The unmanned aerial vehicle (UAV) includes detachable motor arms. In this way, the UAV may be conveniently stored and transported, rapidly assembled in the field, and repaired in the event of a crash. The motor arms are also configured to separate from the fuselage in the event of a crash. An example unmanned aerial vehicle comprises: a fuselage and two motor arms. Each motor arm is detachably secured to the fuselage by two mechanical connectors and comprises a tube having a rotary wing propulsion system on each end and an electrical connector, positioned between the two rotary wing propulsion systems, configured to conductively interface with an electrical connector in an underside of the fuselage. The two mechanical connectors detachably securing each motor arm to the fuselage are configured to facilitate the separation of that motor arm from the fuselage during a crash.

Various implementations described herein are directed to an aircraft. The aircraft includes an airframe; and a windshield system having a transparent canopy coupled to the airframe and a transparent inner panel having a portion thereof disposed within the transparent canopy and coupled to the airframe and to at least a processor.

Airframe localized keel structures are disclosed. An example aircraft includes an airframe, a first engine mounted on a first side of the airframe, a second engine mounted on a second side of the airframe, and an airframe keel positioned on at least one of a lower portion of the airframe or an upper portion of the airframe between the first engine and the second engine, the airframe keel to prevent an object from exiting the first engine and impacting the second engine.

A landing gear of an aircraft of the present invention, includes a core section with a honeycomb structure, including a plurality of cell walls and a plurality of cell holes defined by the plurality of cell walls; a cover section which covers the core section; and a hole provided in the core section to absorb an impact, the hole having a diameter larger than that of the plurality of cell holes and extending in an extending direction of the plurality of cell holes.

A collapsible flap deployment system for an aircraft wing that includes a support beam pivotably connected to a carrier beam, which is connected to a wing flap. A rear spar fitting is connected to a wing rear spar by a first plurality of fasteners and a second plurality of fasteners connects the support beam to the rear spar fitting. A fuse pin connects a link to the rear spar fitting. The fuse pin is configured to shear upon the application of a first predetermined force. An impact load to the bottom of the support beam puts the link into compression applying a force to the fuse pin until the fuse pin shears releasing the link from the rear spar fitting. The impact load rotates the support beam and carrier beam causing the second plurality of fasteners to fail releasing the support beam and carrier beam from the wing rear spar.

An impact resistant fuselage of an aircraft, the fuselage extending along a central longitudinal direction, wherein transversal sections of the fuselage are comprised in a vertical plane perpendicular to the central longitudinal direction. The impact resistant fuselage comprises at least a ballistic material membrane extended along the longitudinal direction for absorbing high energy impacts. The membrane according to a transversal section, comprising at least one section between two tensional elements, wherein the material membrane is located inside the fuselage of the aircraft, the at least one section of the membrane is mechanically linked to the inside of the fuselage by the tensional elements, and the two tensional elements stress the membrane.

A metallic structure defines ribs and a skin supported by the ribs. The ribs may be defined by metal strips and the skin may be attached to the ribs. Alternatively, the skin may be defined by a plurality of tiles and the ribs may be defined by flanges of each of the plurality of tiles that cooperate to define the ribs. Tiles may be attached to separate rib lattice. Structurally weak locations at nodes where ribs intersect may be reinforced. The components may be brazed together and the stiffness of adjacent locations in the structure adjusted in the brazing operation to reduce the difference in stiffness and to reduce resulting stress risers. The metallic structure may be armored using metal foam to absorb the energy of a projectile.

Various implementations described herein are directed to an aircraft. The aircraft includes an airframe; and a windshield system having a transparent canopy coupled to the airframe and a transparent inner panel having a portion thereof disposed within the transparent canopy and coupled to the airframe.

A system for providing an aircraft with protection against hard landings, the system comprising friction energy absorber systems arranged at the points the aircraft will impact against the ground in the event of a crash. The friction energy absorber system has two energy absorber devices arranged between a base and a cover along two non-parallel deformation axes, a support secured to the base and suitable for moving in translation relative to the cover, and a friction device. The friction device is arranged between the cover and the support and it generates a friction force along a translation axis parallel to a deformation axis such that the energy absorber device, after being plastically deformed during a crash, remains in contact with the support and the cover.

Typical composite panels are brittle and unable to support transverse pressure loads that might be imposed on the panels. For example, the use of typical panels around fuel tanks of a vehicle are unable to support transverse pressure loads that might be imposed on the fuel tanks during a crash of the vehicle or a ballistic impact to the fuel tanks. In the embodiments described herein, panels include face sheets that are bonded to a foam core. The foam core includes a corrugated core sheet that is formed from a highly ductile material, such as Polyethylene or Aluminum. When a transverse pressure load is imposed on the panel, core crush of the foam occurs as the core sheet elongates from its original corrugated shape to a curve shape during deformation. This allows the panel to dissipate the energy of the transverse pressure load applied to the panel.