147results about "Weight reduction" patented technology

Powertrains and related methods for an aerial vehicle may include a torque control system associated with the powertrain and configured to receive torque signals indicative of engine torque supplied by a mechanical power source or generator torque generated by an electric power generation device resisting the engine torque. The torque control system may be configured to generate, based in part on torque signals, a torque control signal configured to change the engine torque or change the generator torque. When torque signals indicate a relative reduction in the engine torque supplied by the mechanical power source, torque control signals may be configured to cause a relative reduction in the generator torque resisting the engine torque. When torque signals indicate a relative increase in the engine torque supplied by the mechanical power source, torque control signals may be configured to cause a relative increase in the generator torque resisting the engine torque.

Described herein are methods and systems for forming composite stringer assemblies, or more particularly, methods and systems for shaping a composite charge as the stringer assemblies are formed. A system includes a bladder having a bladder core and a bladder skin. The bladder core is formed of a foam material. The bladder skin is formed of a resilient material and surrounds the bladder core. The bladder is positioned on a charge base as the composite stringer assembly is formed. The charge base subsequently becomes a stringer base, such as a fuselage section or a wing skin. A charge cap is then positioned on the bladder and is configured to conform to the bladder. The combination of the bladder skin and the bladder core provides support during the forming operation and subsequently as the stringer assembly cures. In some examples, the bladder core is collapsible so as to remove the bladder from a cavity of the stringer assembly.

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.

The application discloses a method for constructing a standard oxygen concentration environment in an aircraft fuel tank, and comprises the following steps: step 1, constructing an external environment of the aircraft fuel tank, wherein the aircraft fuel tank is a fuel tank simulated through experiments; step 2, constructing an internal environment of the aircraft fuel tank; step 3, constructing an oil and gas environment of the fuel tank; and step 4, constructing a standard oxygen concentration environment of the fuel tank. The technical scheme of the embodiment of the application solves the problem that calibration conditions are lacked for an aircraft fuel tank oxygen concentration testing device.

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 molding apparatus and method for wedge-shaped metal honeycomb core adhesive products. The molding apparatus includes a tooling plate, multiple corner supports, segmented core molds, gaskets, a primary curing and equalizing plate, and a secondary curing and equalizing plate. The segmented core molds are placed along the inner walls of the beams and ribs of the wedge-shaped metal honeycomb core adhesive product. The multiple corner supports are fixed on the tooling plate and block one side of the segmented core molds. The gaps between the segmented core molds and the beams, ribs, and corner supports are adjusted using gaskets. After the wedge-shaped metal honeycomb core adhesive parts, core molds, and corner supports are assembled, secondary curing and molding are performed using the primary and secondary curing and equalizing plates. This invention solves the problems of mismatch between the wedge-shaped metal adhesive core mold and the parts, and uneven quality of the adhesive layer. It improves the quality of the manufactured parts while simplifying the assembly process, eliminating the need for continuous cold pressing to check the fit between the skin and the beams and ribs, and saving manufacturing time and money.

The application discloses a reconfigurable multi-rotor unmanned aerial vehicle based on a bistable hinge, which comprises a bistable hinge, a rotor assembly, a magnet and a hand grabber, etc. The bistable hinge is composed of a fuselage module, a pin shaft and a tension spring. The unmanned aerial vehicle has two different fuselage configurations, i.e. a ring configuration (configuration A) and a "I" shape configuration (configuration B). The counter torque generated by the same direction clockwise / anticlockwise rotation of the rotor overcomes the magnet attraction and the spring tension, switches the fuselage hinge steady state, realizes the "opening" / "closing" of the fuselage of the unmanned aerial vehicle, changes the wheelbase of the unmanned aerial vehicle while converting the fuselage configuration of the unmanned aerial vehicle. Moreover, the tension of the tension spring and the attraction of the magnet together guarantee the structural stability of the unmanned aerial vehicle under the two fuselage configurations. In addition, the "I" shape fuselage configuration and the hollow ring fuselage configuration respectively make the unmanned aerial vehicle have the ability to shuttle in narrow space and grab objects.

