Abstract: A structural component for an aircraft or spacecraft, comprising: a planar member; a reinforcing member which projects from the planar member and is rigidly connected thereto; the reinforcing member comprising at least a foam layer and a cover layer, a plurality of pins extending at least through the foam layer and the cover layer, and at least the pins and the cover layer comprising a curable matrix.

Abstract: The present invention provides a composite for an aircraft or spacecraft, comprising: a skin portion; a stringer which is attached to the skin portion; and a former made from fibre composite material, which is attached to the skin portion and forms an opening therewith, through which the stringer extends, substantially transverse to the former.

Abstract: The present invention provides an arrangement, in particular for an aircraft or spacecraft, comprising an outer skin portion, the outside of which comes into contact with the atmosphere; a facing portion, which is arranged on the inside of the outer skin portion; a gap which is formed between the outer skin portion and the facing portion; and a device by means of which an airstream can be passed through the gap at a sufficient temperature to displace the dew point into the outer skin portion.

Abstract: The invention relates to a method for manufacturing a core composite (1, 20, 26) having a folded honeycomb core (2, 21, 29) provided on both sides with cover layers (3, 4, 22, 23, 27, 28) wherein the folded honeycomb core (2, 21, 29) has a number of drainage-enabling channels (5). First a core filler material (16, 17) is introduced into the folded honeycomb core (2, 21, 29) at least in some areas in order to provide the filler material with sufficient stability for the subsequent sewing process. The cover layers (3, 4, 22, 23, 27, 28) which are not yet impregnated with a curable plastics material are then placed on the folded honeycomb core (2, 21, 29) and stitched to this along the base lines (6 to 8) and/or apex lines (9, 10) at least in some sections by means of a sewing thread (19). The infiltration of the overall structure with a curable plastics material is then carried out followed by hardening by applying pressure and/or temperature.

Abstract: The present invention relates to the use of connecting elements for fastening CFK components and metal components in an aircraft or spacecraft, wherein a CFK component and a metal component can be connected to each other via at least one connecting element. The connecting element in this case is produced from an electrically non-conducting material, or has a shank and is produced from an electrically conducting or non-conducting material, wherein the shank is provided with a sleeve consisting of an electrically non-conducting material.

Abstract: The invention relates to a method for manufacturing flat, single or double curved core composites 1, 23 with at least one folded honeycomb core 4, 19. Prior to applying the initially not yet hardened cover layers 2, 3, 13, 22 a curable and later removable core filler material 15, 16 is introduced into full-length drainage-enabling channels 5, 6 of the folded honeycomb core 4, 19 in order to prevent telegraphing of the cover layers 2, 3, 13, 22 into the channels 5, 6 of the folded honeycomb core when arranging and/or hardening the cover layers 2, 3, 13, 22 and to produce edge-free and polygon-free surfaces of the core composite 1, 23. The core composites 1, 23 made according to the method have optimum structural mechanical properties, an ideal surface quality from the aerodynamic and aesthetic point of view, whereby a direct reprocessing of the core composites 1, 23 is possible without the need for further time and cost-intensive as well as in some circumstances weight-increasing finishing steps.

Abstract: The present invention provides a composite for an aircraft or spacecraft, comprising: a skin portion; a stringer which is attached to the skin portion; and a former made from fibre composite material, which is attached to the skin portion and forms an opening therewith, through which the stringer extends, substantially transverse to the former.

Abstract: A structural component for an aircraft or spacecraft, comprising: a planar member; a reinforcing member which projects from the planar member and is rigidly connected thereto; the reinforcing member comprising at least a foam layer and a cover layer, a plurality of pins extending at least through the foam layer and the cover layer, and at least the pins and the cover layer comprising a curable matrix.

Abstract: The present invention provides a fuselage arrangement for an aircraft or spacecraft, comprising: a skin portion which consists substantially of fibre composite material and/or of a fibre-composite/metal mixed material; a metal floor structure which is mounted on the skin portion; and a plurality of electrical system components which are interconnected in an electrically conductive manner by means of the floor structure.

Abstract: The present invention provides an arrangement, in particular for an aircraft or spacecraft, comprising an outer skin portion, the outside of which comes into contact with the atmosphere; a facing portion, which is arranged on the inside of the outer skin portion; a gap which is formed between the outer skin portion and the facing portion; and a device by means of which an airstream can be passed through the gap at a sufficient temperature to displace the dew point into the outer skin portion.

Abstract: The invention relates to a method for manufacturing a core composite (1, 20, 26) having a folded honeycomb core (2, 21, 29) provided on both sides with cover layers (3, 4, 22, 23, 27, 28) wherein the folded honeycomb core (2, 21, 29) has a number of drainage-enabling channels (5). First a core filler material (16, 17) is introduced into the folded honeycomb core (2, 21, 29) at least in some areas in order to provide the filler material with sufficient stability for the subsequent sewing process. The cover layers (3, 4, 22, 23, 27, 28) which are not yet impregnated with a curable plastics material are then placed on the folded honeycomb core (2, 21, 29) and stitched to this along the base lines (6 to 8) and/or apex lines (9, 10) at least in some sections by means of a sewing thread (19). The infiltration of the overall structure with a curable plastics material is then carried out followed by hardening by applying pressure and/or temperature.

Abstract: The invention relates to a method for manufacturing flat, single or double curved core composites 1, 23 with at least one folded honeycomb core 4, 19. Prior to applying the initially not yet hardened cover layers 2, 3, 13, 22 a curable and later removable core filler material 15, 16 is introduced into full-length drainage-enabling channels 5, 6 of the folded honeycomb core 4, 19 in order to prevent telegraphing of the cover layers 2, 3, 13, 22 into the channels 5, 6 of the folded honeycomb core when arranging and/or hardening the cover layers 2, 3, 13, 22 and to produce edge-free and polygon-free surfaces of the core composite 1, 23. The core composites 1, 23 made according to the method have optimum structural mechanical properties, an ideal surface quality from the aerodynamic and aesthetic point of view, whereby a direct reprocessing of the core composites 1, 23 is possible without the need for further time and cost-intensive as well as in some circumstances weight-increasing finishing steps.

Abstract: The present invention provides an inherently rigid fuselage of an aircraft or spacecraft comprising shell elements or fuselage barrels of CRP construction, wherein the shell elements or fuselage barrels each have an outer skin and an inner skin, which each at least partially or completely consist of a CRP material, wherein a composite core element is provided between the outer skin and the inner skin, and wherein the fuselage in the interior has a metal framework which is connected to the shell elements or fuselage barrels.

Abstract: The present invention creates a fuselage structural component of an aircraft or spacecraft, with a non-load-bearing outer skin and a load-bearing inner framework structure, wherein a foam layer which effects heat insulation and/or impact protection is arranged between the outer skin and the inner framework structure.

Abstract: The present invention provides a shell element as part of an aircraft fuselage, wherein the shell element is formed as a curved sheet-like element and is at least partially or completely of CRP construction.

Abstract: The present invention relates to the use of connecting elements for fastening CFK components and metal components in an aircraft or spacecraft, wherein a CFK component and a metal component can be connected to each other via at least one connecting element. The connecting element in this case is produced from an electrically non-conducting material, or has a shank and is produced from an electrically conducting or non-conducting material, wherein the shank is provided with a sleeve consisting of an electrically non-conducting material.

Abstract: A load-bearing part for aircraft or space vehicle is integrally made in a uniform manner from fiber-reinforced resin connecting endpieces of a small cross section merged via transition zones with a circular main- and central hollow. A particular method and particular tooling, for carrying out the method and making the hollow, are also disclosed.