Abstract: In a method for reinforcing a fiber composite component for aviation and space flight, a vacuum mat is configured with at least one receiving portion for reproducibly receiving at least one reinforcing element. The at least one reinforcing element is introduced into the at least one receiving portion of the vacuum mat. The vacuum mat with at least one introduced reinforcing element is applied reproducibly in a sealed manner to the fiber composite component to be reinforced to form a mold portion and at least the formed mold portion is cured to connect the at least one reinforcing element to the fiber composite component. The vacuum mat is then removed from the reinforced fiber composite component so that the vacuum mat can be reused.

Abstract: The present invention creates a tape laying device as well as a method for the construction of a laminate by means of a tape laying device, in particular in the field of aviation and aerospace. The tape laying device comprises a laying arrangement for laying tape for the construction of a laminate and a stimulation arrangement by means of which the tape which has to be laid can be stimulated directly and evenly over the thickness thereof so as to be heated. The idea the present invention is based on consists in that instead of the stimulation at the surface of the tape which has to be laid which is known to the Applicant a direct and evenly stimulation of the tape which has to be laid takes place, that means, a stimulation which acts substantially direct and evenly over the thickness of the tape which has to be laid on the tapes components. The heating achieved by this is much faster than by means of convection according to the known approach.

Abstract: The present invention provides a device, in particular within the aeronautical and aerospace fields, comprising a first and second tool and guide means, which guide the first and second tools in such a way that they press a preliminary fabric arranged in a pressing zone of the device along a second axis as said fabric moves along a first axis through the pressing zone, said second axis being substantially transverse to the first axis, so as to form a consolidated preliminary fabric, the guide means guiding the first and second tools parallel to one another, in part, along the first axis in the pressing zone.

Abstract: The invention relates to a method for joining precured stringers to at least one structural component of an aircraft or spacecraft, including depositing covering vacuum films in advance at least on a predetermined region of each of the stringers to be joined, 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 in such a manner that a continuous vacuum arrangement, including 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 the continuous vacuum arrangement formed.

Abstract: A forming tool with a predefined mould portion is provided in a method for producing a reinforced fibre composite component for aviation and aerospace. A forming/support element is moulded by means of the mould portion of the forming tool. A semi-finished fibre product is then deposited at least in portions on the shaped forming/support element. The deposited semi-finished fibre 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 fibre composite component portion by supporting with the forming/support element. The reinforcing portion is cured to form the reinforced fibre composite component. A forming device is provided with a predefined mould portion which holds a forming/support element such that it can be removed.

Abstract: The invention relates to a method for joining precured stringers to at least one structural component of an aircraft or spacecraft. A vacuum arrangement necessary for the joining is produced in two parts. In a first step, each precured stringer is covered in advance with a covering vacuum film. The stringers prepared in this manner are arranged on the structural component. Respective vacuum film strips are subsequently arranged 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.

Abstract: Disclosed is a method for producing a profiled semifinished fiber composite (FC) product from a semifinished product (2) that is made of dry fiber or prepreg material and is moved through a shaping device (1) in the longitudinal direction (L) of the semifinished product (2). The semi-finished product (2) is moved through the shaping device (1) on a support film in the longitudinal direction (L) of the shaping device (1), between the facing shaping contours of tool parts (11a, 12a). The shaping contours of the cross-sections of the tool parts (11a, 12a) continuously change from a cross-sectional contour in the feeding zone to a shaping contour in the discharge zone. Furthermore, the support piece in the drawing-in zone of the shaping device (2) is made to oscillate such that the semifinished product is transformed into a shapeable state as a result of the frictional heat generated between the support piece and the semifinished product.

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

Abstract: The present invention creates a tape laying device as well as a method for the construction of a laminate by means of a tape laying device, in particular in the field of aviation and aerospace. The tape laying device comprises a laying arrangement for laying tape for the construction of a laminate and a stimulation arrangement by means of which the tape which has to be laid can be stimulated directly and evenly over the thickness thereof so as to be heated. The idea the present invention is based on consists in that instead of the stimulation at the surface of the tape which has to be laid which is known to the Applicant a direct and evenly stimulation of the tape which has to be laid takes place, that means, a stimulation which acts substantially direct and evenly over the thickness of the tape which has to be laid on the tapes components. The heating achieved by this is much faster than by means of convection according to the known approach.

