aircraft module 10 in an assembled state.
  • the panel 12 is intended to be connected to the aircraft module 10 via at least one frictional connection.
  • the panel 12 has a plurality of connection elements 18, 20, 22, 24.
  • the connection elements 18, 20, 22, 24 are fixedly arranged on an underside 28 of the panel 12.
  • the underside 28 of the panel 12 is opposite the visible side 26 of the panel 12.
  • the underside 28 of the panel 12 is in , ,
  • connection elements 18, 20, 22, 24 are formed as elevations, which extend away from the underside 28.
  • the connection elements 18, 20, 22, 24 on the underside 28 of the panel 12 are formed as flat pins. In this case, all of the connection elements 18, 20, 22, 24 can be of identical design. In principle, it is conceivable on the basis of structural conditions and / or material combinations that the various connection elements 18, 20, 22, 24, which are connected to the panel 12, have different cross sections and / or lengths.
  • the connection elements 18, 20, 22, 24 are fixed and captive connected to the bottom 28 of the panel 12.
  • the connection elements 18, 20, 22, 24 are materially connected to the underside 28 of the panel 12. Via an adhesive connection, the connection elements 18, 20, 22, 24 are respectively connected to the underside 28 of the panel 12.
  • an adhesive is arranged between the underside 28 of the panel 12 and an adhesive surface of the attachment elements 18, 20, 22, 24, which connects the attachment elements 18, 20, 22, 24 to the panel 12 in a materially bonded manner ,
  • the connection elements 18, 20, 22, 24 on its adhesive surface on several small, not shown elevations.
  • the elevations define the bonding gap between the respective attachment element 18, 20, 22, 24 and the underside 28 of the panel 12, which is filled for connection to the adhesive.
  • the attachment elements 18, 20, 22, 24 are placed by means of a template 12 placed on the panel and then glued.
  • the template is placed with corresponding recesses for the attachment elements 18, 20, 22, 24 defined on the underside 28 of the panel 12.
  • the connecting elements 18, 20, 22, 24 are connected to the underside 28 of the panel 12 via a defined adhesive layer of the adhesive.
  • the connection elements 18, 20, 22, 24 are adhesively bonded to the panel 12 via a robot-controlled placement. In this case, the panel 12 is firmly clamped in a defined basic position in a holder, and computer-controlled robots glue the connection elements - -
  • connection elements 18, 20, 22, 24 fixed to the panel 12 in another manner that would appear to be appropriate to the person skilled in the art are connected, for example via a positive connection.
  • the aircraft module 10 per attachment element 18, 20, 22, 24 of the panel 12 has a recess 30.
  • the recesses 30 are formed as through holes, which are introduced into the aircraft module 10 formed from a sheet metal.
  • the recesses 30 are formed as positive locking elements.
  • the aircraft module covering device per attachment element 18, 20, 22, 24 has a connecting element 32.
  • the connecting elements 32 are intended to be arranged in a mounted state in a recess 30 of the aircraft module 10 and to receive a corresponding connecting element 18, 20, 22, 24 of the panel 12.
  • the connecting elements 32 are each formed as separate components.
  • Each connecting element 32 per se is formed as a separate, separate component.
  • connection of the panel 12 to the aircraft module 10 via the further pairings of corresponding connection element 18, 20, 22, 24, connecting element and corresponding recess in the aircraft module 10 takes place accordingly.
  • the connecting element 32 is designed as a bent sheet metal part. In principle, it is also conceivable that the connecting element 32 is formed from another material which appears appropriate to a person skilled in the art.
  • the connecting element 32 has a U-shaped basic shape.
  • the connecting element 32 In the region in which the connecting element 32 is U-shaped in a cross-section, the connecting element 32 has two side walls which are spaced apart from one another by a smaller width than the width of the recess 30 of the aircraft module 10, in which the connecting element 32 is arranged in the assembled state.
  • the connecting element 32 On its upper, open side, the connecting element 32 has support elements 34, 36 which are each bent by 90 degrees at the two ends.
  • the support elements 34, 36 are intended to be supported in an assembled state on an upper side 38 of the aircraft module 10.
  • ment 32 has two inwardly directed barb elements 40, 42.
  • the barb elements 40, 42 are opposite each other and extend toward each other. Depending on a barb element 40, 42 is each formed by a side wall of the connecting element 32. In an unloaded state, the two barb elements 40, 42 have a minimum distance from one another which is smaller than a thickness of the attachment element 18.
  • the barb elements 40, 42 are designed to be elastically deformable. To connect the connecting element 32 to the connecting element 18, the connecting element 18 is pushed between the two barb elements 40, 42. Since the attachment element 18 is thicker than a distance between the two abutment elements 40, 42 in an unloaded state, the barb elements 40, 42 are deflected outwards.
  • connection element 18 is connected firmly and captively to the aircraft module 10 via the connection element 32.
  • connection element 32 Via the connection by means of the connection element 18 and the connection element 32 and the further connection elements 20, 22, 24 of the panel 12 and the further connection elements, the panel 12 can be connected without damage to the aircraft module 10, ie in particular to the recesses 30 of the aircraft module 10. Further, when disassembling the panel 12 from the aircraft module 10, no other components such as a double-sided adhesive tape, screws or the like need to be removed.
  • connection elements 18, 20, 22, 24 By applying a disassembly force on the panel 12 and on the corresponding connection elements 18, 20, 22, 24, which is directed against a mounting direction in which the connecting elements 18, 20, 22, 24 are inserted into the respective connecting elements 32, the Connecting elements 18, 20, 22, 24 are released from the corresponding connecting elements 32.
  • the disassembly force is greater than one of the - -
  • connection elements 18, 20, 22, 24 clamping force.
  • the disassembly force corresponds per attachment element 18, 20, 22, 24 each 15 Newton.
  • An entire disassembly force results from the individual disassembly forces of all attachment elements 18, 20, 22, 24.
  • the attachment elements 18, 20, 22, 24 slide in the disassembly direction from the corresponding connection elements 32, wherein neither the attachment elements 18, 20, 22, 24 nor the connecting elements 32 nor the aircraft module 10 are damaged or destroyed.
  • no further elements have to be removed from the aircraft module 10 or the panel 12.
  • the panel 12 has not shown in more detail for disassembly markings that point a fitter points of attack, where he can apply the disassembly force on the panel 12 for safe disassembly.
