A light-weight stiffened wing airfoil includes at least one wing segment (52). The wing segment comprises an upper (53b) and a lower skin assembly (53a), wherein each of the upper and lower skin assemblies incorporates a plurality of inwardly facing stringers (74); a first rib (64-1) at a distal end of the wing segment and a second rib (64-2) at a proximal end of the wing segment; a plurality of rib trusses (70) extending from the first and second ribs to the opposing skin assembly (53); and a plurality of support members extending from the inwardly facing stringers to the opposing skin assembly.

Methods aim to reduce and/or eliminate the need for shims in manufacturing assemblies, such as in manufacturing of aircraft wings. Exemplary methods include predicting a set of predicted manufacturing dimensions within a range of predetermined allowances for a first part, manufacturing the first part, scanning the first part to determine a set of actual manufacturing dimensions for the first part, and at least beginning manufacturing a second part before the scanning the first part is completed. The second part may be manufactured based on the set of predicted manufacturing dimensions for the first part. Once the scan of the first part is completed, the set of predicted manufacturing dimensions may be compared to a set of actual manufacturing dimensions to check for any non-compliant deviances between the predicted and actual manufacturing dimensions. Repairs and local re-scans may be performed in the areas of the non-compliant deviances, which may streamline manufacturing.

A jet engine support structure includes an inboard support fitting that is configured to be operatively attached to the jet engine, an outboard support fitting that is configured to be operatively attached to the jet engine, an inboard longeron that is configured to be attached to the inboard support fitting and an exterior underside surface of an aircraft wing, an outboard longeron that is configured to be attached to the outboard support fitting and the exterior underside surface of the wing, and a drag brace fitting that is configured to be attached to the inboard longeron and the outboard longeron and operatively attached to the jet engine. These component parts are employed in operatively attaching the jet engine to the aircraft wing. The same set of component parts is employed in operatively attaching the jet engine to either a left side or port side wing or to operatively attach a jet engine to a right side or starboard side wing. The construction of the longerons enable their attachment to the underside surface of the wing using fasteners already used in the wing construction.

Provided are stiffened stringer panels with integrated shim structures. An example composite panel comprises a skin member having an inner surface, and a stringer positioned on the inner surface. The stringer comprises a first flange portion and a second flange portion with the first flange portion and the second flange portion contacting the inner surface. The composite panel further comprises an integrated shim positioned on the first flange portion. The integrated shim may comprise a sacrificial material configured to be trimmed to the desired geometry. The sacrificial material may comprise a glass fiber reinforced plastic fabric pre-impregnated with resin. The integrated shim may be configured to interface with internal support structures of a wing assembly. The integrated shim may be configured to be machined to fit dimensions and measurements corresponding to the internal support structures of the wing assembly.

A panel including an outer face sheet; an inner face sheet; a foam disposed between the outer face sheet and the inner face sheet; and a core structure (e.g., a stringer comprising a hat structure) between the foam and the inner face sheet; and between the foam and the outer face sheet. The core structure is protected from impact damage and increases flexural stiffness of the panel on an aircraft wing.

The present invention provides an apparatus and method for twisting a wing rib of an aircraft that when deployed across the wing span allows for a wide range of wing shape variations. This variance in shape may be used to steer the airplane without the use of flaps, and change the wings from a high-speed, low-lift shape to a low-speed, high-lift shape, including interim wing configurations, during flight to increase efficiency. The apparatus utilizes high strength-to-weight ratio polymer artificial muscles wrapped in heating wire as the rib twisting actuators. Wing rib twist is accomplished by electrifying the heating wire of the appropriate polymer artificial muscle to alter the wing rib twist. The wing rib apparatus includes a venting design that allows for faster activation of the wing rib twist by using ambient air convection to accelerate cooling of the relaxing polymer artificial muscle.

There is provided a laminated composite structure having improved impact damage resistance and improved strength. The laminated composite structure has a plurality of stacked layers of a composite material. The plurality of stacked layers have one or more interlaminar corrugations formed within the plurality of stacked layers. Each interlaminar corrugation has a substantially sinusoidal shaped profile, and has a depth and a length dependent on a size of the laminated composite structure formed. The laminated composite structure with the one or more interlaminar corrugations has improved strength and improved impact damage resistance at an exposed edge of the laminated composite structure, when the exposed edge is subjected to an impact force.

A composite structure forming system configured to form a contoured elongate composite structure in a continuous process is presented. The composite structure forming system comprises a plurality of charge forming stations and a plurality of conveyor systems. The plurality of charge forming stations is configured to operate in parallel, each charge forming station of the plurality of charge forming stations is configured to form a respective composite charge of the contoured elongate composite structure. Each conveyor system of the plurality of conveyor systems is configured to transport a respective composite charge through a respective charge forming station.

A fiber composite component having a first and a second fiber composite element each bent along a transverse axis opf the fiber composite component to have, respectively, in succession, a first and second base flange, a first and second web section, a first and second top flange and a first and second stiffening web. Respectively, the first and second base flanges are parallel to the first and second top flanges, the first and second web sections are angled with respect to each of the first and second base flanges and the first and second top flanges, the first and second stiffening webs are at right angles with respect to the first and second top flanges, and the first stiffening web and the second stiffening web are congruent with respect to one another, and are connected to one another, along a longitudinal axis of the fiber composite component.

Prior to curing a composite workpiece assembly, an expandable element can be inserted into a cavity of the workpiece assembly. The expandable element is configured to expand when a predetermined change is produced in an attribute of the element. The attribute can be a temperature of the element. The element is expanded by producing the predetermined change, and the workpiece assembly is cured while the expanded element is in the cavity, so that the expanded element applies positive pressure to inner surfaces of the cavity during curing. The expanded element can be removed from the cavity after curing. The expanded element can comprise a plurality of expandable pellets.