3218results about "Weight reduction" patented technology

An electrical propulsion system for a vertical take-off and landing (VTOL) aircraft comprises an electrical motor assembly and an inverter assembly. The inverter assembly comprises a housing, a capacitor assembly, at least one printed circuit board assembly (PCBA), and a plurality of positioning pins. The capacitor assembly comprises a center hole, at least one capacitor, a capacitor housing having at least one busbar, and a plurality of through holes in the capacitor housing. The capacitor assembly and the at least one PCBA are positioned inside the housing. The plurality of positioning pins pass through the through the plurality of through holes of the capacitor housing and the at least one PCBA and are connected to the housing.

A structural shroud for a passenger seat assembly includes a hollow body having a first end, a second end opposite from the first end, and a central cavity between the first end and the second end. In certain embodiments, the structural shroud may include a magnesium alloy and / or an aluminum alloy, and may be formed via casting, forging, and / or additive manufacturing, among others. In some embodiments, a base frame assembly for a passenger seat assembly includes a monolithic base frame having a first leg portion and a first tube portion integrally formed with the first leg portion.

The invention relates to a fiber continuous longitudinal and transverse reinforced integral composite material skin structure which is formed by co-bonding an upper skin (3), a lower skin (1) and a longitudinal and transverse reinforced structure (2). The invention further relates to a forming method of the fiber continuous longitudinal and transverse reinforced integral composite material skin structure. The forming method comprises the step of forming the longitudinal and transverse reinforced structure (2). Paving the skin and positioning the longitudinal and transverse reinforcing structure (2); and curing the paved fiber continuous longitudinal and transverse reinforced integral composite material skin structure, machining a circular opening in each of the left side and the right side of the longitudinal and transverse reinforced structure (2), and completing the whole forming process of the fiber continuous longitudinal and transverse reinforced integral composite material skin structure. The skin structure is light in weight and continuous in fiber, the structural strength is effectively enhanced through integral structure cementing co-curing forming, the integral structure is more excellent in rigidity performance, and the embedded longitudinal and transverse reinforcing structure is high in size and position precision.

The invention belongs to the field of hydraulic transmission system design, and discloses a hydraulic transmission system protection method and system.The protection method comprises the steps that onboard state signals, flight control system state signals and first hydraulic pressure and second hydraulic pressure of each hydraulic pipeline in a hydraulic transmission system are collected in real time; according to the onboard state signal, an electric pump closing instruction and a hydraulic pressure logic judgment instruction are output through an onboard state threshold value and first electric pump control logic, and an electric pump connected to a hydraulic pipeline is closed according to the electric pump closing instruction; and acquiring a flight control system state signal, first hydraulic pressure and second hydraulic pressure according to the hydraulic pressure logic judgment instruction, and automatically controlling the electric pump by adopting a second electric pump control mechanism and a hydraulic pipeline protection mechanism through a low-pressure threshold value and a pressure safety threshold value. According to the invention, the starting of each electric pump can be automatically controlled, a hydraulic passage can be protected through a plurality of mechanisms, and the flight safety can be improved.

The invention provides a bulkhead woven by a composite material and a preparation process thereof, and belongs to the field of bulkhead preparation, and the preparation process comprises the following steps: splicing two half mandrels into a whole mandrel with a rectangular cross section; clamping the whole core mold, and knitting fiber yarns on the whole core mold through a three-dimensional knitting machine, so as to prepare a prefabricated body with a rectangular section; heating the prefabricated body, and carrying out preliminary curing on the prefabricated body, so as to prepare a preliminarily cured prefabricated body; after the preliminarily solidified prefabricated body is cooled, the preliminarily solidified prefabricated body is divided into two U-shaped prefabricated bodies along the cutting plane; one vertical surface of the U-shaped prefabricated body is turned over by more than 90 degrees; and the folded U-shaped prefabricated body is cured again, so that the bulkhead is prepared. According to the invention, the braiding machine is used for braiding the prefabricated body, preliminarily curing the prefabricated body, cutting the prefabricated body, folding the prefabricated body and re-curing the prefabricated body, so that fiber wrinkles, fiber gaps and fiber angles at the bent part of the bulkhead can be better controlled.

