1502results about "Weight reduction" patented technology

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.

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.

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 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.

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 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.

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 present disclosure relates to a reconfigurable battery system and method of operating the same. An example apparatus includes at least one memory, instructions in the apparatus, and processor circuitry to execute the instructions to determine a state of charge of a battery, determine a closed circuit voltage of the battery, determine a value of a parameter based on a ratio of the state of charge and the closed circuit voltage, and control a switch coupled to the battery based on the value of the parameter, the controlling of the switch to either cause the battery to be coupled to a battery string or cause the battery to be disconnected from the battery string.

The utility model belongs to the technical field of flexible structure design, and particularly relates to a perimeter-variable fuselage frame structure which comprises sliding groove sections, sliding key sections, corrugated beams, height-variable supporting columns and length-variable supporting rods. The corrugated beams are arranged on the outer sides of the sliding key sections and the sliding groove sections and are provided with wave crest and wave trough structures; the sliding key sections and the sliding groove sections are located on the inner sides of the corrugated beams, arranged at intervals and distributed in a staggered mode. The corrugated beams, the sliding key sections and the sliding groove sections are all of arc-shaped structures, and wave crests and wave troughs of the corrugated beams are arranged in the radial direction of the arc-shaped structures; the sliding key section and the sliding groove section are mutually nested to form an arc-shaped supporting sliding way, and the wave trough of the corrugated beam slides on the arc-shaped supporting sliding way; the height-variable supporting column and the length-variable supporting rod are hinged to the two ends of the sliding key section respectively. The shape of the fuselage can be greatly changed, the large-size fuselage can provide a large space to carry more fuel, and the small-size fuselage is beneficial for reducing flight resistance and improving the fuel utilization rate.

The invention relates to the technical field of layered composite fabrics, in particular to a flame-retardant-fiber-based multilayer flame-retardant composite fabric for an aircraft cabin and a preparation method of the flame-retardant-fiber-based multilayer flame-retardant composite fabric. The preparation method comprises the following steps: reacting acrylic acid with DOPO to prepare modified DOPO; reacting the epoxy POSS with the modified DOPO to prepare a composite flame retardant; reacting the hydrophilic grafted polyamide with a composite flame retardant and amino-modified nano-montmorillonite to prepare modified polyamide; carrying out melt spinning and chopping on the modified polyamide to obtain flame-retardant fibers; the flame-retardant fibers are spun into flame-retardant yarns, then the flame-retardant yarns are spun into a fabric, and a surface fabric is obtained; the flame-retardant fibers and the polyester fibers are blended to obtain blended yarns, and then the blended yarns are spun into a fabric to obtain a bottom layer fabric; and compounding the surface layer fabric and the bottom layer fabric to obtain the composite fabric. The prepared composite fabric has excellent flame retardant property and antistatic property, and can be applied to aviation cabins.

A holding system for a wheelchair in an aircraft is proposed, including a plate-shaped base unit having first fastening elements at an underside configured to be connected to complementarily shaped second fastening elements in the form of seat rails on a floor or surrounding structure in the aircraft for fastening the base unit, a front holding unit and a rear holding unit arranged on an upper side of the base unit at a distance from each other and defining therebetween a placement area for a wheelchair. The front holding unit and the rear holding unit are configured to form a releasable form-fit connection with the wheelchair disposed on the placement area so that the wheelchair is braced between the front holding unit and the rear holding unit. The holding system is configured to fix the wheelchair stationarily to the floor during operation of the aircraft.

The invention belongs to the technical field of aircraft structure design, and discloses a fuselage frame preparation process and a fuselage frame. According to the fuselage frame preparation process, preparation of the first woven preformed body and the second woven preformed body is achieved through the weaving operation of the weaving machine, the preparation efficiency and the structural integrity of the first woven preformed body and the second woven preformed body are improved, and the influence of manual operation on the fuselage frame forming quality is avoided. The preparation of the first weaving body and the second weaving body is realized by cutting operation of cutting guide parts arranged on the first weaving auxiliary core mold and the second weaving auxiliary core mold in advance, so that the preparation precision is guaranteed, and the forming quality of the fuselage frame is improved. The invention further provides the machine body frame, the machine body frame is prepared through the machine body frame preparation technology, the manual operation participation degree is reduced in the machine body frame preparation process, the influence of manual operation errors on the forming quality of the machine body frame is avoided, the forming quality and forming efficiency of the machine body frame are improved, and the manual labor intensity is relieved.

A seat unit for a passenger of an aircraft, including a seat and a shell, the seat including a sitting portion and a backrest movable on a base, respectively actuated by a motor-driven mechanism, the unit including a means for detecting the presence of the passenger on the seat, the seat being configured to adopt a takeoff / taxiing / landing position when a presence of the passenger is detected, in which the sitting portion is away from the shell and in which the backrest is reclined in the direction of the shell, and a docking position when a presence of the passenger is not detected, in which the sitting portion is close to the shell, and in which the backrest is substantially parallel to the shell, so as to widen the egress between two seat units.

