A composite laminate for an airframe lifting surface including: at least two sides and one ramp area defined by a decreasing staggered laminate extended along a ramp direction, wherein the composite laminate includes: first plies formed by tapes arranged parallel to the ramp direction, second plies formed by tapes arranged orthogonal to the ramp direction, third plies formed by tapes arranged in a first laying up direction, being the first laying up direction different from the ramp direction and the direction orthogonal to the ramp direction, and fourth plies formed by tapes arranged in a second laying up direction, being the second laying up direction different from the ramp direction, the direction orthogonal to the ramp direction and the first laying up direction; wherein in the ramp area, the tapes forming the third and/or fourth plies are extended from one laminate side to another laminate side.

In one aspect, there is a method of making a composite skin for a tiltrotor aircraft including providing a first skin in a mold, the first skin having a periphery defined by a forward edge, an aft edge, and outboard ends; providing a plurality of honeycomb panels having an array of large cells onto the first skin, each cell having a width of at least 1 cm; assembling the plurality of honeycomb panels along the longitudinal axis of the first skin to form a honeycomb core having an outer perimeter within the periphery of the first skin; positioning a second skin onto the honeycomb core, the second skin having an outer perimeter within the periphery of the first skin; and curing an adhesive to create a bond between the first skin, the honeycomb core, and the second skin to form a composite skin.

A composite prepreg laminate such as a hat type-stringer is formed on a contoured mandrel using a combination of mechanical sweeping and vacuum forming.

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.

Methods, systems, and apparatuses are disclosed for the manufacture of composite components having incorporated reinforcing structures machined into composite material substrates, and composite components manufactured according to disclosed methods, and assemblies and larger structures comprising the composite material components.

A method, apparatus, system, and computer program product for designing a composite hollow body. A computer system selects ply layers for the hollow composite body. The ply layers comprise courses having course edges. The computer system positions the course edges between the ply layers throughout the hollow composite body to create a staggering of the course edges between the ply layers for a design of the hollow composite body.

A system and process for bonding involves a pocket made into one article is used to secure that article to another using a flowable, curable material (e.g., resin) which during saturation enters through a passageway and at least partially fills the void. When the article is cured, the article is bonded to another article to which resin has also been applied since the void (now containing cured material) is larger than the passageway.

Systems for and methods of forming structural components from a source, such as one or more rolls, of a fibrous material (e.g., fabric) are disclosed. The fibrous material is made up of multiple layers, with at least one pair of adjacent layers having different lengths.

Provided is a shaping method for shaping a laminated body of multi-layered sheet materials containing reinforcing fibers by using a shaping die. The shaping die has a curved portion formed in a convex shape over a predetermined direction. The shaping method includes: fixing, to the shaping die, a holding member configured to cover the laminated body over the predetermined direction to maintain a state where the laminated body is pressed against the curved portion; sealing the laminated body and the holding member to the shaping die by a sealing member to form a closed space; and depressurizing the closed space to thin the laminated body by sucking air of the closed space, and the fixing fixes the holding member to the shaping die such that the holding member does not come into contact with an end face on one side in the predetermined direction of the laminated body.

A method and apparatus (14) for assembling a reinforcement web (12) for use with a wind turbine blade (10) are provided. A pre-formed flange structure (20) to be integrated with laminate layers (58, 60) to form the reinforcement web (12) is clamped into position against a mould end surface (76) using one or more locating clamps (16). The locating clamps (16) include first and second clamp blocks (80, 82) that are shaped to provide an external profile that avoids resin collection and bridging during resin injection molding, while allowing for clamping to be applied to the flange structure (20) with an easily assembled and disassembled removable engagement of the clamp blocks (80, 82). The locating clamp (16) prevents undesirable dislodgment of the flange structure (20) during the assembly process for the reinforcement web (12), and without necessitating the use of complex or expensive molding equipment or processes.