Abstract: The disclosure relates to a composite panel member for an airframe of an aircraft or spacecraft, the composite panel member having a laminated or sandwich structure including: a first outer layer extending over a first side of the panel member; a second outer layer extending over a second side of the panel member; a core layer between the first and second outer layers; and at least one support element configured as an electrical conductor and provided within the core layer between the first and second outer layers. In this regard, the at least one support element extends within the core layer substantially parallel to the first and second outer layers.

Abstract: A pressure bulkhead for an aircraft fuselage having a skin element, with a first surface and a second surface, a circumferential edge and a center axis, and a reinforcement assembly, which has radially disposed reinforcement elements extending along the first surface radially between an outer end at the edge of the skin element and an inner end, which faces the center axis, wherein, in a cross section, which extends radially along the center axis, each of the radially disposed reinforcement elements defines first and second contour lines, each extending from the outer end of the radially disposed reinforcement elements to the inner end thereof, and the reinforcement assembly resting with the first contour lines thereof on the first surface of the skin element. A pressure bulkhead for an aircraft fuselage can efficiently absorb occurring pressure loads.

Abstract: An aircraft structural component having a skin element, including an inner surface, a multitude of transverse stiffening elements, which lie against the inner surface of the skin element, and a multitude of longitudinal stiffening elements, perpendicular to the transverse stiffening elements against the inner surface of the skin element, wherein between a first and a second transverse stiffening element and also between a first and a second longitudinal stiffening element, a first skin section is defined. An aircraft structural component is provided with the quantity of the longitudinal stiffening elements reduced. A first applied reinforcement includes a core layer having a contour line running along the circumference of the skin element, the first applied reinforcement including a face sheet surrounding the core layer on the side thereof which faces away from the skin element, which face sheet is connected to the skin element along the core layer contour line.

Abstract: A pressurized aircraft fuselage includes a pressure bulkhead installed on a fuselage structure, which forms a barrier between an internal pressurized cabin area and an outside area, wherein the edge of the pressure bulkhead is circumferentially attached to the fuselage structure by fastening means, wherein the fastening means include a pair of ring-shaped attachment frames made of an open or hollow profile, which are disposed at each side of the pressure bulkhead respectively, in order to form a clamping connection for the pressure bulkhead towards the fuselage structure.

Abstract: A method of producing a composite panel member, especially a composite panel member having a foam core sandwich structure for an airframe of an aircraft or spacecraft, including: providing at least one fiber reinforcement layer in a panel molding tool; moving or transferring the molding tool to an infusion station that is pre-heated to an infusion temperature and infusing the fiber reinforcement layer in the molding tool with a polymer resin; moving or transferring the molding tool to a curing station that is pre-heated to a curing temperature and curing the at least one resin-infused fiber reinforcement layer in the molding tool to form a composite panel member; and moving or transferring the molding tool to a cooling station that is provided at a cooling temperature and cooling the composite panel member in the molding tool.

Abstract: A pressure bulkhead for an aircraft fuselage having a skin element, with a first surface and a second surface, a circumferential edge and a center axis, and a reinforcement assembly, which has radially disposed reinforcement elements extending along the first surface radially between an outer end at the edge of the skin element and an inner end, which faces the center axis, wherein, in a cross section, which extends radially along the center axis, each of the radially disposed reinforcement elements defines first and second contour lines, each extending from the outer end of the radially disposed reinforcement elements to the inner end thereof, and the reinforcement assembly resting with the first contour lines thereof on the first surface of the skin element. A pressure bulkhead for an aircraft fuselage can efficiently absorb occurring pressure loads.

Abstract: 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.

Abstract: Structural component including at least one covering shell and a carrier structure, the covering shell designed in the form of a sandwich and formed by an inner skin section, an outer skin section and a shear-load absorbing core layer disposed between the two sections and connecting the inner and the outer skin sections to one another in a planar manner. The carrier structure is formed by at least one panel-shaped connecting part extending between and transversely to the inner and outer skin sections and connected along a reference longitudinal direction to the covering shell. The connecting part is located outside the carrier structure and fastened to the inner skin section resting thereon in a planar manner, wherein at least one profiled support extending along the reference longitudinal direction is provided in the covering shell in order to form a reinforcing section in the connecting region of the panel-shaped connecting part.

Abstract: An aircraft structural assembly includes an area element which has a core and an outer layer. A first surface of the area element is convexly curved, at least in certain sections, relative to an imaginary central area of the area element extending through the core of the area element. By contrast, a second surface of the area element located opposite the first surface is oriented, at least in certain sections, parallel to the imaginary central area of the area element extending through the core of the area element. The area element is configured in the form of a floor panel of a floor system. The first or the second surface of the area element is designed to form a walkable floor surface of the floor system in the state of the aircraft structural assembly when mounted in an aircraft.

Abstract: According to the invention a main-load bearing skin shell for a structural component is provided, wherein the skin shell for being affixed to a support component comprises an outer edge section with an outer edge and comprises: a connection region that does not comprise a core layer, which connection region extends along the edge with the inner skin section and the outer skin section, wherein in an end region of the core layer along the outer edge section of the skin shell reinforcement devices are integrated that project through the shear-force-absorbing core layer.

