69results about "Fuselage frames" patented technology

Apparatus and methods for ultrasonic inspection of elongated composite members in a single scan pass using pulse echo phased arrays (97, 98, 99) operating in a bubbler method. The system concept is fully automated by integrating an inspection NDI probe assembly (32, 34, 36) to a robot (100) and using the robot to move the inspection NDI probe assembly along the part (i.e., outside of an inspection tank); and by integrating tooling fixtures that move out of the way as the inspection NDI probe assembly travels along the length of the part during the inspection. In addition, the system allows for generally elongated composite members having lengthwise variation in shape, curvature and dimensions.

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.

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.

This invention discloses a tiltrotor aircraft, relating to the field of tiltrotor aircraft design technology. It includes a fuselage with an overall spindle-shaped profile and an inverted trapezoidal cross-section. The outer peripheral surface of the fuselage includes an upper surface, a lower surface, and two side surfaces. The two side surfaces are connected between the upper and lower surfaces and are located on opposite sides of the fuselage. The upper surface, lower surface, and two side surfaces are all composed of complete swept surfaces, each formed by a straight-line sweep. This achieves coordinated suppression of the aircraft's full radar cross-section and aerodynamic noise, enhancing its survivability.

A rotary-wing aircraft with a fuselage that comprises a load carrying shell structure and a load carrying stiffening member structure, wherein the load carrying shell structure delimits an aircraft inner volume, and wherein the load carrying stiffening member structure is arranged outside of the aircraft inner volume on an outer surface of the load carrying shell structure.

The application relates to a light main body skeleton structure applied to a tilt-rotor aircraft, which comprises a chassis assembly, a frame assembly and a tail frame, the chassis assembly is provided with oppositely arranged front and rear ends, and the tail frame is located at the rear end; the frame assembly comprises a first truss, a second truss, a third truss and a fourth truss, and the first truss, the second truss, the third truss and the fourth truss are sequentially and fixedly arranged on the chassis assembly from the front end to the rear end; the frame assembly further comprises truss beams, the truss beams are sequentially connected with the first truss, the second truss, the third truss and the fourth truss, so that the frame assembly forms a streamlined shape, has the effect of simplifying the body structure, and facilitates the aircraft to realize long-range high-speed flight.

The utility model relates to the technical field of aircraft, and disclose a machine arm joint mounting structure, including machine arm connector, the outer ring both sides of machine arm connector are provided with fixed mounting block no.

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.

The utility model relates to aviation aircraft technical field, concretely is an aviation aircraft with floating anti -sinking material quality shell, including aviation aircraft main part, be equipped with fuselage frame body on the aviation aircraft main part, and the fuselage frame body upper end is fixed with fuselage framework, the fuselage framework outside cover is equipped with mutually sealed clamping nose shell, backrest shell and chassis shell, and the inside of fuselage framework, backrest shell and chassis shell is clamped with lining shell respectively, and sets up sealable gas filling layer in lining shell, the bottom of fuselage frame body is equipped with chassis protection ring. The aviation aircraft with floating anti -sinking material quality shell whole frame body adopts lightweight design, and through setting up lining shell with gas filling layer in the nose shell, can provide floating dust sink function under the condition that aviation aircraft self load is not increased, guarantees the buoyancy of aviation aircraft when flying, improves the security of aviation aircraft near water area near execution task.

The present application relates to a kind of continuous plant fiber reinforced honeycomb composite gradient damping structure, preparation device and preparation method, the structure includes: upper damping layer, upper panel, intermediate core layer, lower panel and lower damping layer, the upper panel, intermediate core layer and lower panel are sequentially stacked, the upper panel and lower panel are all provided with recess on, and the upper panel and lower panel with the recess constitute gradient thickness structure, the upper damping layer and lower damping layer are respectively arranged in the recess bottom of upper panel and lower panel, the intermediate core layer is gradient height honeycomb structure, and the upper panel and lower panel use multiscale continuous plant fiber reinforced honeycomb composite material.Compared with prior art, the present application has the advantages of realizing excellent broadband damping without increasing mass.

