Systems and methods for printed multifunctional skins are disclosed herein. In one embodiment, an aerodynamic apparatus includes an aerodynamic structure having a first surface exposed to an outside environment, and a second surface exposed to an inside environment. A printed sensor is carried by the first surface of the aerodynamic structure, electronic components are carried by the second surface of the aerodynamic structure, and at least one printed conductive trace is carried by the first surface and the second surface. The printed conductive trace electrically connects the printed sensor with the electronics.

Systems and methods for printed multifunctional skins are disclosed herein. In one embodiment, an aerodynamic apparatus includes an aerodynamic structure having a first surface exposed to an outside environment, and a second surface exposed to an inside environment. A printed sensor is carried by the first surface of the aerodynamic structure, electronic components are carried by the second surface of the aerodynamic structure, and at least one printed conductive trace is carried by the first surface and the second surface. The printed conductive trace electrically connects the printed sensor with the electronics.

A frame component for a frame of a fuselage structure of an aircraft includes a central web extending along a longitudinal direction and having an inner edge region with respect to a radial direction running transversely with respect to the longitudinal direction and an outer edge region with respect to the radial direction. An inner web is bent from the inner edge region of the central web towards a first side. An outer web is bent from the outer edge region of the central web towards the first side. The central web, the outer web and the inner web are produced integrally from a metal sheet and together define a C-shaped cross section of the frame component. At least one stringer recess is formed in the outer web and in the outer edge region of the central web. The central web has, in the region of the stringer recess, a first reinforcing formation which forms a protrusion on the first side of the central web.

A frame component for a frame of a fuselage structure of an aircraft includes a central web extending along a longitudinal direction and having an inner edge region with respect to a radial direction running transversely with respect to the longitudinal direction and an outer edge region with respect to the radial direction. An inner web is bent from the inner edge region of the central web towards a first side. An outer web is bent from the outer edge region of the central web towards the first side. The central web, the outer web and the inner web are produced integrally from a metal sheet and together define a C-shaped cross section of the frame component. At least one stringer recess is formed in the outer web and in the outer edge region of the central web. The central web has, in the region of the stringer recess, a first reinforcing formation which forms a protrusion on the first side of the central web.

A joining system includes a holding fixture assembly configured to hold a fuselage skin. The stringers are temporarily attached to the fuselage skin. The joining system includes an upper beam assembly including an upper beam and a lower beam assembly. The lower beam assembly includes a lower beam and at least one lower heating element. The holding fixture assembly is coupled to the upper beam assembly. The holding fixture assembly is coupled to the lower beam assembly. The upper beam is movable relative to the lower beam to clamp the at least one of the plurality of stringers and the fuselage skin together prior to and during welding.

A joining system includes a holding fixture assembly configured to hold a fuselage skin. The stringers are temporarily attached to the fuselage skin. The joining system includes an upper beam assembly including an upper beam and a lower beam assembly. The lower beam assembly includes a lower beam and at least one lower heating element. The holding fixture assembly is coupled to the upper beam assembly. The holding fixture assembly is coupled to the lower beam assembly. The upper beam is movable relative to the lower beam to clamp the at least one of the plurality of stringers and the fuselage skin together prior to and during welding.

In one aspect, there is a composite skin for a tiltrotor aircraft including a first skin having a periphery defined by a forward edge, an aft edge, and outboard ends; a second skin; and a honeycomb core disposed between the first skin and the second skin, the honeycomb core comprised of a plurality of honeycomb panels positioned along the longitudinal axis of the first skin, the plurality of honeycomb panels having an array of large cells, each cell having a width of at least 1 cm; wherein the second skin and the honeycomb core have an outer perimeter within the periphery of the first skin.

In one aspect, there is a composite skin for a tiltrotor aircraft including a first skin having a periphery defined by a forward edge, an aft edge, and outboard ends; a second skin; and a honeycomb core disposed between the first skin and the second skin, the honeycomb core comprised of a plurality of honeycomb panels positioned along the longitudinal axis of the first skin, the plurality of honeycomb panels having an array of large cells, each cell having a width of at least 1 cm; wherein the second skin and the honeycomb core have an outer perimeter within the periphery of the first skin.

Described herein are methods of forming uncured sealant assemblies and also methods of forming seals between various parts using such assemblies. In some examples, an uncured sealant assembly comprises two protective layers and an uncured sealant layer, disposed in between. The uncured sealant assembly is stored and provided at a cure-inhibiting temperature, selected to minimize the curing rate of the uncured sealant layer. The size and the shape of the uncured sealant layer are specifically selected to ensure the complete coverage of the faying surfaces, filling of all gaps and voids between the faying surfaces, and controlling the shape and size of uncured sealant squeeze out between the faying surfaces. In some examples, the size and shape of the uncured sealant layer maybe be specifically selected to have no uncured sealant squeeze out between parts.

Provided are stiffened stringer panels with integrated plank structures. An example composite panel comprises a skin member having an inner surface, and a plank on the inner surface. The plank comprises a set of layered laminate plies and extends from a first side to a second side. Each laminate ply of the set of layered laminate plies is sized to form a desired geometric profile for each of the first side and the second side. The composite panel further comprises a stringer including a cap portion that spans from the first side of the plank to the second side of the plank to form a first flange portion and a second flange portion, respectively, on the inner surface of the skin member. Base segments of the cap portion conform to the desired geometric profile of respective first and second sides of the plank.