A method for joining structural elements to a component made of at least one fiber-reinforced plastics material, includes providing at least two structural elements made of at least one fiber-reinforced plastics material having a particular matrix; providing at least one film element made of a thermoplastic heavy-duty material; arranging the at least one film element on a surface of one of the structural elements or between surfaces to be joined of the at least two structural elements to be joined; and producing the component to be joined in a curing process.

A measurement arrangement for detecting damage to components that are made of at least one fiber-reinforced plastic material, has a plurality of temperature sensors that are arrangeable or arranged on a component at a spacing from one another. In order to provide a measurement arrangement, by means of which temperature data can be cost-effectively obtained during the production and operation of a component and for it to thus be possible for damage to the component to be recorded and monitored, the plurality of temperature sensors on the component form a sensor array and a change in the thermal material properties of the component is detected by means of the sensor array.

A reinforcing component for a structure of an aircraft or spacecraft has a first component region for reinforcing at least one other element, and at least one second component region, which is permanently connected to the first component region. The reinforcing component is formed using a thermoplastic plastics material in each of the first component region and the second component region. In the first component region, the thermoplastic plastics material forms a matrix in which continuous reinforcing fibres are embedded. In the second component region, the reinforcing component comprises discontinuous reinforcing fibres or is free of reinforcing fibres.

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.

A structural member for an aircraft or spacecraft has at least one fiber plastic composite, the fiber plastic composite having at least one or a plurality of plies. The structural member also has at least one pre-stressing means providing internal stresses to the fiber plastic composite. The fiber plastic composite and the pre-stressing means are configured and arranged to form a balanced system the internal stresses of which are essentially balanced to zero in a cured state of the said system. The balanced system of the structural member is able to counteract loading stresses exerted to the airframe in service such, that a damage caused by the loading stress is easily detectable visually.

A method of making an aircraft acoustic structural panel (10) begins with preforming a core honeycomb laminate (12) having preformed foam (3) bonded inside cells (14) thereof by a distinct adhesive (2). The preformed honeycomb laminate (12) is then stacked between opposite top and bottom structural outer laminates (16,18). The stacked honeycomb laminate (12) and outer structural laminates (16,18) are then compressed together under heat and pressure into a unitary structural panel (10) having the core honeycomb laminate (12) integrally bonded between outer skins (20,22). The outer laminates (16,18) may include imperforate acoustic septums (4) bounding the core honeycomb laminate (12) followed by an outer honeycomb (5) and structural fiber layers (6, 7, 8) defining the outer skins (20,22).

A structural composite component, in particular for an aircraft or spacecraft, has a lightning strike protection layer, and a composite battery comprising a cathode layer, wherein the lighting strike protection layer is formed integrated with the cathode layer.

A fuselage rear end of an aircraft, comprising a structural part comprising a skin and longitudinal and transversal reinforcing members and a fairing. The structural part longitudinally spans over the whole rear end and comprises a first portion in which the transversal reinforcing members occupy the whole perimeter of the corresponding fuselage section and at least a second portion in which the transversal reinforcing members occupy only a portion of the perimeter of the corresponding fuselage section. The fairing is located below the second portion of the structural part.

A rotary wing aircraft comprising a fuselage tail boom, a tail cone and an interface frame, the interface frame in turn having a connecting structure directly attached to the tail cone by means of a tail cone mechanical connection and a connecting sleeve fitting axially the fuselage tail boom, the connecting sleeve and the fuselage tail boom being directly attached by means of a one boom mechanical connection; the interface frame thereby joining the fuselage tail boom and the tail cone.