Abstract: An aircraft structural component having a skin element, including an inner surface, a multitude of transverse stiffening elements, which lie against the inner surface of the skin element, and a multitude of longitudinal stiffening elements, perpendicular to the transverse stiffening elements against the inner surface of the skin element, wherein between a first and a second transverse stiffening element and also between a first and a second longitudinal stiffening element, a first skin section is defined. An aircraft structural component is provided with the quantity of the longitudinal stiffening elements reduced. A first applied reinforcement includes a core layer having a contour line running along the circumference of the skin element, the first applied reinforcement including a face sheet surrounding the core layer on the side thereof which faces away from the skin element, which face sheet is connected to the skin element along the core layer contour line.

Abstract: 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.

Abstract: A method of producing a composite panel member, especially a composite panel member having a foam core sandwich structure for an airframe of an aircraft or spacecraft, including: providing at least one fiber reinforcement layer in a panel molding tool; moving or transferring the molding tool to an infusion station that is pre-heated to an infusion temperature and infusing the fiber reinforcement layer in the molding tool with a polymer resin; moving or transferring the molding tool to a curing station that is pre-heated to a curing temperature and curing the at least one resin-infused fiber reinforcement layer in the molding tool to form a composite panel member; and moving or transferring the molding tool to a cooling station that is provided at a cooling temperature and cooling the composite panel member in the molding tool.

Abstract: The disclosure relates to a composite panel member for an airframe of an aircraft or spacecraft, the composite panel member having a laminated or sandwich structure including: a first outer layer extending over a first side of the panel member; a second outer layer extending over a second side of the panel member; a core layer between the first and second outer layers; and at least one support element configured as an electrical conductor and provided within the core layer between the first and second outer layers. In this regard, the at least one support element extends within the core layer substantially parallel to the first and second outer layers.

Abstract: The present invention provides a structural element, in particular for an aircraft or spacecraft, comprising a core, the material density of which varies, at least in portions, in order to optimize the natural vibration behavior of the structural element. The present invention further provides a method for producing a structural element, in particular for an aircraft or spacecraft, comprising the following steps: providing a structural element which has a core; determining the natural vibration behavior of the structural element; and varying the material density of the core of the structural element, at least in portions, in such a way that the natural vibration behavior of the structural element is optimized. The present invention provides still further an aircraft or spacecraft and a rotor blade, in particular for a wind turbine, comprising such a structural element.

Abstract: The present invention provides a structural element, in particular for an aircraft and spacecraft, comprising a core, the rigidity of which varies at least in portions for optimising the aeroelastic characteristics of the structural element The present invention also provides a method for producing a structural element, in particular for an aircraft and spacecraft, which comprises the following steps: provision of a structural element comprising a core; determination of the aeroelastic behavior of the structural element; and variation, at least in portions, of the rigidity of the core of the structural element such that the aeroelastic behavior of the structural element is optimised. The present invention further provides an aircraft and spacecraft comprising a structural element of this type, and a rotor blade, in particular for a wind turbine, comprising a structural element of this type.

Abstract: An apparatus for producing a component includes a material storage tank for receiving a liquid material (M), a molding tool in which a filling region to be filled with material (M) from the material storage tank is formed, and a material supply line which connects the material storage tank to the filling region of the molding tool. In the region of the material supply line and/or the filling region of the molding tool an optical fiber has been arranged, into which at least one Fibre Bragg Grating sensor for detecting a parameter that is characteristic of the flow of material through the material supply line and/or the filling region of the molding tool is integrated.

Abstract: A pressurized aircraft fuselage includes a pressure bulkhead installed on a fuselage structure, which forms a barrier between an internal pressurized cabin area and an outside area, wherein the edge of the pressure bulkhead is circumferentially attached to the fuselage structure by fastening means, wherein the fastening means include a pair of ring-shaped attachment frames made of an open or hollow profile, which are disposed at each side of the pressure bulkhead respectively, in order to form a clamping connection for the pressure bulkhead towards the fuselage structure.

Abstract: The invention pertains to a planking panel (B) for a structural component (1) that is realized in the form of a sandwich component in its inner region (BI) that extends in a planar fashion and features a first skin section (11), a second skin section (12) and a core section (13) that is situated between these two skin sections, wherein the core section (13) connects the first and the second skin sections (11, 12) to one another in a planar fashion, with the invention being characterized in that a plurality of monitoring lines (14) is provided that extend over a planar section of the core section (13) to be monitored in order to detect damage in the core section (13) and respectively feature a first connection point (15) on a first end (14a) and a second connection point (16) on a second end (16a) in order to apply a monitoring signal (UK), wherein the monitoring lines (14) have a tear strength that lies in the range between 50% and 100% of the tear strength of the core section (13).

Abstract: According to the invention a main-load bearing skin shell for a structural component is provided, wherein the skin shell for being affixed to a support component comprises an outer edge section with an outer edge and comprises: a connection region that does not comprise a core layer, which connection region extends along the edge with the inner skin section and the outer skin section, wherein in an end region of the core layer along the outer edge section of the skin shell reinforcement devices are integrated that project through the shear-force-absorbing core layer.