A method of fabricating a panel for a helicopter airframe configurable in a plurality of helicopter configurations comprises the following steps. Operational characteristics of the helicopter airframe are defined. A skin panel is provided, where the skin panel is configured according to the operational characteristics of the helicopter airframe. A brace region is defined relative to the skin panel based on the operational characteristics of the helicopter airframe and the plurality of helicopter configurations. A brace assembly is operatively connected to the skin panel within the brace region to form a blank panel assembly. Accessories are arranged relative to the blank panel assembly according on one of the helicopter configurations to obtain a configured panel assembly.

An aircraft structure and its method of construction avoids the inefficiencies involved in current methods of constructing the aircraft structure, reduces the manufacturing time required for constructing the aircraft structure and reduces the cost involved in constructing the aircraft structure. The aircraft structure and its method of construction is comprised of a geodesic or multi-grid framework of intersecting thermoplastic composite strips that are joined to an interior surface of a thermoplastic composite fuselage skin panel through overmolding/co-consolidation of the grid framework and the skin panel.

An electrical arrangement on an aircraft fuselage has a fuselage component, an electrical consumer and a conductive electrical connection element. The connection element extends outwardly through an opening in the fuselage component and may have a bearing section on an inside or outside of the fuselage component that is arranged so as to tap or feed a voltage pole. The electrical connection element is electrically insulated from the opening and the fuselage component and is able to be connected to a voltage supply arranged on the inside. The electrical consumer is arranged on the outside of the fuselage component and is electrically connected to that part of the connection element projecting towards the outside of the fuselage.

The present disclosure relates to a cold spray metal process for imparting EMI resistance or lightning protection to the surface of a polymer, and a polymer with surface EMI resistance, or lightning protection, articles coated therefrom, and methods of reducing or eliminating electrochemical interactions between the metallic coating and components of the polymer.

The present disclosure relates to a cold spray metal process for imparting EMI resistance or lightning protection to the surface of a polymer, and a polymer with surface EMI resistance, or lightning protection, articles coated therefrom, and methods of reducing or eliminating electrochemical interactions between the metallic coating and components of the polymer.

A cargo aircraft configured to carry a portion of its payload in a cargo bay volume enclosed by a moveable nose door and supported by a structural extension of the fuselage that extends forward into the moveable nose door is disclosed. Examples include an aircraft with a continuous cargo bay extending from an aft end of a fixed portion into forward portion inside a nose door, where the portion of the cargo bay in the nose door is supported by a cantilevered support structure extending forward a length from a cargo opening into the fixed portion and such that the support structure extends into the nose door when the door is closed. The length of the support structure defines a supported cargo volume of at least about 15% of the overall volume of the fuselage forward of an aft end of the support structure.

Pre-impregnated composite material (prepreg) that can be cured/molded to form aerospace composite parts. The prepreg includes carbon reinforcing fibers and an uncured resin matrix. The resin matrix includes an epoxy resin component, polyethersulfone as a toughening agent, a thermoplastic particle component, a nanoparticle component and a curing agent.

Pre-impregnated composite material (prepreg) that can be cured/molded to form aerospace composite parts. The prepreg includes carbon reinforcing fibers and an uncured resin matrix. The resin matrix includes an epoxy resin component, polyethersulfone as a toughening agent, a thermoplastic particle component, a nanoparticle component and a curing agent.

A method for manufacturing a one-piece reinforced structure and obtained structure is provided, the method using base components made of partially-cured composite material and joining the base components together, applying a coating made of composite material on the base components, and applying heat on the assembly formed by the base components covered with the coating until a complete curing of the assembly is obtained, such that a one-piece reinforced structure made of composite material is obtained formed by the coating and the base components adhered to the coating, wherein the base components that form part of the very manufactured structure act as a mould during the manufacturing process, thus preventing the need to use moulds on which the composite material is deposited that must be subsequently removed from the final obtained structure.

A composite panel assembly including a first panel with a first skin, a second skin, a core, and an assembly flange. A second panel includes a first skin, a second skin, a core, and an assembly flange. A fastening device is adapted to fasten the assembly flange of the first panel to the assembly flange of the second panel, wherein the core of one of the first panel or the second panel includes reinforcing pins and a high-density portion placed along the assembly flange.