Abstract: A textile fiber-composite material precursor and method for producing a component from fiber-composite material. Aircraft components can be produced from polymer fiber-composite materials, a matrix of which can be a high-performance plastics material such as polyether ketone ketone wherein a reinforcement of a non-crimp fabric of carbon fibers is embedded. Large-area non-crimp fabrics and large-area polymer films can be consolidated while being heated and pressed forming simple components. The flexible textile fiber-composite material precursor includes a stack of woven-fabric tiers from a polymer and of non-crimp fabric tiers from carbon fibers. Since both components are capable of draping, the fiber-composite material precursor can be deposited over a large area on curved shape-imparting surfaces and subsequently be consolidated under pressure and heated to form the fiber-composite material.

Abstract: A molding device for producing a shell element reinforced with supporting elements and composed of fiber composite material, including a mold part with a mold surface including a cavity for receiving a supporting element of the shell element, and a mold core to be arranged in the cavity and to support the supporting element on the shell element when the supporting element of the shell element is arranged in the cavity. The molding device enables prevention of the mold cores from falling out in an uncontrolled manner when the shell element is removed from the mold as the mold core includes a holding device to engage with the supporting element of the shell element to be produced located in the cavity and to hold the mold core on the supporting element when the shell element to be produced is released from the mold part.

Abstract: A process for producing a polymeric stiffened sheet-like component, for example a panel, for aircraft construction. Production includes integration of hollow stiffening profiles, for example closed omega stringers, onto a sheet-like component, for example an external skin, where the stringers and external skin are produced from thermoplastic composite material. The stringers are integrated onto the external skin by establishing contact between the stringers and the external skin and melting thermoplastic composite material with exposure to heat and pressure at the areas of contact between external skin and stringers. Melting of the other sections of the stringers is avoided with a pressurized cooling fluid with a temperature significantly below the melting point of thermoplastic composite material, the fluid flowing through the airtight enclosed space in the stringers.

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 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 process for producing a component for an aircraft for the frame, e.g. formers and stringers. Aircraft are being increasingly constructed of polymeric fiber composite to reduce weight. Here, fiber composites were originally composed of thermoset polymer and carbon fibers. Thermoplastic fiber composites are increasingly a research focus. An example is poly(ether ether ketone). However, production of components of thermoplastic fiber composites is complex. An improved process for producing such components includes producing a sheet-like object with a thermoplastic fiber composite having a thermoplastic polymer material and reinforcing fibers embedded therein, forming the sheet-like object for a semifinished part, and solidification of the semifinished part to give the component. Components obtainable by this process are also disclosed.

Abstract: A textile fiber-composite material precursor and method for producing a component from fiber-composite material. Aircraft components can be produced from polymer fiber-composite materials, a matrix of which can be a high-performance plastics material such as polyether ketone ketone wherein a reinforcement of a non-crimp fabric of carbon fibers is embedded. Large-area non-crimp fabrics and large-area polymer films can be consolidated while being heated and pressed forming simple components. The flexible textile fiber-composite material precursor includes a stack of woven-fabric tiers from a polymer and of non-crimp fabric tiers from carbon fibers. Since both components are capable of draping, the fiber-composite material precursor can be deposited over a large area on curved shape-imparting surfaces and subsequently be consolidated under pressure and heated to form the fiber-composite material.

Abstract: A molding device for producing a shell element reinforced with supporting elements and composed of fiber composite material, including a mold part with a mold surface including a cavity for receiving a supporting element of the shell element to be produced, and a mold core to be arranged in the cavity and to support the supporting element on the shell element when the supporting element of the shell element to be produced is arranged in the cavity. The molding device enables prevention of the mold cores from falling out in an uncontrolled manner when the shell element is removed from the mold as the mold core includes a holding device to engage with the supporting element of the shell element to be produced located in the cavity and to hold the mold core on the supporting element when the shell element to be produced is released from the mold part.

Abstract: Processes for determining carbon dioxide emission impact. In one process, a carbon dioxide emission impact is determined for each stage of a multi-stage trip. A traveler may make adjustments to the trip to offset the carbon dioxide emission impact of the entire trip and achieve a desired sustainability level for the trip. In another process, that may be used in the first process, the carbon dioxide emission impact for a vehicle is determined based on the entire life span of the vehicle, including a carbon dioxide emission impact for a production stage, a carbon dioxide emission impact for a usage stage, and a carbon dioxide emission impact for an end of life stage.

