43results about "Fuselage bulkheads" patented technology

An air mobility apparatus and mobility apparatus having a door. The air mobility apparatus includes a fuselage having a first cabin and a second cabin; and a rear door having an inner door partitioning the first cabin and the second cabin, and an outer door partitioning the second cabin and an external space, wherein the inner door and the outer door are independently opened and closed.

A retainer 16, 18 suitable for use on an aircraft 2 is disclosed. The retainer 16, 18 comprises a first retainer part 16 and a second retainer part 18. The first retainer part 16 comprises a catch and a damper 24; in which the catch comprises one or more magnetic elements 20, 22. The damper 24 comprises a damping material, the damping material damps shock impulses; and the catch is supported on the damper.

The application relates to a cooling system of long-time high-power airborne electronic equipment, and belongs to the technical field of aircraft equipment heat dissipation, and solves the problem of insufficient heat dissipation efficiency of high-power electronic equipment of a subsonic aircraft. The cooling system comprises a heat dissipation structural member, a driving pump, a heat storage pool and an aircraft skin; the electronic equipment is in contact with the heat dissipation structural member, the heat dissipation structural member is heated and the electronic equipment is cooled at the same time through heat conduction; the heat dissipation structural member is internally provided with a cooling liquid internal flow channel for circulating cooling liquid; the driving pump is used for driving the cooling liquid to circulate between the heat dissipation structural member and the heat storage pool; the heat storage pool internally stores the cooling liquid; the cooling liquid can circulate between the heat storage pool and the aircraft skin and conduct heat to the external environment. The application can transfer heat to the external environment by driving the cooling liquid to circulate between the heat dissipation structural member, the heat storage pool and the aircraft skin, so that effective and rapid heat dissipation of the electronic equipment is realized.

This invention relates to a semi-mounted composite skin frame hull structure. A first, second, third, fourth, and fifth lower arc frame are all located below the left and right main beams and are fixedly connected to them. These lower arc frames are arranged sequentially from the front docking frame to the rear docking frame. A first, second, third, fourth, and fifth upper arc frame are all located above the left and right main beams and are fixedly connected to them. These upper arc frames are arranged sequentially from the front docking frame to the rear docking frame. This invention provides a clear force transmission path, ensuring the hull structure meets strength and rigidity requirements under various working conditions and loads, enabling rapid installation of the warhead, while simultaneously meeting design weight requirements.

A fuselage portion of or for an aircraft. The fuselage portion comprises an aircraft component which is installed in a shell of the fuselage portion by a plurality of connector elements. Therein, a respective first end of each connector element is held in an associated bore b running through the shell, and a respective second end of each connector element is joined to the aircraft component by a joint providing the respective connector element with a margin of swiveling relative to the aircraft component. Also an aircraft comprising such a fuselage portion, and a method for assembling such a fuselage portion.

The application relates to a subsonic aircraft heat dissipation device and belongs to the technical field of aircraft heat dissipation devices, solving the problems of high heat dissipation requirement and limited heat dissipation space of the subsonic aircraft device. The subsonic aircraft heat dissipation device comprises a heat dissipation fin, an adapter and a skin. The heat dissipation fin is in direct contact with the heat generating electronic device, and the heat generated by the heat generating electronic device is quickly conducted out by using the high heat conduction performance of the heat dissipation fin. The adapter is a connecting device of the heat dissipation fin and the skin, and can conduct the heat of the heat dissipation fin to the skin. The application can quickly transfer the heat to the air outside the skin, thereby forming a heat transfer channel, realizing the quick conduction of the heat, and can be used for the heat dissipation of high heat flow density electronic devices.

A composite fuselage frame for an aircraft includes a web and a shear tie foot. The web includes a width and a first thickness. The web forms an arc between first and second ends in relation to the fuselage. The width extends from the outer side in relation to a skin on the fuselage to an inner side. The shear tie foot is disposed on the outer side of the web between first and second mouse holes and projects from the first thickness of the web. The shear tie foot has a second thickness that extends along the skin between first and second stringers. The second thickness of the shear tie foot is defined by the first thickness and second polymer matrix composite layers formed by first material polymer matrix composite plies that provide a frame filler integrated with second material polymer matrix composite plies of the first thickness.

A composite structure comprises stacked sets of laminated fiber reinforced resin plies and metal sheets. Edges of the resin plies and metal sheets are interleaved to form a composite-to-metal joint connecting the resin plies with the metal sheets.

A composite fuselage frame for an aircraft includes a web and a shear tie foot. The web includes a width and a first thickness. The web forms an arc between first and second ends in relation to the fuselage. The width extends from the outer side in relation to a skin on the fuselage to an inner side. The shear tie foot is disposed on the outer side of the web between first and second mouse holes and projects from the first thickness of the web. The shear tie foot has a second thickness that extends along the skin between first and second stringers. The second thickness of the shear tie foot is defined by the first thickness and second polymer matrix composite layers formed by first material polymer matrix composite plies that provide a frame filler integrated with second material polymer matrix composite plies of the first thickness.

