Airbus Operations Gmbh

Airbus Deutschland Gmbh

Fraunhofer-Gesellschaft Zur Foerderung Der Angewandten Forschung E.v.

Paul Joern

Ulrich Eberth

Markus Mueller

Norbert Heltsch

Karim Grase

Wolf-Dietrich Dolzinski

Markus Joerg Weber

Arne Hartwig

Jens-Ulrich Prowe

Alexei Vichniakov

Hans-Juergen Heinrich

Ichwan Zuardy

Claus Hanske

Bernd Schwing

Arne Roth

Ralf Herrmann

A method for producing a fuselage portion, in particular for an aircraft or spacecraft, has the following method steps: welding a skin portion containing a thermoplastic material with a former containing a thermoplastic material in the region of a predetermined welding zone; and connecting an attachment element, configured as a crack stopper, to the skin portion and to the former in the region of the welding zone.

Method for producing a fuselage portion

A method for integrating a backing-structure assembly in a structure of an aircraft or spacecraft is described. In this case, a plurality of individual elements is joined to form the backing-structure assembly. The individual elements for the backing-structure assembly and a skin portion for the structure are provided. The elements are arranged on a pre-assembly device which comprises retaining devices, which are each configured to hold one of the elements so as to be adjustable with respect to the position and/or location thereof. Some or all of the elements are connected to the skin portion. In the method, by adjusting the retaining devices for tolerance compensation, gaps between joint regions of the elements and the skin portion are eliminated or adjusted.

Method for integrating a backing-structure assembly in a structure of an aircraft or spacecraft

A sandwich structure with an outer cover layer, inner cover layer, core layer, and pins extending through the core layer. The sandwich structure has a large core thickness resistant to crack formation and to detachment of the cover layers from the core layer. An anchor layer is between the outer and inner cover layers and divides the core layer, at least in areas, into an outer core layer section and inner core layer section. An outer group of pins extends through the outer core layer section and an inner group of pins extends through the inner core layer section. The outer group pins are connected to the anchor layer with their first end and to the outer cover layer with their opposite second end, and the inner group pins are connected to the anchor layer with their first end and to the inner cover layer with their opposite second end.

Sandwich structure

A passenger aircraft with a passenger cabin is described, wherein the passenger aircraft comprises a fully functional door, which is openable and closable and which is adapted for an entrance and an exit of passengers, and en emergency exit door, which is exclusively adapted for an emergency exit of passengers from the passenger cabin. Therein, the emergency exit door is arranged in an area of a fuselage of the passenger aircraft, which is arranged ahead of and/or behind of wings of the passenger aircraft.

Passenger aircraft with an emergency exit door

The invention relates to a fixing threadfor sewing together reinforcing fibers, in particular to form a fiber preform, for the production of a fiber-reinforced composite component by means of impregnation with a curable polymer material. According to the invention, the fixing threadhas an interfacefor locally increasing the ductility of a matrix formed by the curable polymer material and has at least one core thread, so as in particular to reduce the tendency for microcracks to form in the matrix of the finished composite component in the regions where the fixing threadspenetrate. The interfacemay be formed, for example, by an enclosure of the core threadby a thermoplastic polymer material, for example by polyethylene, polypropylene or the like.

Fixing thread for sewing together reinforcing fibers

An aircraft fuselage includes an exterior skin; at least one window frame; at least one window package disposed in a breakout of the exterior skin and attached to the at least one window frame; and a plurality of stiffening elements configured to stiffen the exterior skin, wherein at least one of the stiffening elements crosses the at least one window package.

Aircraft fuselage

A pressure bulkhead made of fiber composite material for an aircraft fuselage for the pressure-tight axial closure of a fuselage region configured to be put under internal pressure. The bulkhead comprises a first area element, at least one second area element and a plurality of core elements. The first area element has a plurality of receptacles. Each receptacle is configured to receive a core element. The core elements are received in the receptacles between surfaces facing one another of the first area element and of the at least one second area element. The surfaces facing one another of the first area element and of the at least one second area element are connected to one another in regions without core elements.

Pressure bulkhead for an aircraft and device as well as method for manufacturing an aircraft structural component

A segment and a method for producing such a segment, in particular for an aircraft, having a skin field and a skin field stiffening structure. The skin field and stiffening structure are made from fiber-reinforced plastic materials. The stiffening structure is bonded to an inner surface of the skin field and comprises a multiplicity of first stiffeners and mounts which are formed integrally with the first stiffeners and extend in a transverse direction of the first stiffeners for the connection of second stiffeners to form an integral lattice structure. Reinforcement elements are provided for mechanically fixing the mounts to the skin field. Each reinforcement element has a connection section and a blocking section having an extension in a direction inclined to the connection section. The connection section penetrates the skin field and the mounts, and the blocking section is in contact with an outer surface of the skin field.

