259results about "Fuselage bulkheads" patented technology

A hydrogen aircraft includes: an airframe including a fuselage and a wing; at least one propulsion system fixed to the airframe; a pressurized chamber disposed inside the fuselage; a pressure bulkhead that is disposed at a rear part of the pressurized chamber and has strength to withstand pressurization of the pressurized chamber; a hydrogen tank that is disposed in an accommodation section and stores liquid hydrogen, the accommodation section being installed behind the pressure bulkhead, outside the pressurized chamber, and inside the fuselage; and a supply line that supplies the liquid hydrogen stored in the hydrogen tank to the propulsion system.

An aircraft structure part having a stacked arrangement of layers with the layers being one of an at least first doubler and a basic layer of the outer skin of an aircraft fuselage, with the first doubler and the basic layer being bonded by a first intermediate bond film layer to form a step-shaped configuration of the aircraft structure part, wherein the first doubler includes a first fiber prepreg layer, in particular glass fiber prepreg layer, arranged at the perimeter of the first doubler and wherein the first fiber prepreg layer is bonded to the basic layer by at least one first bond-film segment deposited adjacent to the first fiber prepreg layer in a first wedge-shaped element at least partially incorporating the first fiber prepreg layer. A method is disclosed for providing an aircraft structure part and an aircraft fuselage provided with the aircraft structure part.

An aircraft structure including an outer mold line skin, one or more frames located on an inner surface of the outer mold line skin having an outer flange attached to an inner surface of the outer mold line skin, a first core filler in contact with the outer flange, an inner flange in contact with the first core filler and a web skin, wherein the web skin and the outer flange encapsulate the inner flange and at least a portion of the first core filler, and an arched inner mold line skin located between each frame of the one or more frames, wherein the arched inner mold line skin is in contact with each web skin of the one or more web skins.

A variable fuselage fuel tank structure of a high-speed aircraft comprises a fuselage fixing structure, an external flexible heat insulation skin and an internal skin bearing framework. The machine body fixing structure comprises a lower wall plate and an upper wall plate; the edges of the lower wall plate and the upper wall plate are connected, and the upper wall plate is provided with a fuel circulating hole; the outer flexible heat insulation skin is connected with the edge of the upper wall plate; the external flexible heat-insulating skin comprises a flexible fabric heat-proof outer layer and a microporous aerogel heat-insulating inner layer; the inner skin bearing framework is connected to the inner side of the outer flexible heat insulation skin and comprises reinforcing longitudinal beams, elastic elements and high-temperature-resistant rubber. The longitudinal beams are connected to the inner side of the outer flexible heat insulation skin in the longitudinal direction. A plurality of groups of elastic elements are respectively connected between two adjacent longitudinal beams, are distributed along the longitudinal direction, and are connected to the inner side of the external flexible heat insulation skin; and the high-temperature-resistant rubber is filled between the longitudinal beam and the elastic element, is filled in the elastic element and is adhered to the inner side of the external flexible heat-insulating skin.

The invention discloses a laser processing microstructure electric heating anti-icing and de-icing skin and a using method thereof.The laser processing microstructure electric heating anti-icing and de-icing skin comprises a skin body arranged on the surface of an airplane, the skin body comprises an electric heating microstructure anti-icing and de-icing skin body, a microstructure anti-icing and de-icing skin body and a common skin body, and the input end of the electric heating microstructure anti-icing and de-icing skin body is connected with an airborne power source through a wire; the electric heating micro-structure anti-icing and deicing skin is arranged on the windward side, with the maximum airflow collection coefficient, of the outer surface of the aircraft, the micro-structure anti-icing and deicing skin is arranged on the upstream and downstream of the windward side of the outer surface of the aircraft, and the common skin is arranged on other parts, except the electric heating micro-structure anti-icing and deicing skin and the micro-structure anti-icing and deicing skin, of the outer surface of the aircraft. According to the laser processing microstructure electric heating anti-icing and deicing skin and the using method thereof, through matched arrangement of the electric heating microstructure anti-icing and deicing skin and the microstructure anti-icing and deicing skin, the purposes of being good in anti-icing and deicing effect, low in energy consumption, light in weight and good in maintenance performance can be achieved.