A seat unit for a passenger of an aircraft, including a seat and a shell, the seat including a sitting portion and a backrest movable on a base, respectively actuated by a motor-driven mechanism, the unit including a means for detecting the presence of the passenger on the seat, the seat being configured to adopt a takeoff / taxiing / landing position when a presence of the passenger is detected, in which the sitting portion is away from the shell and in which the backrest is reclined in the direction of the shell, and a docking position when a presence of the passenger is not detected, in which the sitting portion is close to the shell, and in which the backrest is substantially parallel to the shell, so as to widen the egress between two seat units.

The invention relates to the technical field of composite material forming, in particular to a forming method for improving the performance of a transition area of a heat-proof layer of a special-shaped complex structure and the heat-proof layer, and the forming method comprises the following steps: (1) marking the transition area: marking a boundary line of the transition area on a mold according to the position design requirement of the transition area of a product; (2) marking a layering auxiliary line: marking the layering auxiliary line on the mold; (3) parting and cloth cutting; (4) the prepreg cloth is obliquely laid, specifically, the prepreg cloth is laid on the mold through a staggered lapping process, a plurality of layers of prepreg cloth form a unit, butt joints, located at the transition area, of the interlayer prepreg cloth of each unit are staggered in a stepped mode, and the staggered directions of the butt joints of the prepreg cloth of the adjacent units are the same; the laid and pasted prepreg cloth is compacted; (5) curing; and (6) demolding. According to the method, the butt joints of the prepreg cloth can be prevented from being concentrated and gathered, weak areas are reduced, and therefore the defects of the inclined laying butt joint technology are overcome.

The invention provides a wing panel stringer skin accurate butt joint device and a use method. According to the method, the actual space pose of the skin is determined through a composite skin forming mold, and then the theoretical space pose of the stringer is calculated; space attitude adjustment is conducted on the stringers to be in butt joint through the numerical control attitude adjusting and positioning structure, the stringer accurate positioning device grabs and positions the stringers and accurately moves the stringers in the X direction and the Y direction, finally the stringers are accurately positioned and pressed on the composite material skin, and accurate butt joint of the stringers and the skin of the composite material wing wall plate is achieved.

The invention belongs to the technical field of aviation accessory production, and particularly relates to an automatic welding system and method for aviation oxygen stent production. A plurality of hollowed-out areas formed by the hollowed-out holes are punched and machined through the punching assembly; a plurality of groups of reinforcing plates are clamped and mounted by using the reinforcing plate feeding assembly; punching and processing a U-shaped slot by using the slotting assembly; the arc-shaped limiting plate is bent and formed through the bending assembly; the reinforcing plate welding assembly is used for welding and locking the positions of the reinforcing plate and the main plate; a winder is installed on the lower side of the main board in an abutting mode through a winder feeding mechanical arm, and the winder and the main board are welded and locked through a winder welding assembly; in this way, automatic assembly of the combined aviation oxygen stent is achieved, the combined aviation oxygen stent assembled through welding has high stent strength and durability, the oxygen generator clamping function can be provided, and the oxygen supply pipe can be stored and released conveniently.

The invention provides a composite material oil tank and a manufacturing method thereof, relates to the technical field of composite material oil tanks, and greatly simplifies the mold manufacturing and assembling process through the integrated design of a modular numerical control mold and a three-dimensional digital model. The die purchasing and processing cost is greatly reduced through partition die assembly and reuse of the supporting prosthesis, repeated investment of a special core die is avoided through accurate matching of the silicone rubber soft die and the air bag, and the laying quality of corners and narrow parts is improved. Through cooperation of two times of curing and a stepped lap joint structure, defects of fiber wrinkles, cavities, bridging and the like are effectively eliminated, and fiber continuity and resin bonding strength of a complex curved surface area are ensured. The finally formed composite material oil tank shell not only has excellent mechanical properties of high strength, high rigidity and light weight, but also has excellent sealing performance and durability, and an efficient and reliable solution is provided for large-scale application of a composite material oil tank with a complex structure.

A propulsion system for an aircraft can include an electric power source and an propulsion assembly having a propulsor. An electric power bus can electrically connect the electric power source to the propulsion assembly. The electric power source can be configured to provide electrical power to the electric power bus. An inverter converter controller can be positioned along the electric power bus and can be electrically connected to the electric power source at a location downstream of the electric power source and upstream of the electric propulsion assembly.