The application relates to a preparation method of a high-density graphite-based dissipation heat-resistant composite material. In order to solve the problems of high cost of the existing dissipation heat-resistant composite material and poor ablation resistance of the graphite dissipation heat-resistant composite material, the preparation method comprises the following steps: mixing multi-element dissipation agent raw materials, uniformly mixing the multi-element dissipation agent raw materials, loading the multi-element dissipation agent raw materials into a graphite crucible, processing a high-density graphite base into a shape to be prepared, mounting the high-density graphite base on a lifting rod of a vacuum infiltration furnace, preparing a dissipation agent alloy melt, and immersing the graphite base in the dissipation agent alloy melt to obtain the high-density graphite-based dissipation heat-resistant composite material. The high-density graphite-based dissipation heat-resistant composite material is prepared by using the vacuum infiltration method, high-purity graphite is used as the base, the composition of the dissipation agent alloy is adjusted, the dissipation agent is wetted with the graphite base, and the ablation resistance of the workpiece is improved. The application can be used for preparing aircraft engine nozzle throat liners, aircraft nose cones and wing leading edges and the like components.

This utility model discloses an electric vertical takeoff and landing manned aircraft, including an aluminum alloy frame, a vibration damping assembly, a right guardrail connected to one side of the aluminum alloy frame, and a left guardrail connected to the other side via a pivot seat. The vibration damping assembly includes a first carbon fiber tube and a vibration damping seat fitted on the first carbon fiber tube. The upper and lower vibration damping seats are fastened with bolts through the penetration between the first carbon fiber tube and the aluminum alloy frame. The rubber vibration damping pad absorbs high-frequency vibration energy through its own elastic deformation, effectively blocking the direct transmission of vibration to the frame. The axial anti-roll bracket forms an elastic clamp on the sleeve through rubber clips. This clamping force can be flexibly adjusted by the hardness of the rubber clips, which greatly reduces the vibration felt in the cockpit area. Both the first and second carbon fiber tubes are made of carbon fiber composite material, which has a weak ability to reflect sound waves and can reduce the transmission efficiency of propeller aerodynamic noise.

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).

This present invention relates to the field of aircraft and discloses an aircraft and a method for connecting a canard to an arm. The connection method is used to couple a canard and an arm. The canard comprises a coupling insertion portion, which comprises a plurality of fixing portions. The arm comprises an arm insertion cavity, and an insertion opening is provided at one end of the arm insertion cavity. The coupling insertion portion is configured to be inserted into the arm insertion cavity through the insertion opening and is fixed within the arm via the fixing portions. In the present invention a detachable structure is enabled between the canard and the arm, allowing the position of the center of lift to be adjusted by exchanging canards featuring different designs, thereby adapting to aircraft cabins with different centers of gravity.

An assembly platform configured to manipulate at least one aircraft lower shell without deforming it with a view to assembly of an aircraft fuselage barrel is disclosed including at least one rectangular chassis, at least two frames having an adjustable trim with respect to the rectangular chassis, at least two setting systems intended to adjust the trim of the frame and at least two cradles fixed onto the frames and intended to receive the lower shell and to hold the lower shell in a desired form. The cradles of the assembly platform make it possible to prevent a deformation of the lower shell during the assembly of the fuselage barrel.

An electric VTOL aircraft has a fuselage having a nacelle that is flexibly connected to a power plant assembly that includes a plurality of electric rotors and arms. The power plant assembly may pitch and roll relative to and independently of the nacelle to produce efficient forward, rearward, and lateral thrust while the nacelle attitude remains unchanged. This provides a stable passenger experience as well as enhanced performance and controllability, while reducing cost and complexity. In some embodiments, the fuselage is vertically elongated, and the power plant assembly is mounted above the fuselage such that a person can walk fully upright below the rotors. Also disclosed is a VTOL docking station configured to allow an aircraft to land without the use of landing gears. The docking station also includes an electronic component configured to automatically provide charging to the aircraft when the aircraft is docked with the docking station.

