Methods for manufacturing a reinforced composite structure for an aircraft and devices used in such methods are provided. A device includes a base, a first support member fixedly attached to the base, and a second support member fixedly attached to the base and aligned longitudinally with the first support member. The first support member and the second support member are spaced a first distance apart. Two pinching wheels are spaced a second distance apart. The two pinching wheels are positioned proximate to ends of the first support member and the second support member. The second distance is less than the first distance. The two pinching wheels are configured to receive a composite material layout between them and to cause two lengths of the composite material layout to contact each other.

A fan blade platform is provided. The fan blade platform may include a wall, a first sidewall extending from the flowpath to a circular member, and a second sidewall extending from the flowpath to the circular member. A stiffening member may also extend from the circular member to the flowpath and be formed integrally with the first sidewall, the second sidewall, and the flowpath.

Tapered layers of pre-cured composite material are integrated into a tapered, highly stressed laminate structure in order to provide improved compressive strength. The pre-cured composite material can advantageously be cured under tension as pultruded material, to further augment compressive strength. The thickness of composite layers can be tapered on their termination edges by mechanically abrading, chemical abrading, or other methods. Especially preferred embodiments include aircraft structural components such as wings, wing spars, wing skins, fuselage skins, rotor blades, propellers, and propeller blades. Preferred laminates can be constructed to have at least 6, 10, 30, 50, or 100 layers of material, and can have a maximum thickness of at least 0.15, 0.25, 0.5, 1.0, or 5.0 inches.

A composite structure includes a first laminate, a second laminate, and at least one core component located between the first laminate and the second laminate. The at least one core component includes a honeycomb core and at least one locking ply extending from the honeycomb core between the first laminate and the second laminate. The at least one locking ply is configured such that a pressure applied during a cure cycle of the composite structure retains the at least one locking ply between the first laminate and the second laminate thereby preventing movement of core component.

A structure, such as a fin or water sports board, including a core having opposed first and second core outer surfaces. The structure further includes first and second fiber sheets, wherein the first fiber sheet extends along the first core outer surface and the second fiber sheet extends along the second core outer surface. A thermoplastic material impregnates the first fiber sheet and the second fiber sheet and bonds the first fiber sheet and the second fiber sheet to the core. The thermoplastic material defines an outer surface of the structure.

A composite core material and methods for making the same are disclosed herein. The composite core material comprises mineral filler discontinuous portions disposed in a continuous encapsulating resin. Further, the method for forming a composite core material comprises the steps of forming a mixture comprising mineral filler, an encapsulating prepolymer, and a polymerization catalyst; disposing the mixture onto a moving belt; and polymerizing said encapsulating prepolymer to form a composite core material comprising mineral filler discontinuous portions disposed in a continuous encapsulating resin.

Hollow composite dowel bar assemblies, their manufacture, and apparatus for manufacture. The dowel bar assemblies may include an elongate and hollow core, a protective jacket coating at least the sidewall exterior of the core, and a sealing structure coupled with each end of the combined core and jacket, that are configured to protect the core from the environment.

A cover for a cushion includes a first panel, a second panel, and a primary receptacle defined between the first panel and the second panel. The primary receptacle may receive a primary cushioning element, such as a pillow, a pillow insert, or a fill material. At least one of the first panel and the second panel includes a secondary cushioning element. The secondary cushioning element may be defined from a compressible, resilient elastomeric material that defines a plurality of thin interconnected walls that in turn define an array of open cells or columns. Each cell or column may have a hexagonal shape, imparting the array of open cells or columns with a honeycomb appearance. Methods for assembling a primary cushioning element with such a cover are also disclosed.

The invention outlines a method for producing large-sized square tubes from composite materials, incorporating a finishing process. This method addresses the need for high-quality square tube products made from composite materials, which must meet stringent requirements for lightweight design, high mechanical strength, and pressure resistance both internally and externally. The process consists of the following steps: step 1: prepare necessary equipment and materials; step 2: wind the inner layers; step 3: wrap the foam core layer over the inner layers; step 4: vacuum press the foam core layer onto the inner layers; step 5: wind the outer layers; step 6: vacuum press and heat to solidify the entire product; step 7: demould and finish the product.

An aerodynamic profile for an aircraft, comprises a core (6, 7, 8) made of foam, a skin (11, 12) defining an outer surface of the wing and cloths (14) form-ing spars and/or ribs. The profile comprises a wing, a canard, a horizontal stabiliser, a vertical stabiliser, an aileron, a flaperon, a wingtip winglet, an elevator, an elevon, a rudder or a flap.