Methods and apparatus for determinant assembly support are disclosed. A disclosed example apparatus to support a panel component for determinant assembly thereof includes a support frame that is releasably couplable to a jig, a first clamp to support the panel component at a first position of the panel component, a first body of the first clamp or the support frame including a nominal clearance opening to receive a first fastener to operatively couple the first clamp to the support frame, and a second clamp to support the panel component at a second position of the panel component different from the first position, a second body of the second clamp or the support frame including a slotted opening to receive a second fastener to operatively couple the second clamp to the support frame, the second fastener movable about the slotted opening to enable the second clamp to reduce forces imparted to the panel component while held by the support frame.

The invention relates to the technical field of layered products, and particularly discloses a heat-insulating and wave-absorbing skin material as well as a preparation method and application thereof. The heat-insulating and wave-absorbing skin material is of an integrated layered structure based on an aluminum oxide-mullite system, comprising a functional gradient layer with double gradients of fiber density and aperture, an impedance matching layer with a gradient dielectric constant, a resonance wave absorbing layer with a periodic dot matrix formed by silicon carbide and aluminum oxide core-shell microspheres, and a heat insulation layer with a nano-micron double-scale pore structure. According to the preparation method, a green body is constructed mainly through tape casting and silk-screen printing, and then co-sintering integrated forming is carried out. The material has the excellent characteristics of high temperature resistance, toughness, efficient heat insulation, broadband wave absorption and the like, and is suitable for extreme environments such as hypersonic aircraft skin and the like.

The invention discloses a stringer assembly, a stiffened wall plate, a stiffened wall plate assembling method and an aircraft, and belongs to the technical field of aircraft manufacturing. The stringer assembly comprises a cap-shaped stringer and a forming die, the cap-shaped stringer comprises a stringer main body and a lap joint part, the stringer main body extends in the first direction and is provided with a cavity, the forming die comprises a containing part extending in the first direction, the containing part is provided with a containing groove, and the lap joint part is provided with a lap joint part. The containing groove extends in the first direction and is opened in one side of the second direction, the containing groove and the opening of the cavity face the same direction, the stringer body penetrates through the opening of the containing groove and is installed in the containing groove, the end, in the first direction, of the containing groove is closed, and the lap joint part is in lap joint with the forming mold. The cap-shaped stringer is blocked by the closed end of the containing groove of the containing part, it is guaranteed that the cap-shaped stringer cannot penetrate through the containing part and pass through the net size area in the skin, the influence on the surface quality of the skin is reduced, and the forming die cost is reduced.

A partition includes, but is not limited to, a first body comprising a planar, foldable material. The partition further includes a plurality of anti-fold flat substrates. The partition further includes an adhesive. The substrates are adhered to the first body by the adhesive and are arranged in a repeating pattern defining a plurality of columns and rows. Each column and each row is spaced from its adjacent counterparts, defining longitudinal gaps between the columns and lateral gaps between the rows. The first body is folded laterally along the longitudinal gaps in an accordion-like manner and simultaneously folded longitudinally along the lateral gaps in an accordion-like manner. The partition contracts in both the lateral and longitudinal directions when the first body is simultaneously folded laterally and longitudinally. The partition extends in both the lateral and longitudinal directions when the first body is simultaneously unfolded laterally and longitudinally.

The invention relates to a method for automated docking of a stationary part (1) with a movable part (2) capable of being moved towards the stationary part by a robot (3), the stationary part (1) and the movable part (2) each comprising an end (4, 5), the two ends forming a docking interface (6). The method comprises the steps of positioning of multiple profilometers around the docking interface so that the docking interface is located in the field of view of the profilometers, determination of a target profile of the end of the movable part, measurement of a profile of the end of the movable part by the profilometers, comparison of the target profile and the measured profile, generating a speed setpoint in the measurement space, and movement of the movable part towards the stationary part (1) by the robot (3) on the basis of this speed setpoint.