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 application discloses a kind of for aircraft composite wallboard turnover tool's pose adjusting device and method, include multiple groups of turnover pose adjusting beam and control system, every group of turnover pose adjusting beam includes beam frame, pose adjusting mechanism, protective cover, maintenance cover plate, electrical element, beam frame main body is half-closed long strip plate welding structure, two ends are connected with wallboard turnover tool respectively, two groups of pose adjusting mechanism are embedded and fixed in the middle layer main bearing plate of beam frame, control system controls multiple groups of turnover pose adjusting beam pose adjusting and wallboard suction fixation, multiple groups of turnover pose adjusting beam synchronous turnover will wallboard be turned to required pose.

An aircraft nose fuselage skin forming die and a forming method thereof belong to the field of aircraft skin forming. The aircraft nose fuselage skin forming die comprises a bottom plate, two side plates, a rotating device, a top plate and a limiting device. The upper ends of the bottom plate are symmetrically fixedly connected with the two side plates, and the upper ends of the side plates are fixedly connected with the top plate. The rotating device is fixedly connected with the upper end of the bottom plate. The limiting device is movably connected with the rotating device. The aircraft nose fuselage skin forming die can not only turn over the rotating die, move the corresponding dies on different sides of the rotating die to the working position, make one rotating die manufacture multiple different state skins, save the floor area, and make two skins be processed and formed at one time through the up-down reciprocating movement of the hydraulic cylinder, thereby improving the working efficiency.

A propulsion system coupled to a vehicle. The system includes a convex surface, a diffusing structure coupled to the convex surface, and at least one conduit coupled to the convex surface. The conduit is configured to introduce to the convex surface a primary fluid produced by the vehicle. The system further includes an intake structure coupled to the convex surface and configured to introduce to the diffusing structure a secondary fluid accessible to the vehicle. The diffusing structure comprises a terminal end configured to provide egress from the system for the introduced primary fluid and secondary fluid.

The application discloses a folding rudder wing operating mechanism, which comprises a mounting seat body, a cavity is formed in the middle position of the mounting seat body, a rudder handle is mounted in the upper end of the cavity, a folding rudder wing is connected to the end of the rudder handle away from the cavity, an adapter plate is mounted outside the lower end of the cavity, a push rod is connected to the adapter plate, a linkage mechanism is arranged in the cavity, and the linkage mechanism is used for synchronously rotating the adapter plate and the folding rudder wing; the overall structure has a short thrust transmission path, and the response speed of folding rudder wing operation can be effectively improved.

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.

The present disclosure relates to systems and methods of distributing propulsion load power drawn from high-energy and high-power batteries. Specifically, a system and method of distributing load power drawn from a plurality of batteries for electric propulsion of a vehicle. The system includes: a high-energy battery and a high-power battery designed for optimal production of DC power during high specific energy propulsion and high specific power propulsion, respectively; and a battery health management system configured to monitor state of charge and state of health of the batteries and generate battery state signals. The system further includes a propulsion load configured to produce propulsion force using power converted from power generated by at least one of the batteries; and a system controller configured to distribute load power drawn from the high-energy battery and the high-power battery for use by the propulsion load according to a propulsion phase of the vehicle and the battery state.

The application provides a multifunctional breather valve for a helicopter fuel system, which comprises a spring 1, a valve assembly, a valve base 8, and a shell 11, wherein the valve assembly comprises a valve cavity shell 2, a float 3, a float cavity 4, a sealing ring 1 5, a valve cavity base 6, a sealing ring 2 7, a sealing ring 3 9, and a steel ball 10, and wherein the valve base 8 is a hollow cylindrical structure, the bottom end of the valve base 8 is a mounting flange structure, and the upper end of the valve base 8 is provided with an outwardly turned boss structure; the shell 11 is a hollow cylindrical structure, the boss of the upper end of the valve base 8 is assembled in the bottom end of the shell 11, and the top end of the shell 11 is the outlet of the multifunctional breather valve; the valve base 8 and the shell 11 constitute a breather valve shell, the internal cavities of the valve base 8 and the shell 11 are provided with the spring 1 and the valve assembly; one end of the spring 1 is fixed on the upper part of the shell 11, and the other end is fixed on the upper part of the valve cavity shell 2; and the valve assembly is pressed on the valve base 8 under the action of the spring 1.

