Abstract: An apparatus is provided for an aircraft which includes an aircraft component. The aircraft component includes a first plate section, a second plate section and an elastic expansion section laterally between the first plate section and the second plate section. The elastic expansion section is configured to elastically deform to increase a lateral width between the first plate section and the second plate section. The aircraft component is formed from at least an elastomeric matrix and a fiber layer at least partially embedded within the elastomeric matrix. A bead is formed in the fiber layer laterally between a first portion of the fiber layer and a second portion of the fiber layer. The first portion of the fiber layer is disposed in the first plate section. The second portion of the fiber layer is disposed in the second plate section. The bead is disposed in the elastic expansion section.

Abstract: An assembly is provided for induction welding. This assembly utilizes a plurality of heat shields (e.g., mica heat shields) that are aligned/disposed in end-to-end relation, with each such heat shield having a recess. An induction welding coil may be disposed within a heat shield recess during induction welding operations and is movable along a welding path while proceeding along the recesses of the various heat shields. This welding path may be axially extending or may be curved. The induction welding assembly may be used to induction weld a stiffener to a curved skin or shell, for instance where a base of the recess for each heat shield is curved such that the induction welding coil may be moved along a curved welding path and while maintaining a constant spacing between the induction welding coil and the recess base of the various heat shields.

Abstract: A repair method is provided during which a thermoplastic patch is arranged on a thermoplastic aerospace component. The thermoplastic patch contacts the thermoplastic aerospace component. The thermoplastic patch is horizontal ultrasonic welded to the thermoplastic aerospace component using an ultrasonic horn.

Abstract: A door assembly for an aircraft includes a door, a door handle, a door hinge, a support assembly, and a lifting assembly. The door handle is rotatable between an unlatched position and a latched position. The support assembly rotatably mounts the door hinge to the door. The lifting assembly includes a latch cam, a lift track, and a latch idler. The latch cam is mounted to the door handle. The latch cam is rotatable with the door handle about the door handle rotational axis. The latch cam forms an eccentric roller path extending about the door handle rotational axis. The latch idler includes a first roller and a second roller. The first roller is disposed on a lift track of the support assembly and configured to roll along the lift track. The second roller is disposed within the eccentric roller path. The latch idler is mounted to the door. The latch idler is configured to lift the door as the door handle rotates from the latched position to the unlatched position.

Abstract: A fiber-reinforced thermoplastic hollow structure is provided that includes a shell and a mating component. The shell includes a first flange, a second flange, and sidewalls. The shell includes a fiber-reinforced thermoplastic material. The mating component is coupled to the first flange and the second flange and includes a first internal surface. The first internal surface and internal surfaces of the sidewalls defining a cavity. The first flange and the mating component defining a first weld seam, the first weld seam being between 20% and 80% of a lateral length of the first flange.

Abstract: A preform is provided. The preform includes first fabric layers that include a first set of commingled fibers and a second set of commingled fibers. The preform includes a second fabric layer positioned between the first fabric layers. The second fabric layer includes a third set of commingled fibers and a fourth set of commingled fibers. The first set of commingled fibers includes a lower percentage of the at least one of first carbon fibers or first fusible fibers than the third set of commingled fibers, The first set of commingled fibers include a higher percentage of first fugitive fibers than the third set of commingled fibers. The first fugitive fibers and the third fugitive fibers are pyrolyzed from the first fabric layers and the second fabric layer.

Abstract: A door assembly for an aircraft includes a door, a door hinge, and a support assembly rotatably mounting the door hinge to the door. The support assembly includes a forearm and at least one link. The forearm is rotatably mounted to the door hinge along a first rotational axis. The forearm extends along the first rotational axis between and to an upper end and a lower end. The at least one link extends between and to an inner end and an outer end. The at least one link is rotatably mounted to the forearm at the inner end. The at least one link is rotatable relative to the forearm about a second rotational axis. The at least one link is rotatably mounted to the door at the outer end. The at least one link is rotatable about the second rotational axis between and to a lowered position of the door and a raised position of the door.