Abstract: A method for producing a fuselage portion, in particular for an aircraft or spacecraft, has the following method steps: welding a skin portion containing a thermoplastic material with a former containing a thermoplastic material in the region of a predetermined welding zone; and connecting an attachment element, configured as a crack stopper, to the skin portion and to the former in the region of the welding zone.

Abstract: A reinforcing arrangement for an opening in an aircraft fuselage structure has a fuselage skin having an inner side, an outer side and a first opening with a first opening contour, and a one-piece reinforcing structure having a thickening section and at least one projecting profile. The first opening contour is surrounded on the inner side of the fuselage skin by an opening edge surface having a thickness corresponding to a thickness of the fuselage skin surrounding the opening edge surface. The thickening section has a bearing surface matched to the opening edge surface and has a receiving section facing away from the bearing surface and surrounding a second opening. The bearing surface is in surface contact with the opening edge surface, and the reinforcing structure is connected to the fuselage skin, with the result that the first opening and the second opening lie one above the other.

Abstract: A process for producing a polymeric stiffened sheet-like component, for example a panel, for aircraft construction. Production includes integration of hollow stiffening profiles, for example closed omega stringers, onto a sheet-like component, for example an external skin, where the stringers and external skin are produced from thermoplastic composite material. The stringers are integrated onto the external skin by establishing contact between the stringers and the external skin and melting thermoplastic composite material with exposure to heat and pressure at the areas of contact between external skin and stringers. Melting of the other sections of the stringers is avoided with a pressurized cooling fluid with a temperature significantly below the melting point of thermoplastic composite material, the fluid flowing through the airtight enclosed space in the stringers.

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 rivet connection of a fiber composite component. The method includes positioning a first component which contains a fiber composite material in an overlap joint with a second component, laser-drilling a shared through-hole at least through the fiber composite material of the first component, inserting a rivet into the through-hole and fixing the rivet to the first and to the second component.

Abstract: A pressure bulkhead for an aircraft fuselage including a sandwich structure defining a central axis and extending between a circumferential border area, the sandwich structure including an inner face sheet and an outer face sheet extending transverse to the central axis, and a core assembly sandwiched between the inner face sheet and the outer face sheet. The pressure bulkhead can receive pressure loads and can be used as a structural bearing element for receiving loads introduced by other structural parts. The inner face sheet and/or the outer face sheet has a corrugated shape. The distance between the outer face sheet and the inner face sheet increases and subsequently decreases from the border area to a central area about the central axis, forming a thickened area where the distance between the outer face sheet and the inner face sheet has a peak.

Abstract: A method for integrating a backing-structure assembly in a structure of an aircraft or spacecraft is described. In this case, a plurality of individual elements is joined to form the backing-structure assembly. The individual elements for the backing-structure assembly and a skin portion for the structure are provided. The elements are arranged on a pre-assembly device which comprises retaining devices, which are each configured to hold one of the elements so as to be adjustable with respect to the position and/or location thereof. Some or all of the elements are connected to the skin portion. In the method, by adjusting the retaining devices for tolerance compensation, gaps between joint regions of the elements and the skin portion are eliminated or adjusted.

Abstract: A sandwich structure with an outer cover layer, inner cover layer, core layer, and pins extending through the core layer. The sandwich structure has a large core thickness resistant to crack formation and to detachment of the cover layers from the core layer. An anchor layer is between the outer and inner cover layers and divides the core layer, at least in areas, into an outer core layer section and inner core layer section. An outer group of pins extends through the outer core layer section and an inner group of pins extends through the inner core layer section. The outer group pins are connected to the anchor layer with their first end and to the outer cover layer with their opposite second end, and the inner group pins are connected to the anchor layer with their first end and to the inner cover layer with their opposite second end.

Abstract: A method for manufacturing a rivet connection of a fiber composite component. The method includes positioning a first component which contains a fiber composite material in an overlap joint with a second component, laser-drilling a shared through-hole at least through the fiber composite material of the first component, inserting a rivet into the through-hole and fixing the rivet to the first and to the second component.

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 reinforcing arrangement for an opening in an aircraft fuselage structure has a fuselage skin having an inner side, an outer side and a first opening with a first opening contour, and a one-piece reinforcing structure having a thickening section and at least one projecting profile. The first opening contour is surrounded on the inner side of the fuselage skin by an opening edge surface having a thickness corresponding to a thickness of the fuselage skin surrounding the opening edge surface. The thickening section has a bearing surface matched to the opening edge surface and has a receiving section facing away from the bearing surface and surrounding a second opening. The bearing surface is in surface contact with the opening edge surface, and the reinforcing structure is connected to the fuselage skin, with the result that the first opening and the second opening lie one above the other.