This invention proposes a large-diameter oil passage valve, comprising an oil supply pipeline and a piston cylinder, with a flap valve hinged to the oil supply pipeline. A connecting rod is provided between the piston cylinder and the flap valve to enable linkage between the flap valve and the piston. The oil inlet end of the oil supply pipeline communicates with a cavity on one side of the piston in the piston cylinder, while an oil bypass is provided on the cavity on the other side of the piston. A control valve is connected to the oil bypass to change the pressure difference on both sides of the piston, thereby driving the rotation of the flap valve. This invention significantly reduces the internal space through the lever principle, effectively reducing weight and improving working efficiency. By employing a self-locking valve, a pilot-operated self-locking electromagnetic oil passage valve is constructed. The pilot part uses a small-volume self-locking valve to control the on / off state, avoiding the system power loss and electromagnetic coil overheating problems caused by the long-term energization of conventional electromagnets.

This invention discloses a symmetrical non-rotational composite material oblique mesh structure suitable for automated molding, comprising an upper frame, a lower frame, and bidirectional helical ribs; the upper frame, lower frame, and bidirectional helical ribs are integrally molded structures, and the structural materials used for the upper frame, lower frame, and bidirectional helical ribs are composite materials; the bidirectional helical ribs include two mutually symmetrical ribs, and the center line of the outer contour surface of each rib is a helix with equal pitch, equal helix angle, or geodesic. This invention further discloses methods for determining the center line of ribs used in various typical symmetrical non-rotational structures applied in aerospace engineering. The oblique mesh structure of this invention is rationally designed, lightweight, and suitable for automated mass production.

A low-power consumption leg type landing gear and helicopter based on tether coupling relates to the technical field of helicopter landing gear. The existing helicopter landing gear has high energy consumption in the process of ground parking or air hovering and attitude locking, which additionally increases the power consumption of the helicopter and leads to the decline of the endurance. The present application comprises a support frame and at least one set of landing support unit; each set of landing support unit contains a support leg, a differential mechanism, a hip joint rope drive mechanism and a knee joint drive mechanism. Two support legs are symmetrically arranged on both sides of the support frame for helicopter landing support; the differential mechanism corresponds to the support leg one by one to regulate the hip joint angle; the output power of each hip joint rope drive mechanism is divided into two ways and transmitted to the differential mechanism, and after differential adjustment, the support leg is driven to realize angle adjustment; the knee joint drive mechanism adjusts the knee joint angle of the support leg. The tether coupling and differential adjustment structure are adopted to reduce the energy consumption while meeting the landing support and terrain adaptation, and to improve the endurance of the helicopter.

The utility model relates to the field of unmanned plane connecting frame, disclose unmanned plane fuselage and foot stand are connected with integral type connecting piece, including second connecting sleeve, the upper end surface center of second connecting sleeve is provided with first connecting sleeve, the lower extreme of first connecting sleeve is through the upper end surface of second connecting sleeve and leads to inside, the upper end surface of second connecting sleeve and the two side wall junctions of first connecting sleeve are provided with first connecting buckle respectively in first connecting sleeve front and back two sides, the side wall center of second connecting sleeve is provided with third connecting sleeve, one end of third connecting sleeve is through the side wall of second connecting sleeve and the side wall of first connecting sleeve and leads to inside respectively. In the utility model, through setting up as integral moulding type structure between first connecting sleeve, second connecting sleeve and third connecting sleeve, integral type design improves production efficiency, strengthens fuselage structure strength, improves airplane stability and reliability.

This invention discloses a chair seat assembly method. By simultaneously driving two symmetrically arranged positioning plates to move synchronously from top to bottom and from outside to inside, the bottom shell is clamped and pressed down. The relative position of the bottom shell and the seat frame can be continuously adjusted, as well as the posture of the seat frame on the bottom plate. Adjusting the relative posture of the seat frame and the bottom shell quickly makes the first connecting hole of the chair frame and the second connecting hole of the bottom shell coaxial, resulting in better consistency of chairs produced in batches. Furthermore, during the position adjustment process, the positioning plates do not compact the seat frame, thus not affecting the position, posture, and support effect of the mesh on the seat frame.