To provide systems and methods for assembling a section of a fuselage of an aircraft.SOLUTION: The method includes transferring a half barrel section 24 of a fuselage of an aircraft along a track 110-1, 110-2 in a process direction, utilizing an indexing feature associated with the half barrel section 24 to determine a desired contour 361 for a portion of the half barrel section 24, and enforcing the desired contour 361 on the half barrel section 24 using components that enforce an inner mold line 128 and components that enforce an outer mold line 126 when the contour of the half barrel section 24 is out of tolerance from the desired contour 361.SELECTED DRAWING: Figure 3

The present invention relates to the protection of an aircraft fuselage. To provide a protective measure that is as lightweight as possible and easily integrated into the fuselage, a protective insert (10) for the tail section of an aircraft is provided. The protective insert comprises a protective panel (12), an absorption body (14) supporting the protective panel, and a force transmission device (16) holding the absorption body. The protective panel is designed for a pressure-loaded sliding contact with a runway in the event of a tail impact. The protective panel transmits a pressure force acting during the tail impact to the absorption body. The absorption body is designed to absorb the pressure force by damping it, whereby a portion of the energy due to the acting pressure force can be absorbed and converted into an elastic and / or plastic deformation of at least a part of the absorption body.The power transmission device is designed to hold the protective insert on the underside of a tail section of an aircraft and has a plurality of attachment points (18).

The invention relates to a pressurized bulkhead for an aircraft, comprising a membrane (2) having a planar main surface and at least one hinge (3) for connecting the membrane (2) to the fuselage of the aircraft.

The present application relates to composite fuselage frames and methods of manufacturing frame fillers and composite fuselage frames. A composite fuselage frame for an aircraft includes a web and a shear tie. The web includes a width and a first thickness. The web forms an arc between a first end and a second end relative to a fuselage. The width extends from an outboard side to an inboard side relative to a skin on the fuselage. A shear tie is disposed on the outboard side of the web between a first mousehole and a second mousehole and the shear tie protrudes from the first thickness of the web. The shear tie has a second thickness that extends along the skin between a first stringer and a second stringer. The second thickness of the shear tie is defined by the first thickness and a second polymer matrix composite layer formed from a first material polymer matrix composite ply that provides a frame filler integrated with a second material polymer matrix composite ply of the first thickness.

An aircraft skin alignment system includes at least one support structure to support one or more aircraft skins and one or more assembly tools; the one or more assembly tools having a holding mechanism to releasably engage with a first aircraft skin; a pulling mechanism to apply a pulling force to the first aircraft skin; and an edge locator; the pulling mechanism is to pull the first aircraft skin such that a first edge of the first aircraft skin abuts the edge locator, the edge locator preventing the first edge from being pulled past the edge locator; and one or more tool holes can be drilled into the first aircraft skin such that the one or more tool holes are aligned based on an edge index due to a position of the edge locator.

This paper proposes a noise reduction structure for the cabin interior panel based on damping vibration reduction. The outer side of the cabin interior panel is connected to the outer layer mask. Ribs are provided between the outer layer mask and the cabin interior panel. The ribs include vertical ribs and horizontal reinforcing ribs. The vertical ribs are horizontally equidistantly arranged on the inner surface of the cabin interior panel. Damping grooves are provided on the surface of the vertical ribs. The damping grooves are horizontally equidistantly spaced from several vertical ribs on the surface of the vertical ribs. Guide edges are provided on the bottom and both sides of the damping grooves. By improving the original form of directly connecting the cabin interior panel to the ribs on the outer layer mask, damping sliders are set on the ribs. The damping sliders can effectively reduce the vibration radiation of the cabin interior panel, thereby reducing vibration noise inside the cabin. The sealed space between the outer layer mask and the cabin interior panel can also effectively isolate noise.

A method of assembling a riblet panel (200, 300a-d, 400a-d, 500, 600a-c, 700a-d, 800) for attaching riblets (202, 302a-d, 402a-d, 502, 602a-d, 702a-d) to a surface over which a fluid flows, the method comprising: preparing a plurality of riblet films, each riblet film comprising one or more riblets (202, 302a-d, 402a-d, 502, 602a-d, 702a-d); and attaching the plurality of riblet films to a common laminate film in a predetermined arrangement, wherein the riblet panel (200, 300a-d, 400a-d, 500, 600a-c, 700a-d, 800) is configured to adhere to the surface; and wherein the predetermined arrangement is configured such that, when the riblet panel (200, 300a-d, 400a-d, 500, 600a-c, 700a-d, 800) is adhered to the surface, the one or more riblets (202, 302a-d, 402a-d, 502, 602a-d, 702a-d) of the plurality of riblet films are configured to reduce the skin friction of the surface when the fluid flows over them.

An aircraft fuselage includes fore and aft sections, a barrier wall separating the sections and an opening therebetween, a rail attached to the barrier wall, a door moveable across the barrier wall between an open position offset from the opening and a closed position covering the opening, a first roller and a second roller rotatably fixed to the door and moveable on the rail along first and second paths to move the door, the rail including a first ramp engaging the first roller as it moves along the first path and to direct the first roller towards the barrier wall, and a second ramp engaging the second roller as it moves along the second path and to direct the second roller towards the barrier wall, the first roller configured to engage the first ramp as the second roller engages the second ramp to move the door towards the barrier wall.