The application relates to a flight power system and a flying car. The flight power system comprises a flight support and a first arm and a second arm connected to the flight support. The flight support comprises a connecting frame, a first mounting frame and a second mounting frame, the connecting frame has opposite first and second ends, the first mounting frame is connected to the first end of the connecting frame and is bent relative to the connecting frame, the second mounting frame is connected to the second end of the connecting frame and is bent relative to the connecting frame to be arranged in opposite spaced relation with the first mounting frame, the connecting frame, the first mounting frame and the second mounting frame jointly define an accommodating space, the first arm is rotatably connected to the first mounting frame, the second arm is rotatably connected to the second mounting frame, and the first arm and the second arm are accommodated in the accommodating space in a folded state. The flight power system can drive the flying car to vertically take off and land, greatly reducing the influence of the surrounding environment and take-off conditions.

The present invention relates to a fuselage section having a tapered wing rib interface. Disclosed is a fuselage section having a tapered wing rib interface. The disclosed example apparatus includes a rib associated with the fuselage section, and a wing interface surface defined by the rib, wherein the wing interface surface is tapered with respect to a longitudinal axis of the fuselage section from an aft end of the fuselage section to a forward end of the fuselage section.

An assembly comprising at least first and second components, at least one connection element comprising a rod configured to ensure uptake of the traction / compression forces between the first and second components, and at least one pair of first and second contact surfaces provided respectively with first and second bumps and rigidly secured respectively to the first and second components, the first and second bumps having complementary shapes configured to nest in one another and ensure transmission of the shear forces between the first and second components. Also, an aircraft comprising at least one such assembly.

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.

This invention relates to a decorative cover assembly, which includes a decorative cover body; a ring-shaped connector fixed to a wall surface; and a hook connector fixed to the decorative cover body. The hook connector engages with the ring-shaped connector to achieve a detachable connection between the decorative cover body and the wall surface. The hook connector includes a hook member comprising a hook seat, a hook portion, a hook pivot, and an elastic element. The hook seat is fixed to the decorative cover body. The hook portion is pivotally connected relative to the hook seat around the hook pivot between a closed position and an open position. In the closed position, the hook portion is close to the hook seat to engage with the ring-shaped connector. In the open position, the hook portion is away from the hook seat to disengage from the ring-shaped connector. The elastic element biases the hook portion toward the closed position.

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.

The present invention relates to a preform of a bulkhead in an aircraft and a manufacturing method therefor. The preform of the bulkhead is provided with edge strips located at two ends of the bulkhead and a web located in the middle of the preform and connecting the edge strips. The web has a curvature in the longitudinal direction of the web. The web is laid with 0-degree yarns extending without wrinkles at a 0-degree angle relative to the longitudinal direction of the web. By means of the present invention, the 0-degree yarns can be arranged on the web of the preform of the bulkhead in the longitudinal direction of the web without wrinkles, such that the 0-degree yarns have tension in the longitudinal direction of the web, thereby meeting stress requirements for the bulkhead in the longitudinal direction, ensuring smoothness and flatness of fibers of the preform, improving the adaptability of the preform of the bulkhead to curved surfaces, increasing the manufacturing efficiency of the preform of the bulkhead, and ensuring manufacturing quality and reliability.

An aircraft cabin component for use in an aircraft cabin of an aircraft includes a component having at least one parameterized surface. A flexible OLED display assembly is adhered to the at least one parameterized surface of the component and closely conforms to the curvature of the at least one parameterized surface. The flexible OLED display assembly includes a printed electronics layer and an array of OLEDs deposited on the printed electronics layer.