The disclosed invention provides a distributed directional aperture (DDA) system that is installed in a vertical lift aircraft that comprises a fuselage and a rotor system including rotary wings rotatably coupled to the fuselage. The DDA system provides capability to receive and/or transmit signals in one or more frequency bands, and provides communications, signals intelligence (SIGNINT), positional sensing, jamming, and offensive cyber on the vertical lift aircraft. The DDA system of the vertical lift aircraft includes a sensor and emitter array subsystem that includes a plurality of sensors and emitters distributed in the rotary wings, a beamformer subsystem that processes the sensor signals and emitter signals, and a telemetry subsystem that conveys signals between the sensor and emitter array subsystem and the beamformer subsystem.

An apparatus for securing first and second skins to a core in a composite rotor blade includes an elongated member configured to be installed through a passage in the core of the composite rotor blade. The elongated member has a first end configured to be attached to an outer surface of the first skin and a second end configured to be attached to an outer surface of the second skin. The apparatus also includes a first patch configured to adhere the first end to the outer surface of the first skin, and a second patch configured to adhere the second end to the outer surface of the second skin such the elongated member extends from the outer surfaces of the first and second skins through the passage in the core.

Embodiments are directed to systems and methods for two or more cured composite assemblies that are bonded together to form a tubular composite structure, wherein each of the cured composite assemblies do not have a tubular shape. The tubular composite structure may form a spar for an aerodynamic component, for example. The two or more cured composite assemblies may comprise carbon or fiberglass composite materials or a combination of materials. Each of the cured composite assemblies may further comprise axial edges that are configured to be bonded to another of the cured composite assemblies, wherein the axial edges have a sloped shape. An adhesive agent may be applied on the axial edges for bonding two cured composite assemblies. Alternatively, or additionally, one or more fasteners may be used to attach the axial edges of at least two cured composite assemblies.

A method of forming a balanced rotor blade assembly includes measuring a weight of a plurality of sub-components of the rotor blade assembly excluding a core. A configuration of a core of the rotor blade assembly is determined. In combination, the core and the plurality of sub-components achieve a target weight distribution and moment. The core is then fabricated and assembled with the plurality of sub-components to form a rotor blade sub-assembly.

Anti-icing stacks for protecting an aerodynamic surface are described. In some embodiments, an anti-icing stack includes an anti-icing layer, an elastomeric erosion protection layer, and an additional layer. The erosion protection layer is disposed between the anti-icing layer and the additional layer. The additional layer has a thickness greater than the thickness of the erosion protection layer and a tensile modulus of no more than the tensile modulus of the erosion protection layer. The additional layer may be a foam adhesive layer.

A rotor is described that comprises: a hub; a plurality of blades; each blade extends along an axis and comprises: a leading edge and a trailing edge; a top surface and a bottom surface, a chord joining points of the leading edge and the trailing edge; and a closed shell made of composite material and adapted to withstand the torque directed along a first axis of the blade; the shell comprises a first and a second element separate from each other, delimiting the respective shell on mutually opposite sides; a first connection element arranged on the side of the associated leading edge and interposed between the first ends of the respective first and second elements; and a third element connected to the associated first and second elements and arranged on the side of the leading edge; the first connection element is connected to respective first faces, and the third element is connected to respective second faces of the first and second elements; the blade comprises a fourth element interposed between the fourth element and the first connection element.

A method of forming a balanced rotor blade assembly includes measuring a weight of a plurality of sub-components of the rotor blade assembly excluding a core. A configuration of a core of the rotor blade assembly is determined. In combination, the core and the plurality of sub-components achieve a target weight distribution and moment. The core is then fabricated and assembled with the plurality of sub-components to form a rotor blade sub-assembly.

A method for fabricating a composite wing structural component for an aircraft is described. The method comprises extruding a filler material into each mold channel of a plurality of mold channels of a die to form a plurality of filler segments, removing the plurality of filler segments from the plurality of mold channels of the die, and arranging the plurality of filler segments in a space in the composite structural component, the space being defined by a radius of the composite structural component, such that the filler segments are in end-to-end contact. The method further comprises curing the plurality of filler segments in the space to fuse the plurality of filler segments.

A method for fabricating a composite wing structural component for an aircraft is described. The method comprises extruding a filler material into each mold channel of a plurality of mold channels of a die to form a plurality of filler segments, removing the plurality of filler segments from the plurality of mold channels of the die, and arranging the plurality of filler segments in a space in the composite structural component, the space being defined by a radius of the composite structural component, such that the filler segments are in end-to-end contact. The method further comprises curing the plurality of filler segments in the space to fuse the plurality of filler segments.

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.