  • An assembly of the panel 12 on the aircraft module 10 is possible with the same connecting elements 32 and connecting elements 18, 20, 22, 24.
  • the attack points are displayed only in a manual for the fitter.
  • the aircraft module fairing device comprises a further aircraft module 44.
  • the aircraft module 44 is formed as part of the passenger seat module 14.
  • the aircraft module 44 forms part of a shell of the passenger seat module 14.
  • the aircraft module 44 forms a side wall of the passenger seat module 14.
  • the aircraft module 44 is formed from a fiber composite panel 78 and a stiffening structure 80 connected to the fiber composite panel 78. In this case, an outwardly directed side of the fiber composite material plate 78 is formed and the stiffening structure 80 is fixedly connected on an inner side plate 78 with the fiber composite material.
  • the stiffening structure 80 is formed as a honeycomb structure. In principle, it is also conceivable that the stiffening structure 80 is embodied as another structure which appears expedient to the person skilled in the art, for example as a phenolic resin or a foam.
  • the aircraft module fairing device comprises a panel 46, which is provided to cover the aircraft module 44, which is formed from a fiber composite material, in an assembled state.
  • the panel 46 disguises the aircraft module 44 in the assembled state.
  • the panel 46 has a plurality of connection elements 48.
  • the panel 46 is like the panel 12 as a thin - -
  • the panel 46 is also formed as a plastic thermoforming component.
  • the panel 46 in this case has a shape which is formed corresponding to a shape of the aircraft module 44.
  • the attachment members 48 are the same as the attachment members 18, 20, 22, 24 of the panel 12 and connected in a similar manner to the panel 46.
  • the connection elements 18, 20, 22, 24, 48 are always the same regardless of a material of the aircraft module 10, 44 to be covered.
  • the aircraft module 44 per attachment element 48 of the panel 46 each have a recess 50, 52, 54, 56, 58.
  • the recesses 50, 52, 54, 56, 58 are formed as blind holes and plate in the fiber composite material 78 and the stiffening structure 80 of the aircraft module 44 introduced.
  • the recesses 50, 52, 54, 56, 58 have a rectangular basic shape with rounded side edges.
  • the panel 46 has reinforcement zones 82, 84.
  • the reinforcing structure 80 formed as a honeycomb structure is filled with a reinforcing material for reinforcement.
  • the reinforcing material is formed as a filler.
  • the reinforcing material is formed as a resin, in particular as a phenolic resin.
  • the reinforcing zones 82, 84 are thereby larger than the recesses 50, 52, 54, 56, 58.
  • the reinforcing zones 82, 84 in each case enclose the associated recess 50, 52, 54, 56, 58 for connecting the panel 46 to the Airplane module 44, the aircraft module covering device per attachment element 48, a connecting element 60, 62, 64, 66, 68 on.
  • the connecting elements 60, 62, 64, 66, 68 are provided to be arranged in an assembled state in one of the recesses 50, 52, 54, 56, 58 of the aircraft module 44 and to receive a corresponding connecting element 48 of the panel 46.
  • the connecting elements 60, 62, 64, 66, 68 are each formed as separate components.
  • Each connecting element 60, 62, 64, 66, 68 per se is formed as an independent, separate component.
  • the connecting elements 60, 62, 64, 66, 68 are intended to be coupled to an aircraft module 44 made of a fiber composite material. In the following, only one connecting element 60 and a connection of the connecting element 60 to the connecting element 48 and the recess 50 will be described in more detail.
  • the connecting element 60 is formed as a bent sheet metal part. In principle, it is also conceivable that the connecting element 60 is formed from another material which appears appropriate to a person skilled in the art.
  • the connecting element 60 has a U-shaped basic shape. In the region in which the connecting element 60 is U-shaped in a cross-section, the connecting element 60 has two side walls which are spaced apart from one another by a width of the recess 50 of the aircraft module 44 in which the connection is made. is arranged element in the assembled state 60.
  • the connecting element 60 has two inwardly directed barb elements 70, 72.
  • the barb elements 70, 72 are opposite each other and extend toward each other. Depending on a barb element 70, 72 is each formed by a side wall of the connecting element 60. In an unloaded state, the two barb elements 70, 72 have a minimum distance from one another which is smaller than a thickness of the connecting element 48. The barb elements 70, 72 are of elastically deformable design. To connect the connecting element 60 to the connecting element 48, the connecting element 48 is pushed between the two barb elements 70, 72. Since the attachment element 48 is thicker than a distance between the two barb elements 70, 72 in an unloaded state, the barb elements 70, 72 are deflected outwards.
  • the connecting element 60 has two outwardly directed barb elements 74, 76, which are provided for connecting the connecting element 60 to the aircraft module 44. If the attachment element 48 is guided between the inwardly directed barb elements 70, 72 during assembly, the outwardly directed barb elements 74, 76 are pushed outward and thus pressed into the reinforcement zone 82 of the aircraft module 44. As a result, the connecting element 60 is connected in a form-fitting manner in the recess 50 via the outwardly directed barb elements 74, 76 and is thus arranged captively in the recess 50 when the connecting element 48 pushes the barb elements 70, 72, 74, 76 outwards.
  • connection elements 48 can be introduced, the connection elements 48 can be released from the respective connecting elements 60, 62, 64, 66, 68.
  • the disassembly force is greater than one of the connecting elements 60, 62, 64, 66, 68 respectively acting on the connection elements 48 clamping force.
  • the connection elements 48 slide in the dismounting direction from the corresponding connecting elements 60, 62, 64, 66, 68, wherein neither the connecting elements 48 nor the connecting elements 60, 62, 64, 66, 68 nor the aircraft module 44 are damaged or destroyed.
  • An assembly of the panel 46 on the aircraft module 44 is possible with the same connecting elements 60, 62, 64, 66, 68 and connecting elements 48.
  • the aircraft module covering device comprises further aircraft modules 86, 88, 90.
  • the aircraft modules 86, 88, 90 will not be described in detail below.
  • the one aircraft module 88 is designed as a handrail.
  • the aircraft module 88 is designed as a bumper.
  • the aircraft module 90 is formed as a side cover.
  • the aircraft modules 86, 88, 90 are each covered by a panel not described in greater detail, like the aircraft modules 10, 44 described above.
  • the connection of the corresponding panel with the aircraft modules 86, 88, 90 takes place in accordance with the above-described connection of the panels 12, 46 with the aircraft modules 10, 44.

