Abstract: 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.

Abstract: A structure has a net-area-tension fastener pattern formed in the skin panel for receiving fasteners defining a net-area-tension joint for coupling a component attach fitting of a component to the skin panel. The net-area-tension fastener pattern includes two or more rows of fastener holes, including a first row and a last row. Each row is oriented generally perpendicular to a primary load direction of a load that the component is capable of exerting on the skin panel. The first row is located upstream of the last row relative to the primary load direction. The fastener holes in the first row and the last row are respectively the smallest and the largest in the net-area-tension fastener pattern. The rows are spaced apart at a spacing ratio of hole spacing to hole diameter. The spacing ratio of the first row is greater than the spacing ratio of the last row.

Abstract: Aircraft, aircraft wings and associated shear ties are disclosed. An example apparatus includes a first panel coupled to a second panel to define a wing box; a rib disposed chordwise within the wing box; and a stringer disposed spanwise within the wing box immediately adjacent at least one of the first panel or the second panel, the rib including a shear tie including first and second legs extending in opposite directions, the first and second legs to be coupled to at least one of the first panel, the second panel, or the stringer.

Abstract: A wing assembly includes a center wing structure and a pair of outer wing structures. The center wing structure includes a center wing front spar, a center wing rear spar, and an engine mounting location on each side of a wing centerline. Each outer wing structure includes an outer wing front and rear spar configured to be coupled respectively to the center wing front and rear spar to define a wing joint coupling the outer wing structure to the center wing structure. The center wing structure is configured such that the spar terminal ends of the center wing front and rear spars at each wing joint are located no further inboard than an engine centerline associated with the engine mounting location, and no further outboard of the engine centerline than ten percent of a distance between the engine centerline and the wing centerline.

Abstract: A skin panel configured for attachment to an aircraft structure is constructed of a composite material and has a continuous, smooth outer surface; a continuous, smooth inner surface that is configured for attachment to an aircraft structure; a peripheral edge that extends entirely around the skin panel and has a continuous, smooth surface and a plurality of cavities inside the material of the skin panel. The cavities reduce the weight of the skin panel without significantly detracting from the compression strength and tensile strength of the skin panel.

Abstract: Provided is a stringer having a first stiffener extending along a path and having a first flange portion and a first web portion, wherein the first stiffener is formed from a composite including a fabric reinforcement; a second stiffener extending along the path and having a second flange portion and a second web portion, the first web portion being opposed and spaced from the second web portion, wherein the second stiffener is formed from a composite including a fabric reinforcement; and an intermediate stiffener extending along the path, the intermediate stiffener being positioned between the first web portion and the second web portion, the intermediate stiffener having a first side and a second side, the first side being connected to the first web portion and the second side being connected to the second web portion, wherein the intermediate stiffener is formed from a composite including unidirectional reinforcement fibers.

Abstract: A stiffening element that includes one or more integral current flowpaths may include: a first layer including carbon-fiber-reinforced thermoplastic plies; a second layer including one or more glass-fiber-reinforced thermoplastic plies; and/or a third layer including aluminum. The third layer may form an outer surface of the stiffening element. The third layer may form at least part of the one or more integral current flowpaths. A method of using a stiffening element that comprises one or more integral current flowpaths as part of a current return network for a stiffened structure may include: selecting the stiffening element that includes a first layer including carbon-fiber-reinforced thermoplastic plies, a second layer including one or more glass-fiber-reinforced thermoplastic plies, and/or a third layer including aluminum; and/or routing current from the current return network through the one or more integral current flowpaths of the selected stiffening element.

Abstract: An apparatus including one or more panels including an outer face sheet comprising a plurality of first composite materials; an inner face sheet comprising a plurality of second composite materials; and a plurality of foam pieces disposed between the outer face sheet and the inner face sheet, wherein the foam pieces reduce warping of the panels.

Abstract: 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.

Abstract: A stiffening element that includes one or more integral current flowpaths may include: a first layer including carbon-fiber-reinforced thermoplastic plies; a second layer including one or more glass-fiber-reinforced thermoplastic plies; and/or a third layer including aluminum. The third layer may form an outer surface of the stiffening element. The third layer may form at least part of the one or more integral current flowpaths. A method of using a stiffening element that comprises one or more integral current flowpaths as part of a current return network for a stiffened structure may include: selecting the stiffening element that includes a first layer including carbon-fiber-reinforced thermoplastic plies, a second layer including one or more glass-fiber-reinforced thermoplastic plies, and/or a third layer including aluminum; and/or routing current from the current return network through the one or more integral current flowpaths of the selected stiffening element.

Abstract: A wing assembly includes a center wing structure and a pair of outer wing structures. The center wing structure includes a center wing front spar, a center wing rear spar, and an engine mounting location on each side of a wing centerline. Each outer wing structure includes an outer wing front and rear spar configured to be coupled respectively to the center wing front and rear spar to define a wing joint coupling the outer wing structure to the center wing structure. The center wing structure is configured such that the spar terminal ends of the center wing front and rear spars at each wing joint are located no further inboard than an engine centerline associated with the engine mounting location, and no further outboard of the engine centerline than ten percent of a distance between the engine centerline and the wing centerline.