The application belongs to the field of composite materials, and particularly relates to a heat-proof and sound-absorbing structure integrated composite material with low manufacturing cost, heat-proof effect meeting use requirements and weight reduction requirements and a preparation method thereof. The heat-proof and sound-absorbing structure integrated composite material comprises an external surface layer and an internal surface layer, a sound-absorbing and heat-proof layer is arranged between the external surface layer and the internal surface layer, the external surface layer comprises a surface layer and at least one inner layer, the surface layer is made of polysilazane modified silicium-containing arylacetylene resin, oxide powder and quartz fiber, the inner layer is made of boron phenolic resin, oxide powder and quartz fiber, the content of the oxide powder in the inner layer is less than that in the surface layer, the integrated composite material has heat-proof and sound-absorbing capabilities, meets the requirements of multiple functions, has low manufacturing cost and good structural strength.

A system includes a first propeller, a second propeller, an electromagnetic field transmitter coupled to the first propeller, and an electromagnetic field receiver coupled to the second propeller. The electromagnetic field transmitter emits an electromagnetic field, and in response to the electromagnetic field, the electromagnetic field receiver and the second propeller rotate in a first rotational direction, and the electromagnetic field transmitter and the first propeller rotate in a second, opposite rotational direction.

An electric aircraft comprising: a fuselage; one or more wings mounted on the fuselage; a plurality of propellers pivotable between lift and forward propulsion configurations; and a plurality of battery packs, wherein at least two of the plurality of battery packs are configured to power: at least a portion of a first propeller of the plurality of propellers and at least a portion of a second propeller of the plurality of propellers, which is different from the first propeller.

This invention discloses a test method for rapidly adjusting the impact trajectory of a hydraulic lock with an actuator cylinder. The loading force is controllable, and the impact trajectory is adjustable. This is achieved through the following technical solution: an open-end lock ring box equipped with a hydraulic lock is connected to a force sensor via a load transmission shaft. The load transmission shaft is coaxially connected to a load loading mechanism via a hinge shaft at its end and passing through the first support plate. The load loading mechanism is connected to the equipment loading cylinder and, via a hinge shaft, to an impact angle adjustment hinge through the first support plate. A force sensor coaxially mounted on the load transmission shaft and an open-end lock ring box connected to its end are also connected. A hydraulic lock mounting fixture is laterally connected to the first, second, and third claw disc shafts via bearing seats and connected to the open-end lock ring box via a stepped shaft groove on the second support plate. Together, these components constitute a hydraulic lock loading test device with an actuator cylinder and its impact trajectory adjustment mechanism. This invention can cover the testing of all hydraulic locks with impact radii within the cylinder stroke range.

This invention provides a bonding and assembly mold for carbon fiber skin honeycomb panels used in satellites, including a pressure equalizing plate, a positioning base plate, a pad, a retaining strip, retaining strip gaskets, fastening screws, positioning screws, and limiting pins. The pad and positioning base plate are placed sequentially on a roller platform, with their openings aligned. Fastening screws are then passed through the positioning base plate and retaining strip sequentially to install the retaining strip onto the positioning base plate. Limiting pins are then passed through the retaining strip and positioning base plate sequentially to limit the retaining strip's position. The retaining strip gasket is attached to the inner side of the retaining strip. A lower skin with adhesive film applied is placed in the mold, with the adhesive film facing upwards. The honeycomb core is then placed. The positions of the front and side embedded parts are confirmed, and the honeycomb cores at the corresponding positions are cut. Positioning screws are then passed through the positioning base plate and lower skin sequentially to position the front embedded part. Positioning screws are then passed through the retaining strip and retaining strip gasket sequentially to position the side embedded part. An upper skin with adhesive film applied is placed, with the adhesive film facing downwards. The pressure equalizing plate is placed, and positioning screws are then passed through the pressure equalizing plate and upper skin sequentially to position the front embedded part. This invention can be used for the bonding, assembly, and high-temperature curing of carbon fiber skin honeycomb panels for satellites.