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Connection Of Plates (AREA)

Abstract

L'invention concerne un dispositif d'habillage de module d'avion comportant au moins un module d'avion (10, 44) et au moins un panneau (12, 46) qui est destiné à recouvrir le module d'avion (10, 44) et qui est relié solidement au module d'avion (10, 44) pour recouvrir au moins partiellement le module d'avion dans un état assemblé. Selon l'invention, le panneau (12, 46) est prévu pour être relié au module d'avion (10, 44) par au moins une liaison à force et comporte pour cela au moins un élément de liaison (18, 20, 22, 24, 48) qui est disposé de façon fixe sur un côté inférieur (28) du panneau (12, 46).
EP16718615.4A 2015-04-13 2016-04-08 Dispositif d'habillage de module d'avion Withdrawn EP3283377A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102015105619.4A DE102015105619A1 (de) 2015-04-13 2015-04-13 Flugzeugmodulverkleidungsvorrichtung
PCT/EP2016/057762 WO2016166028A1 (fr) 2015-04-13 2016-04-08 Dispositif d'habillage de module d'avion

Publications (1)

Publication Number Publication Date
EP3283377A1 true EP3283377A1 (fr) 2018-02-21

Family

ID=55809075

Family Applications (1)

Application Number Title Priority Date Filing Date
EP16718615.4A Withdrawn EP3283377A1 (fr) 2015-04-13 2016-04-08 Dispositif d'habillage de module d'avion

Country Status (4)

Country Link
US (1) US20180170548A1 (fr)
EP (1) EP3283377A1 (fr)
DE (1) DE102015105619A1 (fr)
WO (1) WO2016166028A1 (fr)

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Also Published As

Publication number Publication date
DE102015105619A1 (de) 2016-10-13
WO2016166028A1 (fr) 2016-10-20
US20180170548A1 (en) 2018-06-21

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