There is provided a method that includes directing one or more infrared cameras at a compaction roller of a composite laying head of a composite layup machine. The one or more infrared cameras are mounted aft of the compaction roller. The method includes applying heat to a substrate by a heater. The heater is mounted forward of the compaction roller. The method further includes using the one or more infrared cameras, to obtain one or more infrared images of the compaction roller, during laying down of one or more composite tows of a composite layup onto the substrate by the compaction roller. The method further includes identifying, based on the one or more infrared images, one or more temperature profiles of the compaction roller, and analyzing identified temperature profiles, to determine one or more of, a layup quality of the composite layup, and a heat history of the composite layup.

Processes for forming fiber reinforced composite aircraft components, i.e., aircraft components formed of a cured fiber-reinforced resin, are provided. According to specific embodiments, a finished surface of the composite aircraft component can be achieved by providing in-mold coating of the cured fiber-reinforced resin by a dried and cured film layer of a water-based primer paint material to thereby achieve a composite aircraft component having an exterior surface that does not necessarily require further finishing.

The present disclosure is directed to a method for applying a multi-colored coating to a composite structure comprising applying a first co-curable film layer comprising a first color marking to a composite tool, applying a second co-curable film layer comprising a second color marking over the composite tool and at least partially over the first co-curable film layer to create a lay-up of a multi-colored marking, applying a composite structure over the lay-up of the multi-colored marking, and curing the lay-up of the multi-colored marking and the composite structure in a single curing step to create a cured multi-colored coating on the composite structure. A multi-colored coating for marking a composite structure and an aircraft part having a multi-colored marking are also provided.

A composite manufacturing system is provided. The composite manufacturing system comprises a fiber placement head, a compaction roller associated with the fiber placement head, and a temperature regulation system associated with the compaction roller. The temperature regulation system is configured to actively control a temperature of the compaction roller. The temperature regulation system comprises a number of temperature sensors, a cooling system, and a controller. The number of temperature sensors are configured to detect the temperature of the compaction roller. The cooling system is associated with the compaction roller and is configured to cool the compaction roller. The controller is in communication with the number of temperature sensors and the cooling system. The controller is configured to cool the compaction roller such that the temperature is below a threshold temperature.

A heated composite structure and a method for forming a heated composite structure. The structure includes carbon fibers embedded within a thermoplastic matrix. The carbon fibers are connected with first and second electrodes that are configured to be connected with an electric source such that applying current to the electrodes causes current to flow through the embedded carbon fibers to provide resistive heating sufficient to heat the composite structure to impede formation of ice on the composite structure.

The present invention relates to a blade (7) comprising: —a composite material structure (17), —a blade root fastening portion (9) further comprising a shoulder (10) extending into the recess from the wall—a base (18) arranged in the recess and comprising a support member configured to abut against the shoulder (10) of the blade root fastening portion (9) and a passage (39) formed in the support member, the sections (23) of the blade root portion (22) of the composite material structure extending through the passage (39), and—a blocking part (19) arranged in the recess between the two sections (23) of the blade root portion (22) such that each section of the blade root portion (23) is pressed against the support member by the blocking part (19).

A method eliminates voids in the bond line and manipulates a thickness of the bond line. This method includes: applying an adhesive at a center of an area of a structure; placing a doubler on the adhesive over the area of the structure, wherein the doubler has a peripheral edge; and maintaining a negative pressure at the peripheral edge of the doubler while simultaneously maintaining a positive pressure at the doubler to purge entrapped gases and control a thickness of the bond line at the peripheral edge of the doubler.

A bonded joint for use in bonding composite materials is provided and includes a composite rib having electrically conductive properties and a composite structure having electrically conductive properties. An electrically conductive preform is provided that facilitates a bond between the composite rib and the composite structure. A mesh composition that bonds the composite rib to the preform and that bonds the preform to the composite structure is provided and is electrically conductive to conduct current between the composite rib and the composite structure.

In one aspect, there is a method of making a composite skin for a tiltrotor aircraft including providing a first skin in a mold, the first skin having a periphery defined by a forward edge, an aft edge, and outboard ends; providing a plurality of honeycomb panels having an array of large cells onto the first skin, each cell having a width of at least 1 cm; assembling the plurality of honeycomb panels along the longitudinal axis of the first skin to form a honeycomb core having an outer perimeter within the periphery of the first skin; positioning a second skin onto the honeycomb core, the second skin having an outer perimeter within the periphery of the first skin; and curing an adhesive to create a bond between the first skin, the honeycomb core, and the second skin to form a composite skin.

Multilayer riblet applique and methods of producing the same are described herein. One disclosed example method includes applying a first high elongation polymer material to a web tool, where the web tool is to be provided from a first roll, and heating, via a first heating process, the first high elongation polymer material. The disclosed example method also includes applying a second high elongation polymer material to the first high elongation polymer material, and heating, via a second heating process, the second high elongation polymer material. The disclosed example method also includes applying, via a laminating roller, a support layer to the second high elongation polymer material.