2400results about "Weight reduction" patented technology

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.

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.

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 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 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.

A vertical take-off and landing (VTOL) aircraft is disclosed. The VTOL aircraft may include: a fuselage, a thrust rotor to produce a propulsion thrust, rotor booms, canard surfaces, wing surfaces, tail surfaces, and lift rotors to produce lifting thrust force.

An aircraft assembly is disclosed having a wing and an engine mounting pylon. An aft end of the engine mounting pylon is connected to the wing by a spigot and at least one fastener. The aircraft assembly is configured such that, during operation of the aircraft assembly on an aircraft, the spigot transfers only lateral load between the engine mounting pylon and the wing and the at least one fastener transfers only vertical load between the engine mounting pylon and the wing.

The present invention provides a welding positioning device for an engine pipe body bracket of a drone, which includes a first positioning bracket, a second positioning bracket and a third positioning bracket, and is respectively used for accurately positioning the first pipe body, the second pipe body and the third pipe body of the engine pipe body bracket. Each pipe body penetrates into a pin column with holes to form an assembly base point of the engine. Bolts and nuts are used to penetrate into the pipe body pin column holes to position and fasten the pipe bodies, which can ensure the positioning accuracy of the engine assembly base point and the reliability of the positioning method. According to the correlation between the pipe bodies of the drone bracket, the engine bracket welding positioning device can combine the pipe bodies to form an accurate positioning of the bracket. By accurately positioning the engine bracket, the present invention enables the engine bracket to meet the accuracy requirements specified in the design drawings, meet the installation conditions for the assembly of the drone engine, ensure the stability of the engine function, and provide reliable power support for the drone.

The invention discloses an aircraft fairing and a fairing body composite material integrated forming process, and relates to the technical field of aviation equipment manufacturing. The forming process comprises the following steps: 1) preparing a prepreg; (2) sectional forming; 3) a curing process; (4) demolding and trimming; 5) an assembly process; the whole of the front section, the middle section and the tail section of the fairing is of a long-size U-shaped streamline structure. According to the invention, through molecular collaborative design of fluorination modification-polysiloxane toughening-cyanate ester interpenetration, the key contradiction of a high-performance fairing material is systematically solved; ultrahigh strength and toughness are realized at the same time by utilizing in-situ nano microphase separation; a dense graphite carbon layer is generated through catalysis, so that efficient halogen-free flame retardance is achieved; by means of the partition design of molecular fluorination and surface metallization, the high wave transmission and lightning protection functions are ingeniously unified. And finally, a new generation of aircraft fairing composite material with force bearing, wave transmission, heat resistance, impact resistance and environmental stability is successfully prepared.

The invention belongs to the technical field of airplane structure design, and discloses a horizontal tail wing beam connecting structure, an airplane and an assembling method, and the horizontal tail wing beam connecting structure comprises a central rib plate, an upper T-shaped connector, a lower T-shaped connector, a left connecting plate, a right connecting plate and an outer side connecting plate. The upper T-shaped joint comprises a lower edge strip, a first left edge strip and a first right edge strip, the lower T-shaped joint comprises an upper edge strip, a second left edge strip and a second right edge strip, the central rib plate is arranged between the left beam and the right beam, and the left connecting plate, the right connecting plate and the outer side connecting plate are combined for use, so that the bearing capacity of a left beam web plate and a right beam web plate is improved; the first left edge strip and the first right edge strip can increase the contact area between the upper T-shaped connector and the left wing and the contact area between the upper T-shaped connector and the right wing, the second left edge strip and the second right edge strip can increase the contact area between the lower T-shaped connector and the left wing and the contact area between the lower T-shaped connector and the right wing, the connecting strength of the left beam and the right beam in the vertical direction is improved, and the lower edge strip, the upper edge strip and the central rib plate are matched with one another. And the connection strength of the left beam and the right beam in the horizontal direction is improved.

The present invention relates to a retractable multi-rotor UAV platform, which belongs to the technical field of UAVs and solves the problems of low folding automation and insufficient storage space for multi-rotor UAVs. The retractable multi-rotor UAV platform of the present invention comprises a fuselage component, an arm component, a propeller component, a landing gear component, an arm folding component and a propeller retracting component; one end of the arm component is hinged to the periphery of the fuselage component through the arm folding component; a plurality of arm components are evenly distributed on the periphery of the fuselage component in the horizontal plane; the propeller component is movably connected to the other end of the arm component through the propeller retracting component, and the propeller component is retracted or unfolded relative to the fuselage component in the vertical plane; the landing gear component is connected to the lower end of the fuselage component. The present invention realizes the horizontal folding of the arm controlled by the servo and the vertical retraction of the unpowered propeller, which reduces the storage space of the UAV to the greatest extent, and the retracting process saves energy and has a high degree of automation.

