An amalgamation plate for joining a first component to a second component has a planar body configured to be placed between the first component and the second component, and a plurality of first protrusions extending from a first side of the planar body. The first protrusions are radially dispersed from an axis of the planar body, and are configured to be embedded within either the first component and the second component. The amalgamation plate may also have a plurality of second protrusions extending from a second side, opposite the first side, of the planar body. The second protrusions are radially dispersed from the axis, and are configured to be embedded within the other of the first component and the second component via application of force substantially along the axis. The first protrusions and the second protrusions may be radially symmetric about the axis of the planar body.

An automated fastener insert installation system for composite panels is provided. A first module receives and secures a composite panel with respect to an origin of a first coordinate system, wherein the composite panel has opposed major surfaces and defines an insert-receiving orifice extending through one of the major surfaces, and is secured such that one of the major surfaces is externally accessible. A second module engages each of a plurality of fastener inserts with an installation aide. Third module determines a configuration of the orifice defined by the composite panel, selects a corresponding one of the fastener inserts engaged with the installation aide, inserts the selected fastener insert into the orifice, and dispenses an adhesive material through the installation aide and into the orifice about selected fastener insert such that the adhesive material secures the selected fastener insert within the orifice. Associated systems are also provided.

Coupling body devices, including methods of making the same, are provided and described. A method of manufacturing a coupling body, for example, can include forming a tube by braiding nonmetal fibers, forming a flange by rolling inwards an end of the tube, providing a metallic connector, inserting at least one Z-pin in an end portion of the connector, and coupling the tube and the connector by inserting the Z-pin in the flange and performing a composite material forming process.

A method of repairing a polymeric composite workpiece. The method comprises identifying a localized area of the polymeric composite workpiece having a defect. A plurality of three dimensional interface structures are aligned adjacent at least a portion of the localized area. The method includes applying a polymeric composite patch to the localized area such that the interface structures are disposed between the polymeric composite workpiece and the polymeric composite patch. An alternating electromagnetic field may be introduced to selectively induce localized heating of the interface structures. The localized heating softens regions of the polymeric composite workpiece and the polymeric composite patch adjacent the interface structures, causing the interface structures to penetrate a distance into the respective polymeric composite workpiece and the polymeric composite patch.

A joint including: a first component and a second component; the first component includes a bond region and an array of projections extending from the bond region, wherein the projections are embedded in the second component.

Systems and methods are provided for blade-and-slot panel affixation. One embodiment is a method. The method includes cutting a first slot into a first composite panel that includes a core between two facesheets. The first slot penetrates through the core of the first composite panel. The method also includes cutting a second slot into a second composite panel that includes a core between two facesheets. The second slot penetrates through a facesheet of the second composite panel as well as the core of the second composite panel. The method further includes inserting a blade into the first slot, and inserting the blade into the second slot.

A method of joining two panels of an airframe or fuselage structure of an aircraft or spacecraft, including: preparing an edge region of a first panel to form a first joining surface; preparing an edge region of a second panel to form a second joining surface; aligning the panels with one another such that the joining surfaces abut or interface one another forming a joint area; and joining the panels at the joining surfaces in the joint area. In an embodiment, the preparing steps include machining, and cutting, the edge regions of the first and second panels in a single operation to form the first and second joining surfaces substantially simultaneously. In another embodiment, the first and second joining surfaces are substantially planar and extend at an oblique angle with respect to a primary plane or surface of the respective first and second panels.

A method of mechanically securing a first object including a thermoplastic material in a solid state to a second object with a generally flat sheet portion, with a perforation of the sheet portion, and with the sheet portion having an edge along the perforation is provided, wherein the first object is positioned relative to the second object so that the edge is in contact with the thermoplastic material and wherein mechanical vibration energy is coupled into the assembly including the first and second objects until a flow portion of the thermoplastic material due to friction heat generated between the edge and the thermoplastic material becomes flowable and flows around the edge to at least partially embed the edge in the thermoplastic material. After the mechanical vibration stops, the thermoplastic material is caused to re-solidify, whereby the re-solidified thermoplastic material at least partially embedding the edge anchors the first object in the second object.

The invention provides a method of manufacturing a metallic-composite joint comprising a metallic component having one hole in a surface thereof, and a metallic fastener having a head portion for attaching to the metallic component and a tail portion for attaching to a composite component, comprising the steps of positioning the metallic fastener such that the head portion of the metallic fastener is retained in the hole, and the tail portion of the metallic fastener extends out of the hole, arranging layered fibers of an uncured composite component on the surface of the metallic component such that the tail portion extends through a gap between adjacent fibers in each fiber layer, and curing the composite component. The invention also provides a metallic-composite joint, a structure, an aircraft structure, an aircraft, a method of manufacturing a fastener assembly, a fastener assembly and a metallic fastener.

A method of forming a joint between a composite component and a metallic component is disclosed. The method includes forming a first groove in an outer surface of the metallic component on a first end of the metallic component. The first groove extends circumferentially on the outer surface relative to a center axis of the metallic component, and the first groove extends radially inward from the outer surface relative the center axis. The method further includes inserting the first end of the metallic component into a first end of the composite component. A second groove on the composite component over the first groove of the metallic component such that the composite component extends radially inward into the first groove of the metallic component. A composite hoop reinforcement is then applied in the second groove in a circumferential direction and the composite component, and the composite hoop reinforcement are solidified.