Abstract: A connecting arrangement for connecting two reinforcing elements of an aircraft or spacecraft, wherein the reinforcing elements have different cross-sectional profiles with in each case 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 in each case.

Abstract: In a method for reinforcing a fibre composite component for aviation and space flight, a vacuum mat is configured with at least one receiving portion for reproducibly receiving at least one reinforcing element. The at least one reinforcing element is introduced into the at least one receiving portion of the vacuum mat. The vacuum mat with at least one introduced reinforcing element is applied reproducibly in a sealed manner to the fibre composite component to be reinforced to form a mould portion and at least the formed mould portion is cured to connect the at least one reinforcing element to the fibre composite component. The vacuum mat is then removed from the reinforced fibre composite component so that the vacuum mat can be reused.

Abstract: 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.

Abstract: The present invention provides a method for the production of a tubular structural component. A tubular moulding tool having an inner mould surface which is shaped corresponding to an outer surface of the structural component is provided. A support surface of an expandable support which is formed such that in an unexpanded state, it fills the moulding tool, leaving an expansion spacing of the support surface to the mould surface, is covered with a tubular fibre-woven fabric. After the support has been arranged in the moulding tool, the fibre-woven fabric is pressed against the mould surface by expanding the support and is infiltrated by a curable matrix. From another point of view, a device is provided for the production of a tubular structural component.

Abstract: The present invention provides a method as well as a tool (1, 48) for producing a T-stringer, in particular in the aircraft and spacecraft sector. The method is characterized by two lay-up strips (3, 4) being arranged on top of one another, wherein areas (5, 6) of the lay-up strips (3, 4) corresponding to the stringer foot (60) that is to be formed, are spaced apart from one another by means of separation films (7, 8) and areas (14, 15) of the lay-up strips (5, 6) corresponding to a stringer web (61) come into direct contact with one another, in particular, forming a material bond. In a further method step, the lay-up strips (3, 4) are clamped at the areas (14, 15) corresponding to the stringer web (61) that is to be formed.

Abstract: The invention relates to a method for joining an uncured stringer to a structural component of an aircraft or spacecraft. The uncured stringer is supplied with at least one joining section and at least one web section. A rapidly curing low-temperature auxiliary material is deposited on web flanks of the at least one web section, which are to be supported. The at least one joining section of the uncured stringer is fitted to the structural component. Curing of the auxiliary material then takes place at a first curing temperature in order to form cured supporting elements for the web flanks of the at least one web section, which are to be supported. Curing of the uncured stringer is undertaken at a second curing temperature which is greater than the first precuring/curing temperature, and then the deposited and cured auxiliary material is removed.

Abstract: This invention provides a method for producing a structural component, particularly in the aerospace sector. An unhardened prepreg fabric is formed from a composite fibre 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 fibre 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.

Abstract: 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 fibre 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.

Abstract: The invention relates to a method for determining at least one characteristic parameter of a CRP specimen (3, 4), in particular a specimen of prepreg material, for aerospace, comprising the following method steps: presenting the specimen (3, 4), irradiating the. specimen (3, 4) with a predetermined spectrum of electromagnetic radiation, recording the interaction between the specimen (3, 4) and the electromagnetic radiation in a data record (20) and determining the at least one characteristic parameter from the recorded data record (20).

Abstract: The present invention provides a device, in particular within the aeronautical and aerospace fields, comprising a first and second tool and guide means, which guide the first and second tools in such a way that they press a preliminary fabric arranged in a pressing zone of the device along a second axis as said fabric moves along a first axis through the pressing zone, said second axis being substantially transverse to the first axis, so as to form a consolidated preliminary fabric, the guide means guiding the first and second tools parallel to one another, in part, along the first axis in the pressing zone.

Abstract: The present invention relates to a production method for a workpiece (15) composed of a fibre-composite material, having the following steps: at least one support means (3) is provided. A main material (6) composed of a pre-impregnated fibre semi-finished product and/or of a secondary material (7) are/is applied to the at least one support means (3) in order to form at least two blank-mould areas (1, 10), in such a manner that at least one of the at least two blank-mould areas (1, 10) has the main material, and at least one of the blank-mould areas (1, 10) has the secondary material. The blank-mould areas (1, 10) are pressure-bonded to one another to form a workpiece blank (14) such that the at least one support means (3) is provided on a surface of the workpiece blank (14). The at least one support means (3) and the secondary material (7) are removed from the workpiece blank (14) in order to form the workpiece (15).