Abstract: A structure has a net-area-tension fastener pattern formed in the skin panel for receiving fasteners defining a net-area-tension joint for coupling a component attach fitting of a component to the skin panel. The net-area-tension fastener pattern includes two or more rows of fastener holes, including a first row and a last row. Each row is oriented generally perpendicular to a primary load direction of a load that the component is capable of exerting on the skin panel. The first row is located upstream of the last row relative to the primary load direction. The fastener holes in the first row and the last row are respectively the smallest and the largest in the net-area-tension fastener pattern. The rows are spaced apart at a spacing ratio of hole spacing to hole diameter. The spacing ratio of the first row is greater than the spacing ratio of the last row.

Abstract: Provided is a stringer having a first stiffener extending along a path and having a first flange portion and a first web portion, wherein the first stiffener is formed from a composite including a fabric reinforcement; a second stiffener extending along the path and having a second flange portion and a second web portion, the first web portion being opposed and spaced from the second web portion, wherein the second stiffener is formed from a composite including a fabric reinforcement; and an intermediate stiffener extending along the path, the intermediate stiffener being positioned between the first web portion and the second web portion, the intermediate stiffener having a first side and a second side, the first side being connected to the first web portion and the second side being connected to the second web portion, wherein the intermediate stiffener is formed from a composite including unidirectional reinforcement fibers.

Abstract: A composite assembly includes a composite skin which includes a plurality of composite plies and a composite stringer secured to the composite skin. A first composite plank ply is positioned between a first pair of composite plies of the plurality of composite plies of the composite skin; the first composite plank ply has a width dimension less than a width dimension of the plurality of composite plies of the composite skin; and the first composite plank ply extends along a length of the composite stringer. At least a portion of the composite stringer is positioned in overlying relationship with the first composite plank ply.

Abstract: Aircraft, aircraft wings and associated shear ties are disclosed. An example apparatus includes a first panel coupled to a second panel to define a wing box; a rib disposed chordwise within the wing box; and a stringer disposed spanwise within the wing box immediately adjacent at least one of the first panel or the second panel, the rib including a shear tie including first and second legs extending in opposite directions, the first and second legs to be coupled to at least one of the first panel, the second panel, or the stringer.

Abstract: Aircraft, aircraft wings and associated shear ties are disclosed. An example apparatus includes a first panel coupled to a second panel to define a wing box; a rib disposed chordwise within the wing box; and a stringer disposed spanwise within the wing box immediately adjacent at least one of the first panel or the second panel, the rib including a shear tie including first and second legs extending in opposite directions, the first and second legs to be coupled to at least one of the first panel, the second panel, or the stringer.

Abstract: Aircraft wings and aircraft including such aircraft wings are disclosed. An example apparatus includes an aircraft wing having a first panel; a second panel; ribs coupled between the first and second panels; and stiffeners coupled between the ribs in a spanwise direction and to the first panel, the coupling between the stiffeners and the first panel to deter axial loads from being received by the stiffeners, the stiffeners to increase a compressional stability of the first panel, wherein the coupling between the stiffeners and the first panel are indirect couplings formed via clips, the couplings formed via the clips deter the axial loads from being received by the stiffeners while increasing the compressional stability of the first panel.

Abstract: An aircraft wing construction of composite materials includes a wing spar having a web constructed of only pre-preg fabric and having first and second spar caps projecting at an angle from opposite edges of the web with the first and second spar caps being constructed of interleaved and alternating portions of the pre-preg fabric of the web and unidirectional pre-preg tapes.

Abstract: Provided are stiffened stringer panels with integrated indexing stacks. An example composite panel comprises a skin member having an inner surface. An index for positioning a support tool comprises a first set and second set of layered laminate stacks coupled to the inner surface such that a channel is formed between the first and second set of layered laminate stacks. The first and second set of layered laminate stacks is configured to cradle a support tool in the channel to automatically align the support tool relative to the panel. A stringer may include a cap portion that extends from the first set to the second set of layered laminate stacks to form flange portions on the inner surface of the skin member. The index and the skin member may be co-cured or co-bonded.

Abstract: Provided are stiffened stringer panels with integrated indexing stacks. An example composite panel comprises a skin member having an inner surface. An index for positioning a support tool comprises a first set and second set of layered laminate stacks coupled to the inner surface such that a channel is formed between the first and second set of layered laminate stacks. The first and second set of layered laminate stacks is configured to cradle a support tool in the channel to automatically align the support tool relative to the panel. A stringer may include a cap portion that extends from the first set to the second set of layered laminate stacks to form flange portions on the inner surface of the skin member. The index and the skin member may be co-cured or co-bonded.