An autonomous flying device achieving a large payload and a long continuous flight time and also accurately adjust position and orientation while flying. The device includes: a main rotor and the like that provide main thrust; a sub rotor and the like that controls the orientation; an engine that generates energy for rotating the main rotor and the like and the sub rotor and the like; and an arithmetic control device that controls rotation of the sub rotor and the like. Also, the main rotor and the like are rotated by being drivingly connected to the engine, whereas the sub rotor and the like are rotated by motors driven by electric power generated from generator and the like operated by the engine. Further, when orientation control to tilt the fuselage is performed, the arithmetic control device increases the output distribution ratio of the sub rotor to above the output distribution ratio of the sub rotor when hovering is performed.

The invention provides a high-lightning-stroke-resistant toughness continuous carbon fiber reinforced composite material, which is obtained by paving a plurality of layers of continuous carbon fiber resin prepregs and intercalation layers to obtain a preform and curing the preform through a conventional composite material molding process, each resin-based prepreg layer used in a preset layer number range close to the anti-lightning side is a first resin prepreg layer, the other resin-based prepreg layers are second resin prepreg layers, an intercalation is arranged between every two adjacent second resin prepreg layers, and the intercalation material is a micron conductive material with the surface density of 2-20 g / m < 2 >; wherein the resin in each resin-based prepreg layer is added with 2-10wt% of carbon nanomaterial, and the compressive strength of the rest resin-based prepreg layers after impact is higher than 250MPa. The composite material provided by the invention realizes integrated compounding of high lightning stroke resistance and high impact toughness. The anti-lightning performance of the composite material can meet the requirements of no internal layering and low fiber damage on the surface, and meanwhile, the compression strength CAI after impact can reach more than 200MPa.

A vertical take-off and landing (VTOL) aircraft provides transportation to users of a network system. The network system may include multiple aircraft or other types of vehicles to provide multi-model transportation. An aircraft may include a fuselage, a truss coupled to the fuselage, and multiple distributed electric propellers coupled to the truss. The distributed electric propellers may be positioned on at least two different planes. The fuselage may include a cabin having one or more seats for the passengers arranged in a configuration that has a compact footprint, provides legroom, provides visibility to surroundings of the aircraft, or facilitates convenient ingress or egress of passengers. The aircraft may open a port cabin door and starboard cabin door for simultaneous ingress or egress of passengers.

The utility model discloses an unmanned aerial vehicle undercarriage, relates to the technical field of unmanned aerial vehicles, and provides the following scheme that the unmanned aerial vehicle undercarriage comprises a connecting plate and a buffer mechanism arranged at the bottom of the connecting plate, the buffer mechanism comprises a working assembly and a buffer assembly, the working assembly comprises a supporting rod, and the outer wall of the connecting plate is fixedly connected with the supporting rod; a hollow column is fixedly connected to the bottom of the supporting rod, a sliding column is slidably connected to the inner wall of the hollow column, the buffering assembly comprises a first spring, the first spring is arranged on the outer wall of the sliding column, a connecting ring is fixedly connected to the bottom end of the sliding column, and a supporting head is fixedly connected to the bottom of the connecting ring. A user descends the equipment for parking, a supporting head can be inserted into the ground, a certain fixing effect is achieved, at the moment, a sliding column can retract into a sliding groove formed in a hollow-out column, the buffering and supporting effect is achieved through a first spring, the impact force of the unmanned aerial vehicle and the ground is reduced, and the internal equipment is prevented from being damaged by vibration.

Apparatus, systems, and methods are contemplated for electric powered vertical takeoff and landing (eVTOL) aircraft. Such are craft are engineered to carry safely carry at least 500 pounds (approx. 227 kg) using a few (e.g., 2-4) rotors, generally variable speed rigid (non-articulated) rotors. It is contemplated that one or more rotors generate a significant amount of lift (e.g., 70%) during rotorborne flight (e.g., vertical takeoff, hover, etc), and tilt to provide forward propulsion during wingborne flight. The rotors preferably employ individual blade control, and are battery powered. The vehicle preferably flies in an autopilot or pilotless mode and has a relatively small (e.g., less than 45′ diameter) footprint.

