Abstract: An arrangement includes a fiber-reinforced composite component or assembly and a monitoring device. The composite component or assembly includes at least first and second reinforcing fiber formation sections stitched to each other using a yarn to connect the fiber formation sections along a seam, the yarn being electrically conductive along a length thereof. The monitoring device is configured and coupled to the yarn so as to be capable of sending an electrical input signal along at least a section of the yarn that forms the seam or part thereof and receiving a response signal on the yarn. Furthermore, an aircraft or spacecraft is provided including at least one arrangement of this type, as well as a method of producing an arrangement including a fiber-reinforced composite component or composite assembly, and a method of monitoring the structural integrity of a fiber-reinforced composite component or assembly.

Abstract: A method for bonding two fiber composite components with each other to form a fiber composite structure includes integrating conductive fibers underneath a bonding surface of at least one of the two fiber composite components, each conductive fiber comprising a carbon fiber coated with an electrically insulating coating, the conductive fibers running along the bonding surface and protruding at least at their ends from the respective fiber composite component; arranging the two fiber composite components against each other at their respective bonding surfaces; passing an electric current through the conductive fibers by electrically contacting the conductive fibers at their protruding ends so that the respective fiber composite component is heated at the bonding surface to a curing temperature; and joining the two fiber composite components with each other at their bonding surfaces via secondary bonding, co-bonding and/or co-curing at the curing temperature, thereby forming the fiber composite structure.

Abstract: A fuselage structure for an aircraft has a frame structure, which extends in a circumferential direction and has a first connecting section and a second connecting section, which is arranged spaced apart from the first connecting section, a reinforcing structure, which extends between the first and the second connecting sections and is connected respectively to the first and the second connecting sections, an outer shell, which is secured on an outer side of the reinforcing structure, and an inner shell, which is secured on an inner side of the reinforcing structure. The reinforcing structure has a profiled cross section when viewed in a longitudinal direction extending transversely to the circumferential direction and, together with the inner shell and the outer shell, forms a plurality of channels, which are adjacent to one another in the circumferential direction and each extend in the longitudinal direction.

Abstract: A method for manufacturing a fuselage component for an aircraft includes building up, on a surface of a two-dimensional shell part composed of a fibre composite material, stiffening profiles crossing at an intersection from a plurality of layers, superimposed on one another in a thickness direction, which are each formed by applying filaments, which each comprise a fibre bundle embedded in a thermoplastic material. Stabilising layers, within which the filaments extend beyond the intersection, are formed from the filaments in each of the stiffening profiles. A fuselage component and an aircraft, which comprises at least one such fuselage component, are furthermore described.

Abstract: 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.

Abstract: A joining method for connecting at least two thermoplastic workpieces is provided to permit the joining even of non-transparent carbon fiber reinforced plastics parts by means of laser welding, in which a splice is produced at the edge regions of the workpieces and the workpieces are subsequently positioned relative to one another in such a manner that the opposite splice regions bound a seam region. Connecting bodies are then inserted into the seam region and heated by means of local heat input by laser beam such that a fixed integrally bonded connection forms between the workpieces and the connecting bodies.

Abstract: 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.

Abstract: A planar pressure bulkhead for an air- or spacecraft, including a first planar wall designed as a pressure seal for a fuselage structure, a second planar wall arranged opposite the first wall, and a reinforcing structure which is arranged between the first and the second wall and which includes at least one circumferential monolithic reinforcing element. Furthermore, the present disclosure provides an air- or spacecraft, including a fuselage structure and a planar pressure bulkhead that is coupled to the fuselage structure in a circumferentially continuous manner.

Abstract: Disclosed is a method for manufacturing a fiber composite member. The method involves the step of producing a fiber layer by depositing at least two fiber webs in a plane adjacent to each other and along a circular guiding arc. The fiber webs are deposited on curved deposition paths that intersect the circular guiding arc with the same predetermined angle, and are formed in the form of portions of circular involutes.

Abstract: A method for connecting two aircraft components composed of thermoplastic composite material by a rivet. The two aircraft components are arranged, in sections, areally on one another, before bores in the two aircraft components are expanded by a mandrel element such that the bores, after the expansion, form one continuous bore for receiving the rivet. Here, the two aircraft components are locally warmed by the mandrel element such that, during the expansion of the bores, the components are thermoplastically deformed. The mandrel element is inserted through the first bore into the second bore. The rivet is introduced into the continuous bore. A system for carrying out the method is also disclosed.

