The Boeing Company

Patrick Bickford Stickler

Forouzan Behzadpour

Alexander Rubin

Steven Paul Walker

William John Koch, Iv

Gary D Oakes

Benjamin Joseph Palmer

David Wesley Trop

Douglas Alan McCarville

Juan C Guzman

Gregory M Santini

Joseph L Sweetin

Jennifer Sue Noel

Jerry Deming Chungbin

Ian E Schroeder

John Makoto Shinozaki

Jason H Inouye

Douglas A Mccarville

Jennifer S Noel

A laminate structure may include: an aluminum layer; a glass composite layer adjacent to the aluminum layer; and/or a carbon composite layer adjacent to the glass composite layer, opposite to the aluminum layer. The glass composite layer may include one or more glass-fiber-reinforced thermoplastic prepreg plies. The carbon composite layer may include one or more carbon-fiber-reinforced thermoplastic prepreg plies. A laminate structure may include: a first aluminum layer; a first glass composite layer adjacent to the first aluminum layer; a first carbon composite layer adjacent to the first glass composite layer, and opposite to the first aluminum layer; and/or a second glass composite layer adjacent to the first carbon composite layer, and opposite to the first glass composite layer. The first glass composite layer may include one or more glass-fiber-reinforced thermoplastic prepreg plies. The first carbon composite layer may include one or more carbon-fiber-reinforced thermoplastic prepreg plies.

Methods of making laminate structures

A material system may include: an aluminum layer; a glass composite layer adjacent to the first aluminum layer; and a carbon composite layer adjacent to the first glass composite layer, and opposite to the first aluminum layer. A method of manufacturing a material system may include: stacking an aluminum layer, glass composite layer that may include thermoplastic prepreg plies, and carbon composite layer so that the aluminum layer is adjacent to the glass composite layer, and the glass composite layer is adjacent to the carbon composite layer; and consolidating the thermoplastic prepreg plies to soften the aluminum layer. A method of manufacturing a material system may include: stacking an aluminum layer, glass composite layer that comprises thermoplastic resin, and carbon composite layer so that the glass composite layer is between the aluminum and carbon composite layers; and adjusting temperature and pressure to consolidate the stack.

Methods of manufacturing material systems

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.

Methods of using stiffening elements that comprise one or more integral current flowpaths

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.

All new multi-purpose internally stiffened composite wing sub-panel

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.

Structure having net-area-tension joint

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.

Aircraft, aircraft wings and associated shear ties

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.

Wing assembly having wing joints joining outer wing structures to center wing structure

A composite hat stringer for stiffening a panel includes a plurality of composite fabric plies arranged to form a cap, a pair of flanges and a pair of webs respectively connecting the cap with the pair of flanges. The cap includes at least one 0&#xb0; composite tape ply interleafed in the composite fabric plies within the cap.

Composite fabric hat stringers having interleafed tape plies

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.

Skin panel of composite material having an internal grid

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.

Stringer and associated composite structure and method for reinforcing a base structure

A method, apparatus, and system for managing a composite part. An acceptable level of a warpage is identified for the composite part. The warpage for the composite part is a change in the composite part as manufactured that deviates from design specifications for the composite part. Orientations in a stacking sequence for plies in the composite part are selected that result in the composite part having the acceptable level of the warpage and a desired strength to form selected orientations. The composite part is manufactured using the selected orientations.

Composite structure warpage control system

Stiffening elements that comprise integral current flowpaths

Material systems and methods of manufacturing material systems

Laminate structures comprising fiber-reinforced thermoplastic prepreg plies

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.

Growing community of inventors

Everett, WA, United States of America

Patrick Bickford Stickler