One variation of a system for generating thrust at an aerial vehicle includes: a primary electric motor; a rotor coupled to the motor; an internal-combustion engine; a clutch interposed between the motor and an output shaft of the internal-combustion engine; an engine shroud defining a shroud inlet between the rotor and the internal-combustion engine, extending over the internal-combustion engine, and defining a shroud outlet opposite the rotor; a cooling fan coupled and configured to displace air through the engine shroud; and a local controller configured to receive a rotor speed command specifying a target rotor speed, adjust a throttle setpoint of the internal-combustion engine according to the target rotor speed and a state of charge of a battery in the aerial vehicle, and drive the primary electric motor to selectively output torque to the rotor and to regeneratively brake the rotor according to the target rotor speed.

The invention relates to the technical field of aero seats, and discloses an aero seat angle adjusting device which comprises a cushion support, a supporting shaft is fixedly installed at the rear end of the cushion support, a backrest support is rotationally connected to the supporting shaft, a hip blocking piece is arranged on the front half portion of the cushion support, and a hip blocking piece is arranged on the hip blocking piece. Pushing assemblies are arranged on the rear sides of the left end and the right end of the hip blocking piece, and a waist supporting piece is arranged on the lower half portion of the backrest support. The output end of the air cylinder extends to drive the sliding base to move forwards along the sliding groove, the first steel wire rope and the second steel wire rope are pulled, and the inclination angle of the backrest support is adjusted. The push rod on the right side is pushed forwards through the sliding seat, so that the first supporting rod and the second supporting rod are driven to deflect; the soft wrapping layer of the cushion support forms a slope face with the front portion higher than the rear portion to support the buttocks of a passenger, in the rotating process of the rotating pipe, the two threaded lantern rings are driven to be close to each other, the waist supporting plate is pushed to move forwards, the inclined rib plate is supported on the waist of the passenger, and the passenger can sit more comfortably.

There is provided a structurally integrated vacuum tank that includes a vacuum tank main portion extending between vacuum tank end portions. The vacuum tank main portion includes a vacuum tank skin forming a cylinder. The vacuum tank skin has a longitudinal cross section with a profile geometry configured for buckling prevention for the vacuum tank skin under external pressure loads. The vacuum tank skin is configured to provide a pressure barrier between an outside ambient pressure and a vacuum in an interior of the vacuum tank main portion. The vacuum tank main portion further includes a plurality of stiffener members coupled to surface portions of the vacuum tank skin. The vacuum tank skin and the plurality of stiffener members are configured to carry structural loads.

The invention discloses a solid rocket engine composite material skirt connecting structure and a forming method thereof.The connecting structure comprises a composite material skirt, a skirt hole reinforcing ring, a connected cabin section and a connecting piece, the composite material skirt and the skirt hole reinforcing ring are integrally formed, and the rear end of the composite material skirt and the rear end of the skirt hole reinforcing ring are connected with the connected cabin section through the connecting piece; the connected cabin section is of an L-shaped nested inner lap joint structure and has the butt joint limiting and bearing functions, and the thickness transition section at the front end of the composite skirt is connected with a fiber winding structure layer of a combustion chamber shell. Compared with a traditional metal skirt connecting structure, the weight reduction effect is obvious, compared with a traditional composite skirt hinge block connecting structure, the connecting mode is simpler, and the fiber continuity of a composite skirt connecting area is guaranteed; the problems that in-plane shear rigidity and strength of the composite material are insufficient, and orifice positions are prone to damage are solved. The structure and the forming process are simple, the maturity is high, and the cost is low.

A structural shroud assembly for a passenger seat includes a structural shroud having a first end, a second end opposite from the first end, and a central cavity between the first end and the second end. An armrest pivot may be supported at the first end of the structural shroud. In some cases, the structural shroud may be an armrest. In other cases, the structural shroud is an upper spreader for the passenger seat configured to support a seat back of the passenger seat, and a lower spreader configured to receive at least one seat beam tube is attached to the upper spreader.