Abstract: A method for repairing a component constructed in multiple layers from a layer composite material comprises a step of removing material of the layer composite material of the component in the region of damage to the component in order to form a scarfed cutout, such that a scarfed peripheral contour formed from a plurality of layers of the component is formed. Furthermore, fiber-reinforced plastic strips, which exhibit a curable plastic material in an uncured state, are applied in layers in the cutout in order to create a replacement-structure semifinished product that is complementary to and butts against the scarfed peripheral contour of the component. In a further step, the curable plastic material contained in the fiber-reinforced plastic strips is cured in order to create a replacement structure and for joining the replacement structure to the component.

Abstract: A joining method for connecting at least two thermoplastic workpieces is provided to permit the joining even of non-transparent carbon fiber reinforced plastics parts by means of laser welding, in which a splice is produced at the edge regions of the workpieces and the workpieces are subsequently positioned relative to one another in such a manner that the opposite splice regions bound a seam region. Connecting bodies are then inserted into the seam region and heated by means of local heat input by laser beam such that a fixed integrally bonded connection forms between the workpieces and the connecting bodies.

Abstract: 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.

Abstract: A method for manufacturing a fuselage component for an aircraft includes building up, on a surface of a two-dimensional shell part composed of a fibre composite material, stiffening profiles crossing at an intersection from a plurality of layers, superimposed on one another in a thickness direction, which are each formed by applying filaments, which each comprise a fibre bundle embedded in a thermoplastic material. Stabilising layers, within which the filaments extend beyond the intersection, are formed from the filaments in each of the stiffening profiles. A fuselage component and an aircraft, which comprises at least one such fuselage component, are furthermore described.

Abstract: A fuselage structure for an aircraft has a frame structure, which extends in a circumferential direction and has a first connecting section and a second connecting section, which is arranged spaced apart from the first connecting section, a reinforcing structure, which extends between the first and the second connecting sections and is connected respectively to the first and the second connecting sections, an outer shell, which is secured on an outer side of the reinforcing structure, and an inner shell, which is secured on an inner side of the reinforcing structure. The reinforcing structure has a profiled cross section when viewed in a longitudinal direction extending transversely to the circumferential direction and, together with the inner shell and the outer shell, forms a plurality of channels, which are adjacent to one another in the circumferential direction and each extend in the longitudinal direction.

Abstract: A method for forming a structural component for an airframe of an aircraft or spacecraft, includes: providing a prefabricated shell element comprising a thermoplastic substrate; and applying a stiffening structure to the shell element by additive manufacturing, wherein a plurality of continuous thermoplastic filaments filled with reinforcing fibers are continuously heated and three dimensionally formed such that the filaments are crossing and bonded to each other at a plurality of crossing points to form a three dimensional grid truss integrally formed on the shell element. A corresponding structural component and an aircraft or spacecraft including such a structural component are also described.

Abstract: A planar pressure bulkhead for an air- or spacecraft, including a first planar wall designed as a pressure seal for a fuselage structure, a second planar wall arranged opposite the first wall, and a reinforcing structure which is arranged between the first and the second wall and which includes at least one circumferential monolithic reinforcing element. Furthermore, the present disclosure provides an air- or spacecraft, including a fuselage structure and a planar pressure bulkhead that is coupled to the fuselage structure in a circumferentially continuous manner.

Abstract: A method for resistance welding of two fiber-composite components to give a fiber-composite structure includes arranging conductive fibers within a jointing region of the two fiber-composite components, where each conductive fiber includes a carbon fiber with an electrically insulating coating. An electric current is passed through the conductive fibers to heat the jointing region to a welding temperature and melt the fiber-composite components in the jointing region. The jointing region is hardened in a manner that bonds the two fiber-composite components by way of the jointing region to give the fiber-composite structure.

Abstract: A method for connecting two aircraft components composed of thermoplastic composite material by a rivet. The two aircraft components are arranged, in sections, areally on one another, before bores in the two aircraft components are expanded by a mandrel element such that the bores, after the expansion, form one continuous bore for receiving the rivet. Here, the two aircraft components are locally warmed by the mandrel element such that, during the expansion of the bores, the components are thermoplastically deformed. The mandrel element is inserted through the first bore into the second bore. The rivet is introduced into the continuous bore. A system for carrying out the method is also disclosed.

Abstract: A method for repairing a component constructed in multiple layers from a layer composite material comprises a step of removing material of the layer composite material of the component in the region of damage to the component in order to form a scarfed cutout, such that a scarfed peripheral contour formed from a plurality of layers of the component is formed. Furthermore, fiber-reinforced plastic strips, which exhibit a curable plastic material in an uncured state, are applied in layers in the cutout in order to create a replacement-structure semifinished product that is complementary to and butts against the scarfed peripheral contour of the component. In a further step, the curable plastic material contained in the fiber-reinforced plastic strips is cured in order to create a replacement structure and for joining the replacement structure to the component.