A coupling assembly includes: a tubular compression collar having a tubular body made of a resiliently flexible polymeric material and having an interior surface and an opposing exterior surface; an end of a tube disposed within a throughway of the compression collar; and a tube fitting disposed within the passageway of the tube. The tube fitting includes: a tubular stem; a flange radially outwardly projecting from an exterior surface of the stem; and an annular barb encircling and radially outwardly projecting from the exterior surface of the stem, the annular barb including a frustoconical outside face that extends along and outwardly slopes away from the stem as the outside face extends toward the flange. The compression collar radially inwardly compresses the tube against the annular barb of the tube fitting so that a liquid tight seal is formed between the tube and the tube fitting.

A wheel includes a rim and a wheel center with several spokes with at least one butt strap. The at least one butt strap protrudes from a radially outer end of the spokes. The rim includes at least one opening arranged corresponding to the at least one butt strap. The at least one butt strap extends across the related opening in the rim and is attached to the outside of the rim.

A method for forming a connection between two tubular sections having a polymeric outer surface jacket, using electrofusion to fusion bond a casing of similar, non-crosslinked polymer to the outer surface of the tubular sections.

The disclosure relates to a component device, in particular for a primary supporting component of an aircraft, the component device having a first component element, a second component element, a bonding providing a connection between the first component element and the second component element, and a detector device having at least one interior space sensor device configured to measure a change in a pressure and/or a concentration of a gas surrounding the interior space sensor device. The first component element, the second component element, and the bonding confine an interior space. The interior space sensor device is arranged in the interior space.

A method for forming a composite part of a gas turbine engine. The method includes assembling the composite part of a first composite material and a second composite material. The second composite material defines an outer surface of the composite part, and is selected to be curable at a cure temperature generated by heat from operation of the engine. The first composite material is selected to have an operating temperature limit less than the cure temperature. The method includes placing the composite part within the engine so that, in use, the second composite material is cured by exposure to the heat generated from operation of the engine. The second composite material thermally shields the first composite material from the heat generated from operation of the engine. The method includes operating the engine to cure the second composite material.

A hinged component fabrication method in which, in a layup stage of fabrication, the component includes a live hinge joining uncured material portions together at a hinge region and comprising a layer of tensile fabric at least partially infiltrated by an uncured elastomer layer at least partially interposed between the tensile fabric and the uncured material portions such that the uncured elastomer blocks the uncured material portions from infiltrating the hinge region. The method may include locating overlapped tensile fabric and elastomer layers in a tool and introducing polymer-based material into the tool such that polymer-based material portions overlap respective opposite ends of the fabric and elastomer layers. The polymer-based material portions are formed to a desired shape using the forming tool so that the fabric and elastomer layers form a live hinge between the polymer-based material portions.

A method for bonding composite substrates is disclosed. A curable surface treatment layer is applied onto a curable composite substrate, followed by co-curing. After co-curing, the composite substrate is fully cured but the surface treatment layer remains partially cured. The surface treatment layer may be a resin film or a peel ply composed of resin-impregnated fabric. If a peel ply is used, the peel ply is peeled off after co-curing, leaving behind a remaining thin film of partially cured resin. A subsequent dry physical surface treatment, such as plasma, is carried out to physically modify the surface of the surface treatment layer. After dry physical surface treatment, the composite substrate is provided with a chemically-active, bondable surface, which is adhesively bonded to another composite substrate to form a covalently-bonded structure.

A method for integrating a fitting (110) fitted between two wings of a composite profile (100) comprising the steps of: i) obtaining a preform of a fitting impregnated with a thermosetting resin, comprising two flanges (111) and a rib (114); ii) pre-curing the laminate preform of the fitting to obtain a partial polymerization of the preform; iii) laying up a laminate preform of the profile so that fibers impregnated with a thermosetting resin included in two of the wings of the laminate preform of the profile are laid up over the two flanges of the preform of the partially polymerized fitting; and iv) completely polymerizing an assembly comprising the laminate preform of the profile and the laminate preform of the partially polymerized fitting while maintaining into contact the flanges of the laminate preform of the fitting with the wings of the laminate preform of the profile.

The invention also pertains to a tooling for implementing the method and a wing spar obtained by this method.

A method for forming a composite part of a gas turbine engine. The method includes assembling the composite part of a first composite material and a second composite material. The second composite material defines an outer surface of the composite part, and is selected to be curable at a cure temperature generated by heat from operation of the engine. The first composite material is selected to have an operating temperature limit less than the cure temperature. The method includes placing the composite part within the engine so that, in use, the second composite material is cured by exposure to the heat generated from operation of the engine. The second composite material thermally shields the first composite material from the heat generated from operation of the engine. The method includes operating the engine to cure the second composite material.

Embodiments provide a method for forming a groove on a pile encapsulation jacket to form an interlocking joint on the pile encapsulation jacket. A spacer may be used to form the groove. First, a pile jacket laminate is formed on a mandrel surface. The pile jacket laminate may be partially cured to have a solid but sticky surface. The spacer may be placed on the pile jacket laminate. An additional layer of laminate may be formed on at least a portion of the pile jacket laminate and at least a portion of the spacer. The entire structure may be cured and the spacer may be removed. A groove is formed between the pile jacket laminate and the additional layer of laminate at a space previously occupied by the spacer. The groove has the same width as the portion of the spacer that was covered with the additional layer of laminate.