Apparatus, systems, and methods are contemplated for electric powered vertical takeoff and landing (eVTOL) aircraft. Such are craft are engineered to carry safely carry at least 500 pounds (approx. 227 kg) using a few (e.g., 2-4) rotors, generally variable speed rigid (non-articulated) rotors. It is contemplated that one or more rotors generate a significant amount of lift (e.g., 70%) during rotorborne flight (e.g., vertical takeoff, hover, etc), and tilt to provide forward propulsion during wingborne flight. The rotors preferably employ individual blade control, and are battery powered. The vehicle preferably flies in an autopilot or pilotless mode and has a relatively small (e.g., less than 45′ diameter) footprint.

The invention discloses a method and tool capable of improving the glue joint quality of a sharp corner area of a composite workpiece, and relates to the technical field of composite workpiece machining. The method comprises the following steps of prefabricating the corner filler, notching the corner filler, pre-filling a sharp corner area, assembling and curing. The tool comprises two R positioning plates which are arranged in an overlapped mode, the upper edges of the opposite faces of the two R positioning plates are matched with the side walls of the two sides of the finished corner filler respectively, a blade clamped and fixed through the R positioning plates is vertically arranged between the two R positioning plates, and the upper end of the blade extends to the positions above the upper edges of the opposite faces of the two R positioning plates. According to the method and the tool capable of improving the glue joint quality of the sharp corner area of the composite material workpiece, the sharp corner area is ensured to be fully filled by optimizing the shape of the corner filling material and a pretreatment process, so that cavities are thoroughly eliminated, and the glue joint quality and reliability of the sharp corner area of the composite material workpiece are improved.

An electrical propulsion system comprises an inverter, an electrical motor assembly, an assembly, and a rotor position sensor. The inverter comprises a printed circuit board assembly (PCBA). The electrical motor assembly comprises a stator and a rotor. The assembly is configured to rotate a propeller and comprises a moving component and a stationary component. The rotor position sensor comprises at least one sensor coupled to the PCBA, and a magnet located on the moving component. The at least one sensor is configured to detect a magnetic field of the magnet through the stationary component.

The invention provides a composite cabin door with a combined sandwich structure. A sandwich layer in a special-shaped area of the cabin door is made of a foam sandwich material; the sandwich layer in the installation area of the accessories of the cabin door is made of filler, and the accessories need to be installed through screws and bolts; the sandwich layer in the load-bearing area of the cabin door is made of a high-density honeycomb material; a sandwich layer in a non-bearing area of the cabin door is made of a low-density honeycomb material; the density and strength of the low-density honeycomb material are smaller than those of the high-density honeycomb interlayer, so that the non-bearing area is designed to be light. The total weight of the cabin door can be reduced to the greatest extent under the condition that the equipment mounting requirement and the rigidity and strength requirements are met; meanwhile, metal parts exposed in the air are reduced, and the problems of corrosion and abrasion of cabin door parts are solved; and splicing areas of different sandwich layer materials are locally bonded and reinforced, so that the damage risk of the composite cabin door is reduced.

The invention discloses a composite material T-shaped beam forming method in the technical field of aviation, a forming die, a laying die and a bottom face forming die which are arranged in a split mode and can be combined are adopted, forming is conducted through a secondary hot-pressing curing process, and the method comprises the following steps that S1, preforming and R corner prefabricating are conducted, the independently-prefabricated prepreg core material is positioned in a T-shaped beam R-angle area formed after the prepreg core material and the T-shaped beam R-angle area are oppositely combined; and S2, first-time mold closing and pre-curing are carried out, specifically, the laying mold and the forming mold are oppositely closed, lateral auxiliary positioning is carried out through the bottom face forming mold, locking and fixing are carried out through a positioning pin and a clamping piece, and first-time hot-pressing curing is carried out on a combined body obtained after mold closing. According to the method, the independently prefabricated core material which is subjected to heat-sealing, pressing and pre-curing is filled in the R-angle area, and a secondary hot-pressing process of pre-curing and final curing is matched, so that sufficient pressure transmission and tight laying of the R-angle area are ensured, the defects of layering, bridging and the like are effectively avoided, and the percent of pass of nondestructive testing of products is remarkably improved.