A stowage bin assembly includes a bin bucket assembly having a bin bucket which includes at least one sidewall, the at least one sidewall including a notch and the at least one sidewall having an interior surface. The stowage bin assembly also includes at least one hinge assembly. The at least one hinge assembly includes a housing which is received in the notch, the housing having an interior surface, wherein the interior surface of the housing is substantially flush with the interior surface of the at least one sidewall when the housing of the at least one hinge assembly is received in the notch. The stowage bin assembly also includes a stowage bin door pivotably coupled to the bin bucket assembly via the at least one hinge assembly, the stowage bin door moveable between an open position and a closed position. Related components, assemblies, and devices are also provided.

The utility model discloses an aerial football unmanned aerial vehicle frame, which relates to the technical field of unmanned aerial vehicles and comprises an outer ring, at least four support rods, an upper clamping plate and a lower clamping plate. The four supporting rods are distributed in a cross shape in the inner circle of the outer ring, one ends of the four supporting rods are fixedly connected with the outer ring, and the other ends are in butt joint in the center of the outer ring. The upper clamping and fixing plate and the lower clamping and fixing plate are located at the upper end and the lower end of the butt joint position of the four supporting rods respectively, clamping and fixing edges parallel to the four supporting rods are arranged on the upper clamping and fixing plate and the lower clamping and fixing plate respectively, the upper clamping and fixing edge and the lower clamping and fixing edge are fixedly connected through bolts, the supporting rod in the middle is fixed, and a stable connection relation is established among the four supporting rods; wherein the outer ring, the supporting rods, the upper clamping plate and the lower clamping plate are all made of carbon fiber materials, energy consumption during flight can be reduced, steering agility is improved, the strength, rigidity, connection strength and stability are extremely high, the structure of the football of the unmanned aerial vehicle is kept stable during fierce collision, and the requirement of a high-strength match is met.

The undercarriage comprises a front undercarriage and two main undercarriages, the front undercarriage is arranged on the front side of the bottom of an aircraft body and comprises a front aircraft wheel and a front buffering supporting column, one end of the front buffering supporting column is connected with the aircraft body and arranged in the aircraft body, and the other end of the front buffering supporting column is connected with the aircraft body. The other end penetrates out of the machine body and is elastically connected with the front machine wheel; the two main undercarriages are arranged on the two opposite sides of the bottom of the aircraft body respectively, each main undercarriage comprises a buffer frame body, the two ends of each buffer frame body are a hinged end and a free end respectively, the hinged ends are hinged to the inner side of the aircraft body, and the free ends are connected with main aircraft wheels. Part of the structure of the front undercarriage is arranged in the aircraft body, and the rocker arm of the main undercarriage is directly hinged with the aircraft body, so that the overall windward area of the undercarriage is reduced, the aerodynamic resistance and the overall weight of the aircraft are reduced, the overall performance and the endurance of the aircraft are improved, and meanwhile, the aircraft is suitable for the aircraft with a shorter aircraft body.

Systems, devices, and methods for an aircraft having a fuselage; a wing extending from both sides of the fuselage; a first pair of motors disposed at a first end of the wing; and a second pair of motors disposed at a second end of the wing; where each motor is angled to provide a component of thrust by a propeller attached thereto that for a desired aircraft movement applies a resulting torque additive to the resulting torque created by rotating the propellers.

The invention discloses a flying backpack for a low-altitude single aircraft, and relates to the field of manned aircrafts. The problems that an existing jet-propelled aircraft garment is large in self weight, inconvenient in pose adjustment and prone to hurting pilots are solved. An oil supply system is installed in a shell, an electrical system is installed on the upper end face of the shell, and a backpack micro turbine engine is installed on the lower portion of the shell. The shell comprises a back plate and a flying back covering part, the back plate is a plate body, a plurality of lightening holes are formed in the upper portion of the back plate, the back plate is made of aluminum alloy, penetrating holes are formed in the periphery of the back plate, and the middle of the safety belt is fixedly installed in the middle of the back plate through bolts and olecranon hooks. A strap of the safety belt penetrates through the penetrating-connecting hole and then is connected with the body of the pilot, the flying back covering part is a shell cover and covers the back plate, and a filiform net is arranged on the lower portion of the flying back covering part. The invention is used in aircrafts.