Airbus Operations Gmbh

Airbus Deutschland Gmbh

Deutsches Zentrum Fur Luft-und Raumfahrt E.v.

Paul Jörn

Markus Müller

Peter Linde

Pierre C Zahlen

Alexei Vichniakov

Axel Siegfried Herrmann

Ulrich Eberth

Norbert Heltsch

Claus Hanske

Ichwan Zuardy

Karim Grase

Remo Hinz

Martin Metten

Markus Kleineberg

Bernd Schwing

Andreas Jespersen

Christian Hühne

Tobias Ströhlein

Scott Clair

Steffen Niemann

Arne Roth

Wouter Brok

A method for manufacturing a structural element for a fuselage of an aircraft. To improve the manufacture of structural elements, a method includes laying up textile material members on a mandrel to form a plurality of structural element preforms that are space apart along an extended direction of the mandrel. The structural element preforms form closed loops and are subsequently cured to obtain annular structural elements. The annular structural elements are used as basic building blocks for stiffening panel members or are directly used as structural frame elements reinforcing cut-outs in a fuselage for windows and/or doors.

Method for manufacturing a structural element for a fuselage of an aircraft

Processes for determining carbon dioxide emission impact. In one process, a carbon dioxide emission impact is determined for each stage of a multi-stage trip. A traveler may make adjustments to the trip to offset the carbon dioxide emission impact of the entire trip and achieve a desired sustainability level for the trip. In another process, that may be used in the first process, the carbon dioxide emission impact for a vehicle is determined based on the entire life span of the vehicle, including a carbon dioxide emission impact for a production stage, a carbon dioxide emission impact for a usage stage, and a carbon dioxide emission impact for an end of life stage.

Processes for offsetting a carbon footprint of a multistage trip and for the lifecycle of a vehicle

A textile fiber-composite material precursor and method for producing a component from fiber-composite material. Aircraft components can be produced from polymer fiber-composite materials, a matrix of which can be a high-performance plastics material such as polyether ketone ketone wherein a reinforcement of a non-crimp fabric of carbon fibers is embedded. Large-area non-crimp fabrics and large-area polymer films can be consolidated while being heated and pressed forming simple components. The flexible textile fiber-composite material precursor includes a stack of woven-fabric tiers from a polymer and of non-crimp fabric tiers from carbon fibers. Since both components are capable of draping, the fiber-composite material precursor can be deposited over a large area on curved shape-imparting surfaces and subsequently be consolidated under pressure and heated to form the fiber-composite material.

Textile fiber-composite material precursor, and method for producing a component from fiber-composite material

A molding device for producing a shell element reinforced with supporting elements and composed of fiber composite material, including a mold part with a mold surface including a cavity for receiving a supporting element of the shell element, and a mold core to be arranged in the cavity and to support the supporting element on the shell element when the supporting element of the shell element is arranged in the cavity. The molding device enables prevention of the mold cores from falling out in an uncontrolled manner when the shell element is removed from the mold as the mold core includes a holding device to engage with the supporting element of the shell element to be produced located in the cavity and to hold the mold core on the supporting element when the shell element to be produced is released from the mold part.

Molding device and method for producing a shell element reinforced with supporting elements and composed of fiber composite material

A process for producing a polymeric stiffened sheet-like component, for example a panel, for aircraft construction. Production includes integration of hollow stiffening profiles, for example closed omega stringers, onto a sheet-like component, for example an external skin, where the stringers and external skin are produced from thermoplastic composite material. The stringers are integrated onto the external skin by establishing contact between the stringers and the external skin and melting thermoplastic composite material with exposure to heat and pressure at the areas of contact between external skin and stringers. Melting of the other sections of the stringers is avoided with a pressurized cooling fluid with a temperature significantly below the melting point of thermoplastic composite material, the fluid flowing through the airtight enclosed space in the stringers. Use of closed airtight thermoplastic omega stringers allows integration of the stringers onto the external skin in absence of any flexible tube within the stringers.

Polymeric stiffened sheet-like component and process for producing the same

A process is proposed for producing a component, in which at least one first two-dimensional laid scrim is produced from a fibre composite material having a constant thickness, an outer face and an inner face. At least one second two-dimensional laid scrim is arranged with respect to the first two-dimensional laid scrim and then consolidated. The first two-dimensional laid scrim can be provided in a continuous, efficient production process, which increases speed and lowers costs.

Process for producing a component which is two-dimensional in regions from a fibre composite material

A reinforcing arrangement for an opening in an aircraft fuselage structure has a fuselage skin having an inner side, an outer side and a first opening with a first opening contour, and a one-piece reinforcing structure having a thickening section and at least one projecting profile. The first opening contour is surrounded on the inner side of the fuselage skin by an opening edge surface having a thickness corresponding to a thickness of the fuselage skin surrounding the opening edge surface. The thickening section has a bearing surface matched to the opening edge surface and has a receiving section facing away from the bearing surface and surrounding a second opening. The bearing surface is in surface contact with the opening edge surface, and the reinforcing structure is connected to the fuselage skin, with the result that the first opening and the second opening lie one above the other.

