A panel assembly is disclosed having a panel, a beam attached to the panel, and a plurality of stiffeners attached to the panel. Each stiffener includes a respective bridge which crosses over the beam at an intersection from a first side of the beam to a second side of the beam. Each bridge has an outer surface facing away from the panel and an inner surface facing towards the panel. The inner surface of each bridge deviates away from the panel to form a recess at the intersection, and the outer surface of each bridge deviates away from the panel to form a protrusion at the intersection.

A connector for connecting a frame to a transversely extending stringer of a fuselage structure of an aircraft includes a base carrier extending in a longitudinal connector direction, a first connecting portion connected to the base carrier and extending in a first transverse connector direction, the first connecting portion having a coupling structure in an end region remote from the base carrier for engaging around an end region of a cross-section of the stringer, and a second connecting portion connected to the base carrier and adapted to be connected to the frame.

A long frame is provided with: a plurality of frame members that have a longitudinal direction and a short direction, have a first flange part folded along a folding line extending along the longitudinal direction, and are arranged close to an inner peripheral surface of a surface plate along an array direction; and a plurality of first connecting parts that connect each of the plurality of first flange parts to the surface plate. The frame is further provided with at least one of a second connecting part that connects second flange parts to each other and a third connecting part that joins together end surfaces of a pair of frame members arranged adjacent to each other in the array direction to connect the pair of frame member arranged adjacent to each other.

A process for manufacturing a metal part (10), is described said part extending in a first direction (Y) and having a section comprising a central member (14) and at least a first side member (16) extending in a second direction. The process comprises the following steps: supply of a metal blank (30); removing material from the blank so as to form an intermediate part (32) comprising the central member (14), a junction zone (12) and at least first (116) and second (118) intermediate side members, with a space (34) between the first and second intermediate side members; and hot forming (104) the intermediate part, including spreading the first and second intermediate side members apart by inserting a first punch (212) between said members.

A beam flange clamp includes a clamp body having a first vise jaw portion with a first gripping surface and a second vise jaw portion with a second gripping surface. The first gripping surface and the second gripping surface are oriented parallel to each other, are facing the same direction and are offset in height by a predetermined clamping distance. The at least one bolting hole is formed through the first vise jaw portion perpendicularly to the first gripping surface.

A beam flange clamp includes a clamp body having a first vise jaw portion with a first gripping surface and a second vise jaw portion with a second gripping surface. The first gripping surface and the second gripping surface are oriented parallel to each other and are offset in height by a predetermined clamping distance. The clamp body further including a bolting portion protruding from the second vise jaw portion opposite of the first vise jaw portion, and at least one bolting hole is formed through the bolting portion.

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 stronger than the first type of material such that one of the sheets of the second type of material is lighter weight than a sheet of equal strength of the first type of material.

An aircraft part has a fuselage portion and a floor, at least a part of which is joined to the fuselage portion via elastic link rods. Each elastic link rod is provided with a first end that is configured to be fixed to the floor and with a second end that is configured to be fixed to the fuselage portion. The elastic link rods are configured to exhibit a stiffness that makes it possible both to transmit loads between the floor and the fuselage portion and to absorb displacements brought about by deformations of the fuselage portion. The aircraft part thus making it possible, at a lower cost and with a lower mass, to obtain a connection between the floor and the fuselage portion that exhibits suitable elasticity to be able to transmit inertial loads while absorbing deformations of the fuselage portion.

A flexible corner fitting configured to contribute to a connection between at least three walls of at least one aircraft box structure, and to make them leaktight with respect to liquids. The corner fitting has at least three fixing plates and at least three edge corners made of flexible material. Each of the at least three fixing plates is connected to each of the other at least three fixing plates by one of the edge corners made of flexible material. The edge corners made of flexible material ensure adjustment of the corner fitting without machining or the addition of shims.

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.