Abstract: A probe fastening system has a plate having at least one coupling-on recess and having at least one alignment element, which is arranged in the outer region of the plate, a backplate which has a fastening device for the fastening of a static probe and at least one recess which corresponds with the alignment element, at least one peelable shim which corresponds with the alignment element and which is arranged between the plate and the backplate, at least one alignment device which corresponds with the alignment element, and a coupling-on element which defines a cavity and which is designed for isobarically coupling the static probe onto the coupling-on recess.

Abstract: A primary component (100) for an aircraft (10) comprises a first surface (102) and a second surface (104) arranged opposite the first surface (102) as well as a heat exchanger (200). The heat exchanger (200) comprises a heat exchanger body (201) and a plurality of ports (212, 214, 222, 224). The heat exchanger (200) defines a first fluid circuit (210) and a second fluid circuit (220), wherein each fluid circuit is fluidically connected to two ports, respectively, and the first fluid circuit (210) is fluidically separated from the second fluid circuit (220). The heat exchanger body (201) is embedded in the primary component (100) in such a way that the heat exchanger body (201) is arranged between the first surface (102) and the second surface (104) of the primary component (100), and each of the plurality of ports (212, 214, 222, 224) is fluidically accessible either via the first surface (102) or via the second surface (104).

Abstract: Methods for producing an object by generative manufacturing, a component, and a computer-readable medium. In one method, a sheet-like part is formed with bulges. In another method, a supporting structure is formed with an arch formed by arch segments converging in the buildup direction and meet at an arch tip. In another method, a supporting structure is formed with a strut of a cloverleaf-like cross-sectional shape. In another method, a supporting structure is formed with a sheet-like part connected to the object and formed to project obliquely forward from the object in the buildup direction. In another method, a supporting structure is formed with a sheet-like part extending substantially along the buildup direction, the sheet-like part formed with apertures.

Abstract: A semi-finished product for the manufacture of a structural component has a plurality of prepreg tapes, each having unidirectionally arranged reinforcing fibers embedded in a thermoplastic matrix material, and with a plurality of connecting strands containing a thermoplastic material. The prepreg tapes and the connecting strands are either joined to form a textile sheet structure or the prepreg tapes are arranged to form a multiaxial fabric, individual layers of the fabric being joined by the connecting strands. Further, a method for manufacturing a curved structural component from such a semi-finished product is described.

Abstract: An assembly system with a sealant application station and an assembly tool is described. The assembly system is used to connect two or more components to each other. The sealant application station applies a sealant to a connector. The sealant application station is arranged in a feed system for the connectors. The sealant-applied connectors are transferred to the assembly tool. The assembly tool contains a drill spindle in which a connector and a drill can be axially aligned with each other so that the drill spindle can both drill a hole in the components to be connected and position the connector in the drilled hole with a linear movement along a processing axis. For these operations, the drill spindle moves only along the processing axis, so that a linear drive can be used for these operations.

Abstract: A method for joining a window frame to a fuselage skin of an aircraft made of a fiber composite material, includes providing the window frame, making a window opening in the fuselage skin, arranging a curable connecting layer around the window opening, pressing the window frame onto the connecting layer, hardening the composite of fuselage skin, connecting layer and window opening.

Abstract: A door frame unit for integration into a fuselage of an air vehicle includes at least one door frame component, wherein the door frame unit includes an inner frame and an outer frame surrounding the inner frame, wherein the inner frame defines a door frame opening with horizontal edges and vertical edges, in which a door is mountable, wherein a plurality of cross struts extend between the inner frame and the outer frame and are rigidly connected to the inner frame and the outer frame, and wherein the outer frame runs at an angle to the cross struts at least at areas directly neighboring corners of the inner door frame.

Abstract: A system for inserting and removing cutting inserts of a cutting tool includes a first processing device, a second processing device, and a controller coupled to the first processing device and the second processing device, wherein the first processing device includes a gripping element for holding or securing a cutting insert, wherein the second processing device includes a fastening tool for fastening or releasing a fastening mechanism holding the cutting insert, and wherein the control is adapted to control the first processing device to hold the cutting insert, to control the second processing device to fasten or release the fastening mechanism when the cutting insert is held.

Abstract: A method for producing a component having locally different thicknesses includes the steps of providing a first laminate, arranging a release layer on the first laminate, providing a second laminate with a second surface correlating with the first surface of the first laminate, placing the second laminate with the second surface on the first surface so that the release section is enclosed, joining the first with the second laminate by heating and/or pressing, guiding a cutting tool along one release section correlating with the release section, and so that the separating section is included, joining the first to the second laminate by heating and/or pressing, guiding a cutting tool along one contour in the second laminate that correlates with the separating section, so that one segment of the second laminate is completely detached, and removing the detached segment so that a recess is formed in each case.

Abstract: An apparatus for thermally joining thermoplastic fiber composite components includes a pressurization arrangement for jointly covering, at least in a region of a joining zone, thermoplastic fiber composite components to be joined and applying pressure to the thermoplastic fiber composite components to press the thermoplastic fiber composite components against one another, at least in the joining zone, the pressurization arrangement being flexible, at least in some section or sections. A welding device is configured for welding the fiber composite components in the joining zone during pressurization. The pressurization arrangement and welding device are configured to weld the thermoplastic fiber composite components in a pressurized state in the joining zone. The pressurization arrangement is configured to maintain pressurization independently of the welding device until the joining zone solidifies.