Reinforcing arrangement for an opening in an aircraft structure

A pressure bulkhead for an aircraft fuselage including a sandwich structure defining a central axis and extending between a circumferential border area, the sandwich structure including an inner face sheet and an outer face sheet extending transverse to the central axis, and a core assembly sandwiched between the inner face sheet and the outer face sheet. The pressure bulkhead can receive pressure loads and can be used as a structural bearing element for receiving loads introduced by other structural parts. The inner face sheet and/or the outer face sheet has a corrugated shape. The distance between the outer face sheet and the inner face sheet increases and subsequently decreases from the border area to a central area about the central axis, forming a thickened area where the distance between the outer face sheet and the inner face sheet has a peak.

Pressure bulkhead for an aircraft fuselage, and an aircraft comprising such a pressure bulkhead

A method for manufacturing a rivet connection of a fiber composite component. The method includes positioning a first component which contains a fiber composite material in an overlap joint with a second component, laser-drilling a shared through-hole at least through the fiber composite material of the first component, inserting a rivet into the through-hole and fixing the rivet to the first and to the second component.

Method for manufacturing a rivet connection of a fiber composite component

An aircraft structural component having a skin element, including an inner surface, a multitude of transverse stiffening elements, which lie against the inner surface of the skin element, and a multitude of longitudinal stiffening elements, perpendicular to the transverse stiffening elements against the inner surface of the skin element, wherein between a first and a second transverse stiffening element and also between a first and a second longitudinal stiffening element, a first skin section is defined. An aircraft structural component is provided with the quantity of the longitudinal stiffening elements reduced. A first applied reinforcement includes a core layer having a contour line running along the circumference of the skin element, the first applied reinforcement including a face sheet surrounding the core layer on the side thereof which faces away from the skin element, which face sheet is connected to the skin element along the core layer contour line.

Aircraft structural component

This pertains to a method for connecting fiber-reinforced structural components. In a first step of the method, a first fiber-reinforced structural component and a second fiber-reinforced structural component are supplied. In another step, the first fiber-reinforced structural component and the second fiber-reinforced structural component are heated by means of a heating device. A penetration through the first fiber-reinforced structural component and the second fiber-reinforced structural component is produced in another step by means of a fastening device. In another step, the first fiber-reinforced structural component is connected to the second fiber-reinforced structural component by means of the fastening device. This furthermore pertains to a structural aircraft element.

Method for connecting fiber-reinforced structural components

Described and illustrated is a pressure bulkhead for an aircraft fuselage, comprising a sandwich structure defining a central axis and extending between a circumferential border area which is configured for being mounted to a fuselage shell, wherein the sandwich structure comprises an inner face sheet extending perpendicularly with respect to the central axis, an outer face sheet opposite the inner face sheet, extending perpendicularly with respect to the central axis, and a core assembly sandwiched between the inner face sheet and the outer face sheet, wherein the inner face sheet, when viewed in a radial cross section along the central axis, has an even shape. The object of the present disclosure, to provide a pressure bulkhead for an aircraft fuselage, wherein the required weight is reduced, is achieved in that the outer face sheet, when viewed in the radial cross section along the central axis, has a convex shape, wherein the distance between the outer face sheet and the inner face sheet increases tangent continuously from the border area to the central axis.

A pressure bulkhead () for an aircraft fuselage () having a sandwich structure () which defines a central axis () and which extends between a circumferential edge area (), wherein the sandwich structure () includes an inner cover layer (), an outer cover layer (), and a core layer (), which extends between the inner and outer cover layers () and connects them. The object of providing a pressure bulkhead () for an aircraft fuselage () which, even in the case of larger fuselage diameters, can particularly effectively absorb the occurring pressure forces and which can also be built with minimal weight, is achieved in that, between the inner and the outer cover layers (), in addition to the core layer (), a support structure () is provided, which is connected to the cover layers () and which extends from the inner cover layer () to the outer cover layer (), and in that the support structure () comprises at least one layer () which, viewed in a cross section () parallel to the central axis (), extends from a first section () of the edge area () to an opposite second section () of the edge area ().

Pressue bulkhead for an aircraft fuselage

The invention relates to a method for tolerance compensation between two fiber composite component parts () for aircraft. In order to compensate in simplified fashion tolerances between two fiber composite component parts being joined, the method according to the invention has the steps:

Method for tolerance compensation between two fibre composite components

An aircraft fuselage element includes an outer shell and an inner shell that is disposed with clearance relative to the outer shell. An intermediate space between the outer shell and the inner shell is connectable to an air source. A ventilation aperture that is formed in the aircraft fuselage element and connectable to the air source is provided with a semi-permeable membrane. The semi-permeable membrane is permeable to air flow in both directions and permeable to water in a flow direction into the intermediate space from outside the outer shell or the inner shell.

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Hamburg, Germany

Paul Jörn