According to the hard process cover plate forming method capable of achieving directional control over the R angle thickness of the composite part, the carbon fiber fabric prepreg is laid on the hard cover plate forming tool to prepare a hard cover plate, and the inner R angle of the hard cover plate is compensated to adapt to R angle forming of parts with different thicknesses. By adjusting the value of the R angle in the hard cover plate, the part prepreg is prevented from being excessively extruded in the R angle in the forming process, so that glue flowing is serious, the forming quality of the R angle of the part is guaranteed, and the percent of pass of the composite material part with the R angle is effectively improved.

The invention belongs to the technical field of vibration control, and particularly relates to a composite material laminated board structure embedded with NiTi alloy wires and a preparation method and application of the composite material laminated board structure. The preparation method comprises the following steps: respectively fixing two ends of a plurality of NiTi alloy wires on a wire embedding mold to form an alloy wire surface; the first layer plate part and the second layer plate part are symmetrically arranged on the upper surface and the lower surface of the alloy wire surface, an initial pressing bottom plate is arranged below the first layer plate part, and an initial pressing top plate is arranged above the second layer plate part; first pressure is applied to the initial pressing top plate and the initial pressing bottom plate, so that the first layer plate part, the alloy wire surface and the second layer plate part form a prefabricated structure body; and high-temperature and high-pressure curing molding. The composite material laminated board structure is prepared by the method. The laminated board structure can be applied to a wall board of a new energy aircraft. The composite material laminated board structure provided by the invention is light in weight and good in fatigue resistance, the vibration resistance of the composite material structure is enhanced, and the composite material laminated board structure can be integrally processed.

The present invention provides a long-short carbon nanotube reinforced and toughened fiber composite material and a preparation method thereof. The preparation method provided by the present invention comprises: a) mixing short carbon nanotubes, a thermosetting resin and an additive to obtain a resin-based slurry; b) pouring the resin-based slurry into a fiber preform and curing and forming it to obtain a long-short carbon nanotube reinforced and toughened fiber composite material; wherein: the fiber preform comprises: an upper fiber cloth layer, a long carbon nanotube fiber yarn layer, and a lower fiber cloth layer stacked and contacted in sequence; the long carbon nanotube fiber yarn layer is a yarn-like structure formed by long carbon nanotubes; the length of the short carbon nanotubes is 0.5 to 3 μm, and the average length is ≤2 μm; the length of the long carbon nanotubes is 50 to 1000 μm, and the average length is >100 μm. The present invention optimizes the spatial layout of long and short carbon nanotubes in the composite material, simultaneously achieving the dual purposes of intra-layer reinforcement and inter-layer toughening of the fiber composite material.

A movable seat unit, including armrest side plates supported on the ground; a mechanical extension device including a seat frame configured to attach a seat portion and a backrest assembly disposed at a rear portion of the seat frame and configured to attach a backrest, where guide rails extending in a front-rear direction of the seat unit are provided on the armrest side plates, and the seat frame is slidably engaged with the guide rails; the backrest assembly includes a backrest mounting member and a guide chute member fixedly mounted relative to the armrest side plates, a lower end of the backrest mounting member is pivotally connected to a rear portion of the seat frame, and a middle portion of the backrest mounting member is provided with a guide pin configured to slidably engage with the guide chute.

The invention belongs to the technical field of grating stiffened wall plates, and particularly relates to a grating stiffened wall plate for an aircraft structure and a forming mold and method.The grating stiffened wall plate comprises a grating frame, a skin plate is arranged on the grating frame, and the skin plate comprises an outer skin arranged on the bottom face of the grating frame and an inner skin arranged on the grating frame; the grille frame comprises a transverse rib core material and a longitudinal rib core material which are connected with each other, the inner skin is located between the transverse rib core material and the longitudinal rib core material, the connecting position between the transverse rib core material and the longitudinal rib core material is called as a connecting node, and each connecting node is provided with a forming block; the strength and the stability of the grating frame are effectively enhanced through the forming blocks. In addition, the transverse rib core materials and the longitudinal rib core materials are tightly fixed through flanging wrapping of the inner skin, and therefore the strength and the structural efficiency of the grating stiffened wall plate are remarkably improved. And the unique requirements of aircraft wall plate structures such as airplanes and the like on the aspects of transmitting and bearing large loads are met.