Abstract: A pressure tool for applying a pressure force during the welding of plastics components has at least one metallic first component contact section, wherein the pressure tool has at least one elastically deformable second component contact section and is thermally controllable. A Pressure apparatus has at least one such pressure tool for pressing together plastics components during welding. A method for welding plastics components is carried out by such a pressure apparatus and plastics module having at least two welded-together plastics components, wherein the at least two plastics components are welded using such a method.

Abstract: A probe fastening system has a plate having at least one coupling-on recess and having at least one alignment element, which is arranged in the outer region of the plate, a backplate which has a fastening device for the fastening of a static probe and at least one recess which corresponds with the alignment element, at least one peelable shim which corresponds with the alignment element and which is arranged between the plate and the backplate, at least one alignment device which corresponds with the alignment element, and a coupling-on element which defines a cavity and which is designed for isobarically coupling the static probe onto the coupling-on recess.

Abstract: An assembly system with a sealant application station and an assembly tool is described. The assembly system is used to connect two or more components to each other. The sealant application station applies a sealant to a connector. The sealant application station is arranged in a feed system for the connectors. The sealant-applied connectors are transferred to the assembly tool. The assembly tool contains a drill spindle in which a connector and a drill can be axially aligned with each other so that the drill spindle can both drill a hole in the components to be connected and position the connector in the drilled hole with a linear movement along a processing axis. For these operations, the drill spindle moves only along the processing axis, so that a linear drive can be used for these operations.

Abstract: An assembly system with a sealant application station and an assembly tool is described. The assembly system is used to connect two or more components to each other. The sealant application station applies a sealant to a connector. The sealant application station is arranged in a feed system for the connectors. The sealant-applied connectors are transferred to the assembly tool. The assembly tool contains a drill spindle in which a connector and a drill can be axially aligned with each other so that the drill spindle can both drill a hole in the components to be connected and position the connector in the drilled hole with a linear movement along a processing axis. For these operations, the drill spindle moves only along the processing axis, so that a linear drive can be used for these operations.

Abstract: A method of manufacturing an arc-shaped fiber composite component includes forming a preform with a planar fiber layer arrangement formed along an arc and having an outer edge assigned to a convex outer side of the arc. The outer edge is formed with gaps extending into the arrangement in such a manner that a contour of the gaps is formed at least in sections near a target contour of a respective recess to be provided in the component. The preform is formed such that a first region of the arrangement, adjacent to the outer edge and extending substantially in the direction of the arc, is bent or angled relative to a second region of the arrangement, adjacent to the first region remote from the outer side of the arc. The gaps that the preform has prior to reshaping merge into recesses of the formed preform and remain open.

Abstract: A frame component for a frame of a fuselage structure of an aircraft includes a central web extending along a longitudinal direction and having an inner edge region with respect to a radial direction running transversely with respect to the longitudinal direction and an outer edge region with respect to the radial direction. An inner web is bent from the inner edge region of the central web towards a first side. An outer web is bent from the outer edge region of the central web towards the first side. The central web, the outer web and the inner web are produced integrally from a metal sheet and together define a C-shaped cross section of the frame component. At least one stringer recess is formed in the outer web and in the outer edge region of the central web. The central web has, in the region of the stringer recess, a first reinforcing formation which forms a protrusion on the first side of the central web.

Abstract: A method for coating a metallic support profile with a layer of a corrosion-resistant material, includes providing the metallic support profile with a surface to be coated, providing a layer element of the corrosion-resistant material adapted to the surface to be coated, applying an adhesive, curable substance to at least one of the surface to be coated and the layer element, applying the layer element to the surface to be coated, covering the arrangement of support profile and layer element by a compacting element, pressing the layer element onto the surface to be coated by the compacting element to achieve a predetermined joint thickness occupied by the adhesive substance, curing the adhesive substance while maintaining the joint thickness to generate a bond, and removing the compacting element.

Abstract: A fuselage component for an aircraft, including a shell part composed of a fiber composite material and a stiffening structure which is connected to an inner surface of the shell part and has shaped parts composed of a fiber composite material, wherein the shaped parts each have a flat base section which extends along the inner surface of the shell part, and a step-shaped profile edge which extends in an encircling manner around the base section. The shaped parts are arranged in the form of parquetry with overlapping profile edges. As a result, the overlapping profile edges form stiffening profiles intersecting at an intersection. Furthermore, an aircraft which has at least one such fuselage component and a method for producing a fuselage component are described.

Abstract: Methods for producing an object by generative manufacturing, a component, and a computer-readable medium. In one method, a sheet-like part is formed with bulges. In another method, a supporting structure is formed with an arch formed by arch segments converging in the buildup direction and meet at an arch tip. In another method, a supporting structure is formed with a strut of a cloverleaf-like cross-sectional shape. In another method, a supporting structure is formed with a sheet-like part connected to the object and formed to project obliquely forward from the object in the buildup direction. In another method, a supporting structure is formed with a sheet-like part extending substantially along the buildup direction, the sheet-like part formed with apertures.

Abstract: A pressure bulkhead for the pressure-tight axial closure of a pressurized fuselage of an aircraft or spacecraft that can be put under an internal pressure, comprises a pressure dome, which is designed with a concave inner surface and a convex outer surface; inspection hatches, which each form an aperture in the pressure dome from the inner surface to the outer surface; and pressure covers for pressure-tight closure of the inspection hatches, wherein the pressure covers each have a cover and a pressure seal, wherein the cover is designed for the repeatedly releasable fixing of the pressure cover on the concave inner surface over the inspection hatch and wherein the pressure seal is formed at the edge of the cover for the pressure-tight closure of the cover with the pressure dome.