A method of examining the integrity of an aircraft structure including determining an electrical conductivity or resistivity of the metal matrix composite of the aircraft structure. An apparatus for performing such a method is also provided. A method of estimating damage in an aircraft structure and a method of estimating the remaining operational life of an aircraft structure are also provided.

Systems and methods are provided for curing a composite part. The method includes the steps of: disposing a heat blanket at a composite material; applying, with a controller, power to heaters distributed across multiple cells of a heat blanket to heat the composite material at the heat blanket; monitoring, with the controller, a temperature of the composite material at each of the multiple cells via thermocouples distributed across the multiple cells; and individually adjusting, with the controller, an amount of power applied to the heaters, for each of the multiple cells, in response to the monitored temperature and a target temperature.

In summary, the present application relates to an aircraft interior paneling (10) comprising:—a planar replaceable surface element (40);—a structural element (20);—at least one connection element (30); wherein the replaceable surface element (40) is detachably fastened to the structural element (20) by means of the at least one connection element (30), and wherein the replaceable surface element (40) comprises:—a support element (42), wherein, in the thickness direction of the support element (42), the support element (42) comprises a plurality of successive layers which comprise woven fabrics;—a decorative element (50) which is attached to a decorative connection surface (52) of the support element (42), wherein the support element (42) is dimensionally stable; and a replaceable surface element (40) for an aircraft interior paneling (10); the use of a replaceable surface element (40) for an aircraft interior paneling (10); a method for producing a replaceable surface element (40) for an aircraft interior paneling (10); and a method for updating an existing aircraft interior paneling (100).

Provided herein is a method to reinforce a junction of fiber-composite aircraft components forming an elongate conduit. The method comprises (a) abrading the junction and (b) applying sealant along the junction from a tool moveable through the conduit, the tool exerting force on an interior surface of the conduit opposite the junction.

A method for producing a structural section of a vehicle comprises the steps of providing multiple separate skin panels of a fiber-reinforced plastic having an inner side, an outer side and a border running peripherally around the respective skin panel; arranging at least one stiffening component of a fiber-reinforced plastic on each skin panel, on the respective inner side; integrally connecting the respective at least one stiffening component to the skin panels concerned to form a structural component; arranging at least two structural components on a carrier, so that at least regions of the borders of the structural components concerned are in surface-area contact; and integrally connecting the regions of the borders that are in surface-area contact to one another.

We describe a way of detecting delamination of a laminated structure that is heated by DC powered heaters by passing the source and return wires that supply current to the heaters through a toroidal transformer core. Should there be a breakdown in the laminations, current flowing through the heater will flow into the structure, resulting in less current being present in the return wire than in the source wire. The current imbalance between the source and return wires causes the transformer core to saturate. Using the core saturation, caused by the DC current imbalance, the delamination (or imminent delamination) of the laminated structure can be detected.

The present embodiment relates to a composite fibre structure (100) and a method (200) of manufacturing the composite fibre structure (200). The composite fibre structure (100) includes a core (102) and an outer layer (108) enclosing the core (102). The core (102) further includes at least one of a permanent core (104) and a temporary core (106). The permanent core (104) is 3-D printed along with the temporary core (106) to form the core structure (102). The permanent core (104) and the temporary core (106) are placed alternatively along the section, extending throughout the length of the composite fibre structure (100), or the permanent core (104) and temporary core (102) can be alternate along the length of the composite fibre structure (100). The layer (108), made of a reinforcement material, wraps the core (102) to form the composite fibre structure (100).

A stiffener is disclosed including a core and a shell which surrounds the core. The shell is formed from a fibre material, and the core includes first and second battens arranged side by side, and a foam spacer between the battens. The stiffener extends in a lengthwise direction, and the battens and the foam spacer have respective lengths which extend in the lengthwise direction of the stiffener. The core is assembled with the spacer between the battens, then surrounded with the shell. The stiffener may be a stringer for an aircraft wing.

Various aerospace vehicle systems and methods are disclosed. In one embodiment, a fuel efficient, low emissions aerospace vehicle includes a fuselage having a fineness ration of equal to or greater than 8. The fuselage is comprised of at least 50% composite materials. The aerospace vehicle also includes a first wing, a second wing, and a third wing coupled to the fuselage, each wing having an aspect ratio of equal to or greater than 35. The wings each have a span within 10% of one another and an aspect ratio within 10% of one another. Each wing is comprised of at least 50% composite materials. The aerospace vehicle also includes at least one stabilizing unit coupled to the fuselage. The stabilizing unit includes first and second stabilizer surfaces configured in a V-tail configuration. The aerospace vehicle further includes at least one propulsion system.

A mechanical fastener system for EME protection, including at least one plated component, wherein the at least one plated component includes a base material, a bonding layer disposed over the base material, and a metal plating disposed over the bonding layer.