Abstract: A variable area nozzle assembly includes a fixed nozzle structure, a translating sleeve, a plurality of nozzle segments, and a plurality of segment hoop locking assemblies. The translating sleeve is movably mounted to the fixed nozzle structure. The translating sleeve is translatable along the nozzle axis within the fixed nozzle structure between and to a first axial position and a second axial position. The plurality of nozzle segments form a variable area nozzle extending circumferentially about the nozzle axis. Each nozzle segment of the plurality of nozzle segments is pivotably mounted to the translating sleeve. The plurality of nozzle segments includes a first nozzle segment and a second nozzle segment. The plurality of segment hoop locking assemblies include a first segment hoop locking assembly. The first segment hoop locking assembly includes a hoop crank.

Abstract: A method is provided for forming a thermoplastic structure. During this method, a substrate is provided. The substrate includes a substrate thermoplastic and a plurality of substrate fibers disposed within the substrate thermoplastic. The substrate fibers are arranged in a substrate pattern. A first layer is deposited on the substrate using an automated fiber placement device. The first layer includes a first layer thermoplastic and a plurality of first layer fibers disposed within the first layer thermoplastic. The first layer fibers are arranged in a first layer pattern that is different than the substrate pattern. A blank is stamped into the thermoplastic structure. The blank includes the substrate and the first layer where the first layer thermoplastic is consolidated with the substrate thermoplastic.

Abstract: A method of repairing a thermoplastic composite material is disclosed herein. The method includes placing the thermoplastic composite material on a supporting anvil, placing an energy transmission material on top of the thermoplastic composite material, and applying an ultrasonic welder to the thermoplastic composite material to consolidate and repair the thermoplastic composite material. Another method of repairing a thermoplastic composite material is disclosed herein. The method includes placing an energy transmission material on a supporting anvil, placing the thermoplastic composite material on top of the energy transmission material, and applying an ultrasonic welder to the energy transmission material to consolidate and repair the thermoplastic composite material.

Abstract: A method is provided for accessing an engine core for an aircraft propulsion system. During this method, a first outer structure section of a nacelle is uncoupled from a second outer structure section of the nacelle. The second outer structure section is pivoted to an open position. A first inner structure section of the nacelle is uncoupled from a second inner structure section of the nacelle. The second inner structure section is pivoted to an open position. The first outer structure section and the first inner structure section are pivoted to an open position while the first inner structure section is structurally linked to the first outer structure section.

Abstract: An assembly for an aircraft propulsion system includes a nacelle inlet structure which extends axially along and circumferentially around a centerline. The nacelle inlet structure includes an inlet lip, an inner barrel, an outer barrel and a plurality of structure segments. The inlet lip forms a leading edge of the nacelle inlet structure. The inner barrel projects axially aft away from the inlet lip. The outer barrel projects axially aft away from the inlet lip. The outer barrel is radially outboard of and axially overlaps the inner barrel. Each of the structure segments includes an exterior skin, a mount and an electric heater configured to heat the exterior skin. The exterior skin forms a respective circumferential section of the inlet lip. The mount is connected to the exterior skin. A fastener extends radially through the mount and attaches a respective one of the structure segments to the inner barrel.

Abstract: An apparatus includes a chevron ring, a first seal and a plurality of fasteners. The chevron ring extends axially along an axis to and forms a downstream trailing edge of an exhaust nozzle. The chevron ring includes a ring base and a plurality of chevrons. The ring base includes a plurality of blind apertures. Each blind aperture projects radially inward towards the axis partially into the ring base. The fire seal is disposed radially outboard of the chevron ring. The fire seal axially and circumferentially overlaps the ring base. The fasteners mount the fire seal to the chevron ring. Each fastener includes a shank and a head. The shank projects longitudinally out from the head, through the fire seal, and is threaded into a respective one of the blind apertures. The fire seal is radially clamped between the head of each of the fasteners and the ring base.

