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: The present disclosure pertains to a window panel for a body structure of a vehicle, especially an airframe of an aircraft or spacecraft, including a first skin which extends over a first side of the panel to form an outer skin of the vehicle body structure; a second skin which extends over a second side of the panel to form an inner skin of the vehicle body structure; and a core, especially a foam core, located between and covered by the first and second skins in a sandwich structure. The window panel includes at least one window aperture formed through the first layer, the core, and the second layer. The core may be confined to or extends over a limited extent, region or part of the panel. Thus, the first skin and/or the second skin may extend over or cover a greater area of the panel than the core.

Abstract: A pressure bulkhead of an aircraft with a fuselage shell and a pressure bulkhead disposed therein preferably in the rear region for forming a fuselage-internal pressure region, which pressure bulkhead is attached on the edge in the interior of the fuselage shell, wherein the pressure bulkhead has a lenticular cross-section and a sandwich construction that comprises a foam core sandwiched between two opposing exterior shells.

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: A curing tool for a shell component, includes a first tool component, which is a polyhedron and has a wedge-shaped bevel forming a first sheared edge on a first side edge, and a second tool component, which is a polyhedron, has a wedge-shaped bevel forming a second sheared edge on a first side edge and has a first corner bevel on a second side edge diagonally opposite the first side edge. The wedge angles of the first and the second sheared edge are equal.

Abstract: A support strut for use as a primary load-bearing structural component for the hinged coupling of at least one additional primary load-bearing structural component. The support strut includes: a rod-shaped base body extending in a longitudinal direction, a fiber composite with an annular cross section, central hole originating from the first end of the base body, and with annular holes spaced apart relative to each other in the circumferential direction of the annular cross section and originating from the first end of the base body as well as running along the longitudinal direction in the base body; roving bundles imbedded in a matrix material that fills up the respective annular hole; and a hinged connecting body with a bearing receptacle for holding a swivel joint; as well as a guiding system with such a support strut and a method for manufacturing such a support strut.

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 curing tool for a shell component, includes a first tool component, which is cuboid and has a wedge-shaped bevel forming a first sheared edge on a first side edge, and a second tool component, which is cuboid, has a wedge-shaped bevel forming a second sheared edge on a first side edge and has a first corner bevel on a second side edge diagonally opposite the first side edge. The wedge angles of the first and the second sheared edge are equal.

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 pressurized fuselage of an aircraft includes a fuselage structure and a pressure bulkhead mounted therein for forming a fuselage-internal pressure region. The pressure bulkhead is attached on the edge region, by way of mounting means, to the inside of the fuselage structure. The mounting means comprise at least one pivoting socket arrangement, arranged on the edge side of the pressure bulkhead, which edge side faces away from the pressure region, and interacts with the fuselage structure, as well as a supporting frame element, arranged on the opposite edge side and supporting itself on 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: The present disclosure pertains to a window panel for a body structure of a vehicle, especially an airframe of an aircraft or spacecraft, including a first skin which extends over a first side of the panel to form an outer skin of the vehicle body structure; a second skin which extends over a second side of the panel to form an inner skin of the vehicle body structure; and a core, especially a foam core, located between and covered by the first and second skins in a sandwich structure. The window panel includes at least one window aperture formed through the first layer, the core, and the second layer. The core may be confined to or extends over a limited extent, region or part of the panel. Thus, the first skin and/or the second skin may extend over or cover a greater area of the panel than the core.