A load structure may include a panel having a core, and a coating of polyurethane around the core. The load structure may also have a layer of carpet or felt on at least a portion of a first side of the panel. The load structure may further include an over mold coating on a second side of the panel.

A method of repairing a core stiffened structure, including removing a damaged portion of the core stiffened structure; bonding a shelf onto a first core member; bonding a second core member to a shelf; and securing a skin patch over the second core member.

A method and tool for repairing a damaged zone of an aircraft radome, using the actual radome itself as a heating chamber to avoid resorting to an autoclave and the removal of certain parts of the radome which cannot withstand the curing temperature needed for the repair. The method for repairing a radome, wherein the radome comprises a composite panel exhibiting a double curvature forming a dome, the panel comprising at least one damaged zone, comprises the steps of closing an interior space of the panel using a cap to form a repair chamber. The surface of the panel forms the repair chamber containing the damaged zone. Another step is heating the space inside the repair chamber to form a heating chamber allowing the curing expected for repairing the damaged zone.

In-line inspection and control system to in-situ monitor an extrudate during extrusion. A light beam illuminates a line on the outside circumference of the extrudate skin recording the curvature. A master profile of the illuminated defect-free skin is recorded and compared to successive monitoring of the illuminated skin. Differences from the comparison indicate skin and/or shape defects. A real-time feedback to automatically adjust process control hardware reduces or eliminates the skin and shape defects based on the monitoring and comparison.

The invention relates to a timepiece component made of composite material including at least one reinforcement and one matrix, the reinforcement having a three-dimensional honeycomb structure with a plurality of cells into which the matrix is injected. The invention also concerns a method for manufacturing such a timepiece component.

An example system includes: (i) a resin container defining a cavity; (ii) a plurality of rods extending from an inner base surface of the resin container and into the cavity; (iii) a plurality of light sources arranged to emit radiation into the plurality of rods, such that when the cavity contains liquid resin, radiation passing through a given one of the rods cures liquid resin that surrounds the given rod; and (iv) a control system configured to: (a) receive data specifying a three-dimensional structure; (b) determine a shape for a layer of a plurality of layers that collectively form the three-dimensional structure; and (c) determine one or more of the light sources that correspond to the shape of the layer; and (d) form the layer by operating the one or more determined light sources that correspond to the shape of the layer.

One aspect of a method of manufacturing a honeycomb core includes positioning a first thermoplastic columnar cell adjacent a second thermoplastic columnar cell, modifying a thermoplastic property of the first thermoplastic columnar cell, wherein the modified thermoplastic property permits joining a circumferential surface of the first thermoplastic columnar cell to a circumferential surface of the second thermoplastic columnar cell. The method also includes joining the circumferential surface of the first thermoplastic columnar cell having the modified thermoplastic property to the circumferential surface of the second thermoplastic columnar cell resulting in the honeycomb core.

A composite sandwich panel comprises a first composite skin, a second composite skin, a hollow cell core between the first composite skin and the second composite skin, and a first over-crush edge region with a first edge. The first edge has a first thickness at least 40% less than a nominal thickness of the composite sandwich panel. The first over-crush edge region has a length of at least 0.25 inches over which a thickness of the composite sandwich panel decreases.

Described are novel composite structural panels and methods of forming such panels. In some examples, a method comprises wrapping a mandrel with a composite tape to form a composite tube. This wrapping operation allows forming composite tubular structures with any cross-sectional profiles defined by the mandrel. The wrapping is also used to control the fiber orientations in the composite tubular structures. The composite tube is then cut into composite tubular structures. In some examples, the composite tube is partially cured prior to the cutting, which allows removal of the mandrel while preserving the shape of the composite tube. This cutting operation allows forming composite tubular structures with different lengths, shapes, and orientations of the ends. The composite tubular structures are disposed on a support structure and are bonded to each other. In some examples, this bonding operation also involves final curing of the composite tubular structures.

A composite panel assembly and method of forming the same includes a core, an inner skin coupled to a first side of the core, and an outer skin coupled to a second side of the core. A peripheral edge including portions of the core, the inner skin, and the outer skin, is compressed to close a path into the core.