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.

This invention provides a hyperelastic flexible structure, forming tooling, and a method for fabricating the hyperelastic flexible structure. The hyperelastic flexible structure includes a first helical wire, a second helical wire, and welded mass microspheres. The length direction of the first helical wire is orthogonal to the length direction of the second helical wire. The welded mass microspheres are positioned at the intersection of the first and second helical wires to connect them. This invention connects the orthogonally arranged first and second helical wires with welded mass microspheres to form a hyperelastic flexible structure. Compared to existing hyperelastic structural materials, it possesses excellent multifunctional mechanical properties such as large deformation recovery elastic deformation and zero Poisson's ratio. It can be used as a skeletal reinforcement for novel multiphase composite hyperelastic structural materials or as a standalone component.

The invention provides an aircraft composite heat insulation plate based on welding forming, a flap lower wall plate and an assembling method, and the aircraft composite heat insulation plate based on welding forming comprises an upper frame and a lower frame which are sequentially stacked to form a frame assembly; the outer edge of the outer skin and the frame assembly are aligned and stacked, the lower surface of the outer skin abuts against the lower frame, and the upper frame, the lower frame and the outer skin are formed at a time in a welding mode; the aerogel layer is arranged on the lower surface of the outer skin and located in the frame assembly. The invention has the beneficial effects that the nano porous structure of the aerogel limits the motion of gas molecules and reduces convective heat transfer, and the complex nano skeleton increases the heat conduction path, improves the heat radiation absorption, reduces the solid heat conduction and radiative heat transfer, and can tolerate the high temperature of 650 DEG C in the jet flow area of an engine.

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 invention is related to a structural panel for use in aeronautical applications. The structural panel includes a composite skin (200) that comprises composite material and a stiffening framework that is rigidly attached to the composite skin (200). The composite skin (200) is configured as a load bearing monolithic structure (370) with a structure thickness (310) of less than 1 mm. The composite skin (200) comprises a composite laminate (480) with a plurality (400) of carbon composite plies (405). Each one of the plurality (400) of carbon composite plies (405) comprises a ply thickness (410) in a range from 0.02 mm to 0.05 mm.

The application discloses a flexible metal composite skin structure capable of regulating local out-of-plane deformation and belongs to the technical field of metal skin structures. The application introduces a paper-cut structure on the surface of a metal skin, regulates mechanical response by changing characteristic parameters such as a paper-cut pattern and a metal skin thickness, so that the structure has high ductility and small plastic deformation characteristics, and the structure can recover to original deformation after the expansion of a protrusion of a driving device, thereby ensuring the continuity of a wing aerodynamic shape. The application meets the basic requirements of application to a flexible skin of a deformation wing, provides a feasible selection for the fusion of a gasbag anti-icing mechanism and a deformation wing, and is expected to be widely applied in the fields of aircraft deicing, spacecraft deicing and deformation wings.

A method for manufacturing a fuselage structural component (200) for a triangular section of an aircraft is disclosed, comprising the following steps: providing a one-piece sheet (211, 212, 213) of frame material, the sheet having an opening (250), and deforming at least one inner edge of the sheet (211, 212, 213) forming the opening (250) into an inner flange (215, 216, 217), and deforming at least one outer edge of the sheet (211, 212, 213) into an outer flange (214). The one-piece and deformed sheet (211, 212, 213) forms, adjacent to the opening (250), a frame section (210), a bottom support (212), and a vertical support (213).

The application provides a carbon-glass hybrid composite material skin for an aircraft and a preparation process thereof, and relates to the technical field of aircraft accessories.A carbon-glass hybrid composite material skin for an aircraft, the material of the skin is a mixture of carbon fibers, glass fibers and resin, and the skin is an integrally formed structure; the mass percentage of the mixture of carbon fibers and glass fibers ranges from 60% to 70%, and the mass percentage of the resin ranges from 40% to 30%; the skin is provided with a non-thickened area and a thickened area.The application solves the technical problems of the prior art, such as the heavy weight of the aircraft skin made of metal material, the complex structure design, the poor impact resistance and stability of the product during use.The material of the skin is a mixture of carbon fibers, glass fibers and resin, and the skin is integrally formed, so that the weight is light and the overall structural strength is improved.