The invention relates to the technical field of unmanned aerial vehicles, and particularly discloses an unmanned aerial vehicle capable of landing and buffering, which comprises a fuselage shell, connecting arms are fixedly connected to the opposite angles of the side surfaces of the fuselage shell, and fan blade protection rings are fixedly connected to the ends, away from the fuselage shell, of the connecting arms. A driving motor is fixedly connected to the inner wall of the fan blade protection ring through a support, a driving shaft of the driving motor is fixedly connected with flying fan blades, a mounting frame is fixedly connected to the bottom of the fuselage shell, a damping device is fixedly connected to the bottom of the inner wall of the mounting frame, and a buffering mechanism is fixedly connected to the end, away from the supporting frame, of a fixing rod. According to the unmanned aerial vehicle capable of landing and buffering, the buffering mechanism is arranged, so that the impact force generated when the unmanned aerial vehicle descends can be buffered, and the damping device is arranged, so that the buffered vehicle body can be further damped; damage of parts in the unmanned aerial vehicle caused by violent vibration during descending is avoided.

The invention relates to the technical field of aircraft assembly, and discloses a six-degree-of-freedom attitude adjusting system and method based on a series-parallel connection structure, the attitude adjusting system comprises a moving seat, a mounting frame, a multi-shaft driving assembly and a mounting platform, each driving assembly is provided with a servo motor and a hand wheel dual power source, and is connected with a lead screw module through a bevel gear transmission mechanism; an elastic deflection mechanism is arranged between the power source and the bevel gear and used for achieving auxiliary torque supply, and the multi-shaft driving assembly can achieve adjustment of six postures of front-back, left-right, up-down, pitching, rolling and rotation in the horizontal plane when the large power part of the aircraft is assembled. According to the six-degree-of-freedom attitude adjusting system, through the innovative design of series-parallel mechanism fusion, double-power-source redundant drive, electric cooperative control and the like, high-precision assembly and installation of heavy-weight parts such as aero-engines can be achieved, and high-precision assembly and installation of the heavy-weight parts are high in precision, high in reliability, low in labor intensity and high in adaptability.

The seating systems comprise a seat module and a seat track assembly, a seat track, and an electrical distribution system. The electrical distribution system comprises a plurality of track electrical contacts disposed along the seat track. The seat module comprises a base coupler configured to couple with the seat track and a seat module electrical system having a seat electrical contact. The base coupler and the seat track are configured to position the seat electrical contact in connection with the track electrical contact to which the base coupler is coupled. The methods comprise coupling the base coupler to a seat module, which comprises forming a connection between the seat electrical contact and the respective track electrical contact, and providing electricity to the seat module via the electrical connection.

Disclosed is a middle frame assembly, including a middle plate and a frame disposed around an outer edge of the middle plate, where the middle plate includes a first carbon fiber reinforced resin composite material base body and a first metal plating layer compounded on a surface of the base body. Some embodiments of this application further provide a preparation method for a middle frame assembly and an electronic device in the middle frame assembly. Some embodiments of this application use a carbon fiber reinforced resin composite material as a middle plate base body of the middle frame assembly.

The invention discloses a forming method of a composite material flange and a mold thereof in the technical field of aviation structure connecting piece manufacturing. The forming method comprises the steps of mold pretreatment, layered paving design, twisted wire core material preparation and soft mold auxiliary forming. Fiber arrangement is optimized through gradient paving of the R area and spaced arrangement of the flexible gaskets, and the structural stability is ensured in combination with tooth groove lockrand packaging. A stepped heat-seal pressing curing process is adopted, and synchronous compaction of an R area and a plane area is achieved in cooperation with air bag pressurization. The process comprises core material filling verification, vacuum preloading and temperature control demolding, and finally the high-precision flange component is obtained. According to the method, the problems of poor fiber attachment, non-uniform pressure and stress concentration of the R area of the composite material flange are effectively solved through a synergistic process of progressive pre-compaction and stepped heat-sealing pressing curing, so that continuous arrangement of complex curved surface fibers is realized, the internal stress of an interface is remarkably reduced, and finally, a high-quality flange component with a compact structure and uniform performance is obtained.

The invention relates to the technical field of curved surface forming of fiber composite materials, in particular to a fiber composite material curved surface multi-point flexible gas-assisted hot pressing pressure control method and system.The method comprises the steps that the fiber direction is obtained by obtaining a design mold, and a deformation mismatching area is determined according to the fiber direction and the obtained airbag deformation direction; according to the method, a deformation mismatch area is obtained, a mismatch dynamic control network is constructed to carry out adaptive adjustment on the deformation mismatch area, local pressure compensation is carried out according to obtained adaptive adjustment feedback, and finally, the mismatch dynamic network is controlled by collecting a mold gap and a gap fitting standard. The problem of deformation mismatch caused by elasticity modulus difference in the primary and secondary directions of the fibers is effectively solved, and the direction and position accuracy of the fibers in the forming process is ensured through dynamic adaptive adjustment of a deformation mismatch area, so that the structural integrity and the mechanical property of a formed object are improved, the pressure can be more accurately controlled, and the production efficiency is improved. The requirements of complex mold curved surfaces and fiber layout are met, and the forming precision is improved.