Shell segment of an aircraft and a production method

A fuselage cell structure for an aircraft includes at least two skin panels including at least one double shell skin panel and at least one monolithic skin panel. At least one longitudinal brace or a transverse brace is disposed so as to form at least one of longitudinal seam or a transverse seam between the double shell skin panel and the monolithic skin panel. The fuselage cell structure includes at least one of a longitudinal bracket and a load transfer point. The longitudinal bracket is disposed in a region of the longitudinal seam and includes a first and a second longitudinal flanges disposed offset with respect to one another and connected by a web. The load transfer point is disposed in a region of the transverse seam so as to connect the longitudinal brace disposed on the monolithic skin panel to the double shell skin panel.

Fuselage cell structure for an aircraft in hybrid design

The present invention relates to a fuselage of an aircraft or spacecraft, with at least one shell element and a structural element. An interspace to which air can be admitted by an air stream is provided between the at least one shell element and the structural element. The fuselage is distinguished by the fact that, to form the air-admitting air stream as an outgoing/incoming air stream of a pressurized interior space of the fuselage, the interspace is connected to a corresponding outgoing/incoming air connection of the interior space. The invention also relates to a corresponding aircraft or spacecraft and to a method of actively insulating such a fuselage.

Fuselage of an aircraft or spacecraft and method of actively insulating such a fuselage

The invention relates to a fuselage of an air- or spacecraft. The fuselage comprises at least one shell element and one insulation element. The insulation element is configured as a passive, watertight insulation element and can be mounted on an inner side in a completely air- and watertight manner. The invention furthermore relates to a respective aircraft or spacecraft.

Fuselage of an aircraft or spacecraft and corresponding aircraft or spacecraft

Disclosed is a method for producing single- or multi-layered fiber preforms by the TFP process with fiber strands which are aligned in particular such that they are oriented with the flux of force, wherein the fiber preforms have virtually any desired material thickness without troublesome backing layers and have virtually any desired surface geometry, comprising the steps of: laying and attaching the fiber strands on a flexible and elastic base, in particular a base formed by an elastomer, with a fixing thread led through a sewing head to form the fiber perform; and lifting the fiber preform off the elastic and flexible base. The fiber preforms produced by means of this method have a virtually optimum fiber alignment, that is to say substantially oriented with the flux of force, and no appreciable flaws in the arrangement of fibers, and consequently make it possible to create composite components that can withstand extreme mechanical stress and are at the same time lightweight, for example by subsequent processing in the RTM process.

Method for producing single-or multi-layered fiber preforms by the TFP process

A window element for a double-shell skin field of an aircraft fuselage cell includes a window frame having a window module disposed in the frame, wherein the frame is disposed in an opening of the double-shell skin field, and wherein the frame is connected to the double-shell skin field by a surrounding embrasure profile positioned against a core structure of the skin field at least in some regions, wherein the core structure includes an inner and an outer covering layer.

Window element for a double-shell skin field of an aircraft fuselage cell

A method for producing single- or multi-layered fiber preforms by the TFP process with fiber strands, aligned substantially to be oriented with the flux of force, laid on a backing layer and attached by fixing threads, to form a fiber preform with a desired material thickness and without a backing layer. A release layer is applied at least in certain regions to the backing layer and, after completion of the TFP process, the fiber preform is introduced into a fixing device, in which at least some of the fixing threads are melted. Subsequently the backing layer, separated from the fiber preform by the release layer, is detached from the fiber preform. The fiber preforms have a virtually optimum fiber structure without flaws along with a desired material thickness and are predestined for the production of composite components for load-bearing components that must withstand high mechanical loads.

Method for producing single- or multi-layered fiber preforms by the TFP process

A guiding device for a device for producing fibre preforms by the TFP process, the guiding device guiding the fibre strand. The guiding device has an opening with a variable cross-section for leading through the fibre strand, wherein the guiding device is fitted in the region of a sewing head of the TFP process device. The width of the fibre strands to be laid and attached can be increased, for example in portions of the path curve with small radii of curvature, so that undesired local thickenings are avoided in these portions of the path curve. An open-loop and closed-loop control device undertakes the path control of the sewing head and the control of the variation of the cross-sectional geometry of the guiding device. Existing devices for producing fibre preforms for composite components by the TFP process can be retrofitted with the guiding device.

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Hamburg, Germany

Paul Joern