A wingtip device to attach to a wing of an aircraft includes an aerofoil portion and a connection spar arranged to be received in a connection spar receiving portion of the wing of an aircraft. The spar includes a first through hole in a first portion of the connection spar and a second through hole in a second rearward portion of the connection spar, the second portion being rearward of the first portion with respect to a direction of flight. The first and second through holes are arranged to receive a retaining pin that is to be inserted through both the first and second through holes.

A wing assembly include at least one fuel tank having a tank outboard end. In addition, the wing assembly includes a stout wing rib located proximate the tank outboard end and extending between a front spar and a rear spar. The wing assembly also includes at least one outboard wing rib located outboard of the stout wing rib and defining an outboard wing bay. The wing assembly also includes an upper skin panel and a lower skin panel each coupled to the front spar, the rear spar, the stout wing rib, and the outboard wing rib. A plurality of bead stiffeners are coupled to the upper skin panel and/or the lower skin panel and are spaced apart from each other within the outboard wing bay.

Methods and apparatus are disclosed for deployable wing portions of an aircraft. An example method of deploying an aircraft includes separating the aircraft from a launch vehicle, the aircraft having a wing pivotably coupled to a fuselage, rotating, about an axis of rotation, the wing relative to the fuselage from a first rotational orientation to a second rotational orientation different from the first rotational orientation, wherein, in the first rotational orientation, the wing extends along a direction that substantially aligns with a longitudinal axis of the fuselage, and extending the wing in a lateral direction away from the fuselage in the second rotational orientation.

A composite component (16) for a vehicle (10) includes a laminate (18) made from a composite material, a first pad-up area (22) applied to the laminate (18), where the first pad-up area (22) includes a plurality of first tows laid next to one another in a side-by-side arrangement and where the first pad-up area (22) defines a first fiber orientation, and a second pad-up area (24), where the second pad-up area (24) includes a plurality of second tows laid next to one another in a side-by-side arrangement and where the second pad-up area (24) defines a second fiber orientation that differs by a predetermined angle from the first fiber orientation. The first pad-up area (22) and the second pad-up area (22) intersect at an intersecting area and together form a first pad-up ply on the laminate (18).

Sound absorbers and airframe components comprising such sound absorbers are disclosed. In one embodiment, an airframe component comprises an aerodynamic surface (48) and a sound absorber (38). The sound absorber (38) comprises a perforated panel (40) having a front side exposed to an ambient environment outside of the airframe component and an opposite back side. The panel (40) comprises perforations extending through a thickness of the panel for permitting passage of sound waves therethrough. The sound absorber (38) also comprises a boundary surface spaced apart from the perforated panel. The boundary surface and the back side of the perforated panel (40) at least partially define a cavity in the airframe component for attenuating some of the sound waves entering the cavity via the perforations in the perforated panel (40).

An apparatus is provided for analysis of a leading edge rib of a fixed leading edge section of an aircraft wing. The apparatus may identify geometric or inertial properties of a plurality of stiffeners of the rib in which respective stiffeners are represented by a collection of geometry within a solid model of the rib, and perform an analysis to predict a failure rate of the leading edge rib under an external load. From the failure rate, the apparatus may determine a structural integrity of the leading edge rib under the external load. Identifying the properties may include, extracting a section cut of the geometry that corresponds to and has one or more properties of the respective stiffener, and identifying the properties of the cross-section and thereby the respective stiffener based on a correlation of the cross-section to a generic profile of a plurality of different cross-sections.

An apparatus is provided for analysis of a leading edge rib of a fixed leading edge section of an aircraft wing. The apparatus may identify geometric or inertial properties of a plurality of stiffeners of the rib, and based thereon perform an analysis to predict a failure rate of the leading edge rib under an external load. From the failure rate, the apparatus may determine a structural integrity of the leading edge rib under the external load. Performing the analysis may include importing a plurality of section cuts into a finite element model of the rib and thereby identifying nodes proximate the section cuts. Under an external load, internal load distributions may be extracted from elements proximate the nodes and elements, and the failure rate of the leading edge rib under the external load may be predicted based on the internal load distributions of the elements.

A beam flange clamp includes a clamp body having a first vise jaw portion with a first gripping surface and a second vise jaw portion with a second gripping surface. The first gripping surface and the second gripping surface are oriented parallel to each other, are facing the same direction and are offset in height by a predetermined clamping distance. The at least one bolting hole is formed through the first vise jaw portion perpendicularly to the first gripping surface.

A beam flange clamp includes a clamp body having a first vise jaw portion with a first gripping surface and a second vise jaw portion with a second gripping surface. The first gripping surface and the second gripping surface are oriented parallel to each other, are facing the same direction and are offset in height by a predetermined clamping distance. The at least one bolting hole is formed through the first vise jaw portion perpendicularly to the first gripping surface.

A beam flange clamp includes a clamp body having a first vise jaw portion with a first gripping surface and a second vise jaw portion with a second gripping surface. The first gripping surface and the second gripping surface are oriented parallel to each other and are offset in height by a predetermined clamping distance. The clamp body further including a bolting portion protruding from the second vise jaw portion opposite of the first vise jaw portion, and at least one bolting hole is formed through the bolting portion.