An electrical propulsion system for a vertical take-off and landing (VTOL) aircraft comprises an electrical motor assembly and an inverter assembly. The inverter assembly comprises a housing, a capacitor assembly, at least one printed circuit board assembly (PCBA), and a plurality of positioning pins. The capacitor assembly comprises a center hole, at least one capacitor, a capacitor housing having at least one busbar, and a plurality of through holes in the capacitor housing. The capacitor assembly and the at least one PCBA are positioned inside the housing. The plurality of positioning pins pass through the through the plurality of through holes of the capacitor housing and the at least one PCBA and are connected to the housing.

An unmanned aerial vehicle (UAV) having wings stowed against a fuselage of the UAV in a first arrangement is disclosed. Methods and systems for deploying the wings into a second arrangement are disclosed. For example, after a launch of the UAV, the UAV monitors for at least one precondition. The at least one pre-condition being a pre-condition associated with deploying wings of the UAV into the second arrangement. Upon detecting the at least one precondition, the wings of the UAV are deployed into a second arrangement. Deploying the wings comprises activating, in response to detecting the at least one precondition associated with the UAV, a gearbox configured to transition the wings from the first arrangement to the second arrangement. Roll control may be maintained throughout launch and deployment.

The utility model discloses a flap control system and an airplane. The flap control system comprises a first torque tube, a second torque tube, a flap actuator, a driving rocker arm and a coupling connecting piece. The coupling connecting piece is provided with a first end, a second end and a connecting section arranged between the first end and the second end, and the first end and the second end are fixed to the first torque tube and the second torque tube respectively; one end of the driving rocker arm is fixed to the connecting section, the other end of the driving rocker arm is connected with the flap actuator, and when the flap actuator works, the driving rocker arm can be driven to rotate, and meanwhile the first torque tube and the second torque tube on the coupling connecting piece are driven to rotate synchronously. According to the scheme, the principle and the structure are simple, and the flap actuator can conveniently drive the left flap and the right flap of the aircraft to work. The flap control system provided by the utility model is used for replacing a complex synchronous cross-linking mechanism and electrical control, so that the flap control is not easy to make mistakes, and the control is more reliable.

The invention discloses an aircraft movable part hinge joint reaming method, which adopts an aircraft movable part hinge joint reaming tool to assist in reaming a hinge joint connected between an aircraft movable part and an aircraft fixed part, and the aircraft movable part hinge joint reaming tool comprises a tool main body, and at least two groups of hinge joint positioning assemblies, a plurality of tool positioning pieces and a plurality of airplane fixing part positioning pieces which are in one-to-one correspondence with the hinge joints are detachably fixed on the tool main body. According to the invention, the coaxiality of reaming holes in different hinge joints is effectively ensured, and the mounting precision of movable parts of an aircraft is effectively improved; the reaming tool for the hinge joint of the aircraft movable part can be repeatedly used, reaming positioning is convenient and fast, the assembly cost of the aircraft movable part is reduced, and meanwhile the assembly efficiency of the aircraft movable part is effectively improved.

The utility model discloses an aircraft single wallboard assembly fixture which comprises a main frame, a clamping plate assembly, a stringer end positioner, a stringer axis surface positioner, a stringer puller, a K hole positioner, a skin supporting block, a floating frame presser and a bandage, and the main frame comprises a frame, an adjustable support, a clamping plate support and a laser tracker target ball positioning seat. The laser tracker target ball mounting seat and the frame are fixedly mounted, the clamping plate supports are located at the upper end and the lower end of the frame, the adjustable support is arranged at the bottom of the frame, and the long honing and the floating frame can be conveniently fixed.