The present disclosure relates to a thermoforming composite, comprising a thermoforming base layer having a first principal surface and a second principal surface opposing the first principle surface; a thermoforming stringer having a base part and an elongated protruding part, wherein the base part is attached to the first principal surface of the thermoforming base layer to enclose an elongated hollow part between the first principal surface of the thermoforming base layer and the protruding part; and an elongated flexible hollow tube core inserted into the elongated hollow part. During consolidation or curing of the thermoforming base layer and the thermoforming stringer, the elongated flexible hollow tube core is pressurizable by a first pressure so that the elongated flexible hollow tube core expands to essentially the same shape as the thermoforming stringer, a second pressure is a uniform pressure exertable to the second principal surface, and the first pressure is essentially equal to a second pressure.

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 invention discloses a flexible metal composite material skin structure capable of regulating and controlling local out-of-plane deformation, and belongs to the technical field of metal skin structures. The paper-cut structure is introduced to the surface of the metal skin, the mechanical response of the paper-cut structure is regulated and controlled by changing characteristic parameters such as a paper-cut pattern and the thickness of the metal skin, so that the structure has the characteristics of high ductility and small plastic deformation, and the structure can recover to original deformation along with retraction of the driving device after expanding and protruding along with the driving device; and the aerodynamic configuration continuity of the wing is ensured. Basic requirements of flexible skin applied to the deformable wing are met, a feasible choice is provided for fusion of an air bag anti-icing and deicing mechanism and the deformable wing, and the method is expected to be widely applied to the fields of aviation, spacecraft deicing, deformable wings and the like.

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.

An airframe design may include a bonded frame or assembly, and one or more components (645) that may be removably attached to the bonded frame. The bonded frame may include struts (102), central bulkheads (106), a tail section, a plurality of wing sections, and motor mounts (650) that are adhered together using adhesive. The one or more attachable components (645) may include a forward fuselage (640), motors (756), propellers (860), motor pod fairings (862), stabilizer fins, and landing gear (972) that are attached using fasteners. The bonded frame may reduce the number of parts of the airframe design and may also reduce complexity, cost, and weight, while also increasing stiffness and strength. Further, the various attachable components (645) may facilitate fabrication, assembly, and maintenance of an aerial vehicle having the airframe design.

An aircraft (10, 10") has an aft main cargo cabin (50) with a low deck floor (52), and a rear access (60) configured to provide access to the aft main cargo cabin from behind the fuselage (20). The deck floor of the aft main cargo cabin is lower than a first deck floor (32) of a forward main cabin (30) of a forward section of the fuselage. The aft main cargo cabin has a second height (H2) that is greater than the first height (H1) of the forward main cabin, allowing the aft main cargo cabin to accommodate oversize cargo with larger dimension(s). The aircraft may be built with such an aft main cargo cabin and rear access, or may be a modified aircraft that includes such an aft main cargo cabin and rear access. Related methods of reconfiguring a base aircraft into a modified aircraft having such an aft main cargo cabin and rear access are also disclosed.

A method to manufacture a structural component in composite material reinforced with stiffening stringers and comprising a skin and plurality of stringers having a hollow section fixed to said skin. The method comprising: a) laminating a plurality of first layers of composite material onto an external surface of a cure tool, b) laminating pluralities of second layers of composite material on longitudinal hollow expandable inserts, c) positioning each reinforcing element on a face of a second skin, d) holding each reinforcing element on the face of the second skin; e) housing the tool, skin and reinforcing elements inside a vacuum bag; and f) compacting together the first layers forming the skin with the second layers forming the stringers; g) housing skin and reinforcing elements in a further vacuum bag; and h) sealing the further vacuum bag; and i) applying preset temperature and pressure to the outside of the vacuum chamber.