Abstract: A manufacturing method is provided. A preform is arranged over a surface of an inner mold line. The preform is folded over sides of the inner mold line. An end of the preform is pressed into a grip strip coupled to a side of the inner mold line. The grip strip is translated in a first direction to tension the preform into a shaped body.

Abstract: An exhaust center body includes an inner skin, a perforated outer skin, a first bulkhead, a second bulkhead, a baffle and an acoustic chamber. The first bulkhead extends radially between the inner skin and the perforated outer skin. The second bulkhead extends radially between the inner skin and the perforated outer skin. The baffle includes a panel and a flange formed integral with the panel. The panel extends radially between the inner skin and the perforated outer skin. The panel extends axially along the axial centerline between the first bulkhead and the second bulkhead. The panel is attached to one of the inner skin, the first bulkhead or the second bulkhead by the flange. The acoustic chamber is formed radially between the inner skin and the perforated outer skin, axially between the first bulkhead and the second bulkhead, and circumferentially adjacent the baffle.

Abstract: A latch mechanism for joining two portions of a gas turbine engine. The latch mechanism is moveable between an open and a closed position and includes a first member configured to be attachable to a first portion of a gas turbine engine and a second member pivotally attached to the first member. The first member is attached to the second member at a first position. The latch mechanism also includes a handle pivotally attached to the second member that is attached to the second member at a second position, and third member pivotally attached to the second member. The third member is attached to the second member at a third position, the third position being between the first position and the second position, and wherein the third member comprises securing means for securing the third member to a second portion of a gas turbine engine.

Abstract: A nacelle system having a translating structure is disclosed. In various embodiments, the nacelle system includes a fixed structure; a thrust reverser having a translating sleeve configured to translate relative to the fixed structure and in response to a first dual-circuit hydraulic system; and a variable area fan nozzle having a translating nozzle configured to translate relative to the translating sleeve and in response to a second dual-circuit hydraulic system.

Abstract: A latch assembly is provided that includes first and second lateral side panels, a handle, a latch pin linkage, and a latch pin guide block. The first and second lateral side panels extend lengthwise between proximal and distal ends of the latch assembly. The handle is pivotally attached for rotation about a handle pivot axle between handle closed and handle fully open configurations. The latch pin linkage includes a drive link, a drag link, and a latch pin. The drive link is configured to stop rotation of the latch pin linkage in opposing first and second rotational directions. The drag link is pivotally attached to the drive link and the latch pin. The latch pin guide block is disposed between the first and second lateral side panels at the proximal end of the latch assembly. The latch pin guide block has an inner bore configured to receive the latch pin.

Abstract: A thrust reverser for an aircraft propulsion system includes a passage, a blocker door, an actuation linkage and a damper. The thrust reverser is configured to move the blocker door from a stowed position to a deployed position using the actuation linkage. The blocker door is located radially outboard of a flowpath relative to an axis when the blocker door is in the stowed position. The blocker door is configured to redirect air flowing in the flowpath radially outward through the passage when the blocker door is in the deployed position. The actuation linkage is pivotally coupled to the blocker door at a longitudinal distal end of the actuation linkage. The damper is disposed at the longitudinal distal end of the actuation linkage. The damper is configured to engage the actuation linkage and damp vibrations in the actuation linkage. The damper is configured from or otherwise includes elastomeric material.

Abstract: According to an aspect of the present disclosure, an ultrasonic welding tool having a longitudinal axis is provided that includes an ultrasonic tool, a tapered step horn, an ultrasonic transducer, and a plurality of helical grooves. The ultrasonic tool (UT) has a diameter, a UT length extending between a UT first end and a distal tool tip, and a UT exterior surface that extends around a UT circumference of the ultrasonic tool. The tapered step horn (SH) has an SH length that extends between an SH first end and an SH second end, a major diameter disposed at the SH first end, a minor diameter disposed at the SH second end, a middle diameter between the SH first end and the SH second end, and an SH exterior surface that extends around an SH circumference of the step horn. The SH second end of the step horn is connected to the UT first end of the ultrasonic tool. The ultrasonic transducer is in communication with the step horn and the ultrasonic tool.