Airbus Operations Gmbh

Airbus Deutschland Gmbh

Hauke Lengsfeld

Hans Marquardt

Peter Sander

Christian Luebbering

Volker Reye

Roland Brandenburg

Stefan Tacke

Felix Caspar Helfrich

Barnaby Law

Olaf Rocker

Christian Lübbering

Rainer Graeber

Wolfgang Cruys

René Schroeder

Rudolf Duwald

Christian LÜBBERING

Michael Bechtold

In a fiber composite component for an aircraft or spacecraft which has electrically conductive fibers at least in portions, the fibers are coupled with an electrical energy source for charging with current to heat up the fibers and/or for measuring the electrical resistance of the fibers.

Fiber composite part for an aircraft or spacecraft

A method for the preparation of a dry textile preform for a large surface area fiber-reinforced composite component, in which resin-free textile material layers for purposes of manufacturing the preform are tensioned in each case in a tensioning frame, and in the stretched state are laid down on a moulding body by means of a linear relative movement between the tensioning frame and the moulding body. Also disclosed is a device for the execution of a method of this type with a tensioning frame for purposes of tensioning the material layers, and with a clamping unit for purposes of fixing the laid-down material layers on a moulding body.

Method and a device for the preparation of a dry textile preform

A method for producing a tubular structural component and device for production thereof are disclosed. A tubular molding tool has an inner mold surface shaped corresponding to an outer surface of the structural component. A support surface of an expandable support, formed corresponding to the molding tool shape such that in an unexpanded state the support fills the molding tool while the support surface and the mold surface run parallel leaving an expansion spacing of the support surface to the mold surface, is covered with a tubular non-crimp fiber fabric, such that fibers of a fiber layer of the non-crimp fiber fabric run diagonally around the support. After the support has been arranged in the molding tool, the non-crimp fiber fabric is pressed against the mold surface by expanding the support with radial expansion of the non-crimp fiber fabric. The non-crimp fiber fabric is infiltrated by a curable matrix.

Method and device for the production of tubular structural components

An apparatus for the forming of a lay-up of fiber composite material, in particular in the aerospace sector is provided. In this case, the apparatus has at least one forming tool, which has a surface contour corresponding to a desired geometry of the lay-up after the forming, and a tool carrier, to which the at least one forming tool is releasably fastened. From a further viewpoint, the invention makes available an apparatus with a modular structure, it then being possible for any desired forming tools to be combined with the tool carrier according to requirements. This leads in particular to less laborious handling, since it obviates the need for handling of the comparatively large tool carrier.

Apparatus for the forming of a lay-up of fibre composite material

A forming tool with a predefined mold portion is provided in a method for producing a reinforced fiber composite component for aviation and aerospace. A forming/support element is molded by means of the mold portion of the forming tool. A semi-finished fiber product is then deposited at least in portions on the shaped forming/support element. The deposited semi-finished fiber product is deformed by means of the forming/support element to form at least one reinforcing portion. The forming/support element is then removed from the forming tool together with the reinforcing section formed on this element, followed by defined positioning of the reinforcing section thus formed with respect to an associated fiber composite component portion by supporting with the forming/support element. The reinforcing portion is cured to form the reinforced fiber composite component. A forming device is provided with a predefined mold portion which holds a forming/support element such that it can be removed.

Method and shaping device for producing a composite fiber component for air and space travel

Methods and apparatus for manufacturing a profiled semifinished fiber-reinforced composite (FRC) product from a semifinished product of dry fiber or prepreg material that is moved through a forming device in its longitudinal direction (L), where the semifinished product is moved between the facing forming contours of tool parts in the longitudinal direction of a forming device on a carrier foil and the forming contours of the cross sections of the tool parts continuously change from a cross-sectional contour in the entry region to a forming contour in the outlet region, where a supporting part in the intake region of the forming device is set in vibration in such a way that the semifinished product is transformed into a ductile state due to the frictional heat generated between the supporting part and the semifinished product, and where the speed of the movement of the semifinished product is controlled in dependence on the temperature of the semifinished product.

Forming device for manufacturing profiled semifinished products, system with such a forming device and method for manufacturing profiled semifinished products

The present invention relates to a production method for a workpiece () composed of a fibre-composite material, having the following steps: at least one support means () is provided. A main material () composed of a pre-impregnated fibre semi-finished product and/or of a secondary material () are/is applied to the at least one support means () in order to form at least two blank-mould areas (), in such a manner that at least one of the at least two blank-mould areas () has the main material, and at least one of the blank-mould areas () has the secondary material. The blank-mould areas () are pressure-bonded to one another to form a workpiece blank () such that the at least one support means () is provided on a surface of the workpiece blank (). The at least one support means () and the secondary material () are removed from the workpiece blank () in order to form the workpiece ().

Production method for a workpiece composed of a fibre-composite material

This invention provides a method for producing a structural component, particularly in the aerospace sector. An unhardened prepreg fabric is formed from a composite fiber material, hardenable at a first hardening temperature, into a predetermined shape. The unhardened prepreg fabric is then connected to at least one hardening part of a composite fiber material partially hardened at a second hardening temperature for holding the unhardened prepreg fabric in the predetermined shape, wherein the second hardening temperature is lower than the first hardening temperature. In a further method step the unhardened prepreg fabric connected to the at least one holding part is hardened at the first hardening temperature to form the structural component. The idea on which this invention is based consists in replacing expensive tools of special steel previously used for holding dimensionally unstable prepreg fabrics with at least one holding part of composite fiber materials, wherein the at least one holding part has a lower hardening temperature than the prepreg fabric, and can therefore be produced with a lower cost tool.

