An aircraft panel with a laminate structure is provided that comprises a stack of a plurality of metal sheet layers, at least one fiber reinforced adhesive layer, and at least one cover segment. At least one outer layer of the metal sheet layers comprises at least two separate metal sheets that overlap with each other along their respective commonly adjoining edges, providing an overlapping joint of the two separate metal sheets. The at least one fiber reinforced adhesive layer comprises fiber elements embedded in a matrix structure. One of the at least one fiber reinforced adhesive layers is arranged between two adjacent metal sheet layers. Further, the at least one cover segment is arranged on an outside surface of the laminate structure, the cover segment covering a region of the overlapping joint. Still further, the at least one cover segment comprises at least one layer of reinforcement fibers.

A conjugate damper for a structural panel includes a constraining sheet extending between a first edge and a second edge. Each of the first edge and the second edge is at least partially coupled to a first surface of the structural panel. The conjugate damper also includes a damping layer coupled between the constraining sheet and the first surface such that, when the structural panel is in a compressively deformed state, a thickness of the damping layer in a direction generally normal to the first surface is decreased relative to a baseline state. The damping layer includes a viscoelastic material.

A replaceable web structure for an aircraft frame is described and includes a plurality of structural members fabricated from a first type of material, the replaceable web structure comprising a plurality of web members fabricated from thin sheets of a second type of material, wherein each of the web members is fastened to a web frame of one of the structural members to replace a web portion of the one of the structural members; wherein the second type of material is lighter weight than the first type of material; and wherein the second type of material is more heat-resistant than the first material.

A self-aligning structural attachment and a corresponding method are disclosed. The self-aligning structural attachment includes a main body having a proximal end and a distal end, wherein the main body is attached to (i) a CIP of a support system and (ii) a clevis on an airframe of an aircraft. Further, the self-aligning structural attachment includes a first attachment fitting disposed on the proximal end, wherein the first attachment fitting is a slide-and-swivel attachment fitting. Still further, the self-aligning structural attachment includes a second attachment fitting disposed on the distal end, wherein the second attachment fitting is a slide attachment fitting. Yet still further, the self-aligning structural attachment includes a pivoting hinge disposed between the first attachment fitting and the second attachment fitting.

The invention relates to a thin metal sheet which is made of an alloy based on substantially recrystallized aluminum and which has a thickness of 0.25 to 12 mm, the alloy comprising, in percent by weight, Cu 3.4-4.0; Mg 0.5-0.8; Mn 0.1-0.7; Fe #0.15; Si #0.15; Zr #0.04; Ag #0.65; Zn #0.5; inevitable impurities #0.05 each and #0.15 in total, the remainder consisting of aluminum. The invention also relates to a process for manufacturing such a metal sheet and to the use thereof as a fuselage panel or sheet metal for the production of composite products such as fiber metal laminates (FML) for wing or fuselage applications in the aeronautical industry.

Applicant discloses a two part polymer mix for use in the aircraft industry that is applied with pneumatic mix and spray gun. A two part cartridge is used in the mix and spray gun so the mix is applied immediately upon mixing, but the two components are kept separated unless the gun is applying the mix. The mix cures to form a clear sealant that allows for inspection of cracks and corrosion beneath the sealant. It cures quickly so that the coated part may be further processed, for example in the assembly line during aircraft build.

A joint for a metallic skin structure includes a first end portion of the metallic skin structure and a second end portion of the metallic skin structure, wherein the first end portion and the second end portion of the metallic skin structure are secured together along a line of securement. The joint further includes at least one reinforcing fiber embedded within at least one of the first end portion of the metallic skin structure or the second end portion of the metallic skin structure extends orthogonal relative to the line of securement.

A method for producing a supporting element for a transportation device or vehicle, having steps of providing a flat base chord made of a first material having a first surface and a second surface, applying at least one layer of at least one second material to the first surface using a friction-based application or welding method, and machining the at least one layer of the second material to produce a desired profile cross section of the supporting element, wherein the first material and the second material are different metallic materials.

The method for producing an assembly is a method for producing an assembly equipped with a member to be reinforced (20), a reinforcing member body (41), and a filler (42), wherein the reinforcing member body (41) has a pair of flanges (44) arranged spaced on the surface (20B) of the member to be reinforced (20), a web (45), and a connection portion (46) which connects the flanges (44) and the web (45) and forms a filler space (V) with the surface (20B). The method for producing an assembly includes a step for inserting a filler (42) into the filler space (V), a step for attaching a crack control member (43) to cover the end of the filler (42), a step for joining the flanges (44) and the member to be reinforced (20), and a step for curing at least the member to be reinforced (20).

An aircraft fuselage includes an element forming a skin and a load-carrying structure supporting the element forming the skin. The load-carrying structure includes a plurality of elements forming a spar disposed parallel to an axial direction defined by the fuselage and a plurality of elements forming a frame disposed spaced apart along the axial direction. Each element forming a frame being arranged substantially perpendicularly to the elements forming a spar, where the element forming the skin is fastened on an external perimeter of each element forming a frame by means of fastening elements configured to keep the element forming the skin away from the external perimeter and where the element forming the skin is made of a transparent material.