The invention relates to an assembly (2) for reinforcing a fuselage (10) of an aircraft (1), said aircraft (1) comprising a fuselage (10) having an opening (12) and a removable panel (14) intended to be received in said opening (12), said assembly (2) comprising: - a first support (22) intended to be fixed to an outer face of the fuselage (10) as well as around and outside an opening (12) of the fuselage; - a sub-assembly (24) intended to be fixed to an outer face of a removable panel (14) of the fuselage, where said sub-assembly (24) is housed inside said first support (22); - first means (32) for fixing said first support to said fuselage; and - second means (34) for fixing said sub-assembly to said removable panel. This makes it possible to reinforce the fuselage of an aircraft when removing the removable panel, for example when maintenance of the hydrogen tanks is necessary.

This invention discloses a skin and its variable stiffness deformation control method, comprising: a shape memory polymer matrix and several liquid metal tubes; the liquid metal tubes are installed in the shape memory polymer matrix; the liquid metal tubes are heated to melt them, and the heat is conducted to the shape memory polymer matrix to achieve variable stiffness deformation. The skin structure of this invention can be applied to various morphing aircraft and deformable wings, improving upon the shortcomings of weak aerodynamic load-bearing capacity of rubber-based deformable skins and pure shape memory polymer skins; it also has high structural stiffness, meeting the aerodynamic load-bearing capacity requirements of wings; furthermore, it overcomes the shortcomings of traditional shape memory polymer skins, such as decreased deformation capacity, uneven heating, and low heating efficiency caused by embedded resistance wires or external heating films. By electrically heating the liquid metal inside the skin, the skin can be heated quickly and efficiently, changing its stiffness.

A skin assembly of an aircraft cylindrical section (100). The skin assembly comprises a plurality of wallboards which are sequentially connected in a circumferential direction and have flush inner surfaces, wherein the inner surfaces of all the wallboards form a smooth cylindrical surface; the outer surface of each wallboard comprises a plurality of equal-thickness areas (101) and transition areas (102) which are alternately arranged in a heading direction; the thickness of any equal-thickness area (101) is not less than a corresponding minimum load thickness; the surface of any transition area (102) is an annular surface arranged in the circumferential direction; the boundaries of two sides of any transition area (102) are respectively in contact with adjacent equal-thickness areas (101), with the thicknesses of the boundaries of the two sides of the transition area (102) being respectively equal to the thicknesses of the equal-thickness areas (101) in contact therewith; and any transition area (102) has a preset slope in the heading direction. A frame of the aircraft cylindrical section (100), the frame comprising the skin assembly of the aircraft cylindrical section (100). Such arrangements ensure the aerodynamic performance of an aircraft and also simplify the configurations of stringers and circumferential frames fitting closely onto the inner surface of an aircraft skin, thereby being conducive to reducing the manufacturing costs of the aircraft.

Disclosed herein is a system that is mountable to an outer panel of a vehicle, such as an aircraft. A disclosed version of the system includes an an antenna component, a carrier frame, support fittings, and an aerodynamic fairing. The antenna component is coupled to the carrier frame, which includes a plurality of clevis attachment features. Each support fitting is couplable between one of the clevis attachment features and the outer panel of the vehicle. The fairing has an upper section and a lower section. The upper section has an opening defined therein. The lower section is configured and arranged for coupling to the outer panel of the vehicle. The opening is shaped and sized to accommodate at least a portion of the carrier frame, which is coupled to the upper section of the fairing.

This invention discloses an integrated load-bearing and sweating skin structure for additive manufacturing. Its key feature is that it is composed of three structures sequentially composited from the inside out: a thin-walled dense shell structure, a spatial lattice support-fluid infusion structure, and a capillary sweating structure. The thin-walled dense shell structure layer and the spatial lattice support-fluid infusion structure layer are integrally formed through additive manufacturing, with the lattice structure in the spatial lattice support-fluid infusion structure layer extending into the capillary sweating structure layer. This structure possesses excellent stress resistance and high sweating cooling efficiency, and can replace existing passive thermal protection structures for hypersonic vehicles. It can achieve miniaturization, lightweighting, and long-range capabilities, while reducing the failure risk caused by connecting multiple structures.

This invention relates to a protective device for an aircraft fuselage. To provide protection with the lowest possible weight and good integration into the fuselage, a protective insert (10) is provided for the tail region of the aircraft. The protective insert has a protective panel (12), an absorber (14) supporting the protective panel, and a force transmission device (16) for holding the absorber. The protective panel is configured to slide into contact with the landing / takeoff runway under applied pressure during a tail impact. The protective panel transmits the pressure acting during a tail impact to the absorber. The absorber is implemented in a force-buffering manner to absorb the pressure, wherein a portion of the energy generated by the acting pressure can be absorbed and converted into elastic and / or plastic deformation of at least a portion of the absorber. The force transmission device is configured to hold the protective insert under the tail region of the aircraft and has multiple fastening areas (18).