A vehicle body may have an internal skeleton forming a wing shape, and a skin formed over the internal skeleton. The skin may include a matrix material, and a plurality of continuous fibers encased within the matrix material. The plurality of continuous fibers may curve from a base end near a fore/aft center of the wing shape outward toward leading and trailing edges of the wing shape at a tip end.

A spar structure can connect a moveable component to an aircraft. The structure is formed from a single continuous body of material defining a plurality of attachment/actuation brackets and a pair of continuous hinge lines extending through the body.

A method of manufacturing a spar for a propeller blade comprises forming a hollow inner core from a composite material, and then braiding carbon fibres around the inner core to form an outer spar structure. The hollow inner core may be formed using an inflatable bladder and a mould defining an outer shape of the spar. This construction avoids the need for a foam core to be provided within the spar.

Methods aim to reduce and/or eliminate the need for shims in manufacturing assemblies, such as in manufacturing of aircraft wings. Exemplary methods include predicting a set of predicted manufacturing dimensions within a range of predetermined allowances for a first part, manufacturing the first part, scanning the first part to determine a set of actual manufacturing dimensions for the first part, and at least beginning manufacturing a second part before the scanning the first part is completed. The second part may be manufactured based on the set of predicted manufacturing dimensions for the first part. Once the scan of the first part is completed, the set of predicted manufacturing dimensions may be compared to a set of actual manufacturing dimensions to check for any non-compliant deviances between the predicted and actual manufacturing dimensions. Repairs and local re-scans may be performed in the areas of the non-compliant deviances, which may streamline manufacturing.

Disclosed is a wing in which, while a continuous surface is maintained, a chord length and camber of an airfoil may be modified via only a rotational drive alone, whereby a structure is simple and aerodynamic efficiency may be improved.

A wing leading-edge device includes a flow body having a front side, a back side, a plurality of ribs arranged in ribs, wherein at least one of the ribs is a load introduction rib having at least one first lug for coupling with a drive mechanism, a second load path component, which includes at least one second lug, wherein the second load path component is at least connected to the load introduction rib, such that a second opening of the at least one second lug is co-axial with a first opening of the at least one first lug.

Structural composite airfoils include a primary structural element and a secondary structural element defining a trailing edge of the structural composite airfoil. The primary structural element includes an upper skin panel, a lower skin panel, and a middle C-channel spar that is coupled to the upper skin panel and the lower skin panel. An upper flange of the middle C-channel spar is coupled to the upper skin panel, while a lower flange of the middle C-channel spar is coupled to the upper skin panel and the lower skin panel. An internal volume is defined by the upper skin panel and the lower skin panel aft of the middle C-channel spar, and is defined by the upper skin panel forward of the middle C-channel spar. The leading edge region of the primary structural element defines the leading edge of the structural composite airfoil.

Composite assemblies are described that include composite spars that are co-cured with one or more sacrificial members on their flanges, forming an integrated sacrificial surface for the composite spars. In one embodiment, the composite assembly includes a composite spar having a web and flanges that project from sides of the web. The composite assembly further includes a sacrificial member of composite materials co-cured with the composite spar on an outer surface of at least one of the flanges. In addition, the sacrificial member has an outer surface that has been machined into conformance with an inner surface of at least one skin panel for an aircraft structure to form a contact surface with the at least one skin panel.

Composite assemblies are described that include composite spars that are co-cured with one or more sacrificial members on their flanges, forming an integrated sacrificial surface for the composite spars. In one embodiment, the composite assembly includes a composite spar having a web and flanges that project from sides of the web. The composite assembly further includes a sacrificial member of composite materials co-cured with the composite spar on an outer surface of at least one of the flanges. In addition, the sacrificial member has an outer surface that has been machined into conformance with an inner surface of at least one skin panel for an aircraft structure to form a contact surface with the at least one skin panel.

Structural composite airfoils include a front C-channel spar having an upper flange coupled to an upper skin panel, and a lower flange coupled to a lower skin panel. The lower skin panel includes a joggle adjacent a lower leading edge end, with the joggle being positioned forward of the front C-channel spar. An upper skin surface of the lower leading edge end faces an internal volume defined between the upper skin panel and the lower skin panel. A lower skin surface of the lower leading edge end faces an internal surface of the upper leading edge end of the upper skin panel, and the lower leading edge end of the lower skin panel is coupled to the upper leading edge end of the upper skin panel forward of the front C-channel spar. The leading edge has a bullnose shape defined by the upper skin panel.