Method for producing a structural component

The invention relates to a method for joining cured stringers to at least one structural component of an aircraft or spacecraft, including placing the stringers onto the at least one structural component, depositing and pressing respective vacuum film strips such that each strip is attached to a pair of stringers and present between the pair of stringers with the use of vacuum sealing means in such a manner that a continuous vacuum arrangement, comprising the covering vacuum films and the vacuum film strips, is formed, and pressurized joining of the stringers to the at least one structural component by means of the continuous vacuum arrangement formed.

Method for joining cured stringers to at least one structural component of an aircraft or spacecraft

The present invention provides a method for producing a profile from composite fiber material, in particular in the aviation and aerospace field, having the following steps: A preliminary fabric, particularly made of pre-impregnated fiber material, is first packed in a vacuum bag. Then support elements are laid on the packed preliminary fabric to support the latter. The vacuum bag is then supplied with a vacuum. Next, the preliminary fabric is hardened to the profile under the action of heat, in particularly in an autoclave. With the claimed method the support elements can be formed advantageously from favorable material, in particular aluminum, instead of highly expensive nickel-36 steel, since the support elements are uncoupled mechanically from the preliminary fabric in a longitudinal direction by means of the vacuum bag, enabling movement of the support elements in the longitudinal direction relative to the preliminary fabric.

Method for producing a profile from composite fiber material

The invention relates to a method for joining precured stringers to at least one structural component of an aircraft or spacecraft. A vacuum arrangement required for the joining is produced in two parts. In a first step, each precured stringer is covered in advance by a covering vacuum film. The stringers prepared in this manner are arranged on the structural component. Vacuum film strips are subsequently arranged in each case on adjacent stringers and over an intermediate space between the adjacent stringers. With the use of a vacuum sealing means, the vacuum film strips and the covering vacuum films 8 form a continuous vacuum arrangement. The stringers are subsequently joined under pressurization to the structural component with the use of this vacuum arrangement.

Method for joining a profiled first structural component to a second structural component

The present invention provides a method as well as a tool () for producing a T-stringer, in particular in the aircraft and spacecraft sector. The method is characterized by two lay-up strips () being arranged on top of one another, wherein areas () of the lay-up strips () corresponding to the stringer foot () that is to be formed, are spaced apart from one another by means of separation films () and areas () of the lay-up strips () corresponding to a stringer web () come into direct contact with one another, in particular, forming a material bond. In a further method step, the lay-up strips () are clamped at the areas () corresponding to the stringer web () that is to be formed. In an even further step, the areas () corresponding to the stringer foot () that is to be formed are folded open in opposite directions () by means of a folding-open mandrel (), and by means of a pressing roller () the areas () are pressed against corresponding pressing surfaces () of the clamping device () such that the areas () are properly flush. Finally, the folded lay-up strips () are cured to form the T-stringer that is to be produced. The method according to the invention has the advantage over the known method for producing T-stringers that a smaller number of steps is required to produce the T-stringer and can be automated easily.

Method for producing a T-stringer

A connecting arrangement for connecting two reinforcing elements of an aircraft or spacecraft, wherein the reinforcing elements have different cross-sectional profiles with at least one foot portion and at least one comb portion, at least one foot portion connecting element which on one side can be adapted to the geometric shape of the foot portion of the first reinforcing element and on the opposite side can be adapted to the geometric shape of the foot portion of the second reinforcing element and can be connected rigidly thereto; and at least one comb portion connecting element which on one side can be adapted to the geometric shape of the comb portion of the first reinforcing element and on the opposite side can be adapted to the geometric shape of the comb portion of the second reinforcing element and can be connected rigidly thereto.

Connecting arrangement for joining two stiffening elements having different cross-sectional profiles for an aircraft or spacecraft, and shell component

The invention relates to a method for determining at least one characteristic parameter of a CRP specimen (), in particular a specimen of prepreg material, for aerospace, comprising the following method steps: presenting the specimen (), irradiating the. specimen () with a predetermined spectrum of electromagnetic radiation, recording the interaction between the specimen () and the electromagnetic radiation in a data record () and determining the at least one characteristic parameter from the recorded data record ().

Method for determining a characteristic parameter of a CRP specimen

The present invention relates to a method for producing profile parts, each having an L-shaped cross-section, or an assembled profile part having, in particular, a T-shaped cross-section. For this purpose a lay-up made of composite fiber material is initially placed on a forming tool. In a further step, the lay-up is deformed by the forming tool to form a profile having a preferably U-shaped cross-section, the profile thus exhibiting the book effect at each opposite end. In a further step, the profile is preferably cut in the longitudinal direction in order to produce two profile parts. The idea on which the invention is based is to allow the layers in the lay-up to shift during the deformation process and subsequently to produce a substantially right-angled chamfered end on the profile parts by means of a cutting process. The disadvantageous formation of corrugations and complex clamping of the lay-up are thus avoided.

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Helmste, Germany

Hauke Lengsfeld