1049results about "Undercarriages" patented technology

The invention belongs to the field of electromechanical control, relates to an electric pump work control technology, and discloses an electric pump work comprehensive control method and system.The work comprehensive control method comprises the steps that onboard state signals, flight control system state signals and hydraulic pressure of each sub hydraulic system in a hydraulic system are collected in real time; starting and stopping of an electric pump connected to the hydraulic sub-system are controlled through a first electric pump control logic and an anti-error working mechanism according to the onboard state signal; and according to the flight control system state signal, the hydraulic pressure and other electric pump control instructions, the electric pump connected to each sub hydraulic system is automatically controlled through second electric pump control logic, an anti-error working mechanism and a time-staggered working mechanism. The working accuracy of the electric pump can be ensured, meanwhile, system damage caused by simultaneous starting of the electric pump is avoided, and the system safety is improved.

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.

Heat exchangers for cryogenic fuel distribution systems and related methods are disclosed herein. An example aircraft disclosed herein includes a cryogenic fuel system including a cryogenic fuel, a hydraulic system including a hydraulic fluid, a gas turbine engine fluidly coupled to the cryogenic fuel system, and a heat exchanger fluidly coupled to the cryogenic fuel system, the heat exchanger is used to exchange heat between the cryogenic fuel and at least one of (1) hydraulic fluid or (2) working fluid thermally coupled to the hydraulic system.

The invention belongs to an electric pump work control technology in the aviation field, and discloses an anti-error work control method and system for a multi-redundancy automatic electric pump, and the method comprises the steps: collecting an onboard state signal, a flight control system state signal and two paths of hydraulic pressure of each sub-hydraulic system in a hydraulic system in real time; according to the onboard state signal, a hydraulic pressure logic judgment instruction is output through the first control logic or an electric pump mistake proofing mechanism is started, and an electric pump connected to the sub hydraulic system is controlled to be closed through the electric pump mistake proofing mechanism; a flight control system state signal and two-way hydraulic pressure are obtained according to the pressure logic judgment instruction, and an electric pump connected to each hydraulic sub-system is automatically controlled through second control logic and an electric pump mistake proofing mechanism. Through the design of the first control logic, the second control logic and the electric pump mistake proofing mechanism, the working accuracy of the multi-redundancy automatic electric pump can be ensured, and the system safety is improved.

The invention belongs to the field of airborne electromechanics, and discloses a hydraulic mechanism heterogeneous control method and system.The method comprises the steps that hydraulic systems are divided into two groups, and on-board state signals, flight control system state signals and first hydraulic pressure and second hydraulic pressure of sub hydraulic systems are collected; giving an on-board state threshold value, a low-pressure threshold value, a hydraulic safety threshold value and a sub-hydraulic system relation table; starting and stopping automatic control is conducted on the electric pumps in the first group through the established first control logic architecture according to the data; and the electric pumps in the second group are automatically controlled to start and stop by adopting the constructed second control logic architecture through the data. The system comprises a grouping module, a control logic construction module and two control units, wherein each control unit comprises an acquisition module, a data scheduling module, a staggered starting module, an anti-overvoltage protection module and an anti-error working module. All the electric pumps are controlled through two heterogeneous logic architectures, and the control precision and safety are improved.

The invention provides a broadband sound absorption structure suitable for undercarriage cabin door noise reduction and a design method, and belongs to the field of aircraft noise reduction. The sound absorption structure comprises a plurality of perforated plate type Helmholtz cavity unit cells and a plurality of embedded pipe type Helmholtz cavity unit cells. In the perforated plate type unit cells, one or more first through holes are formed in the perforated plate. And in the embedded pipe type unit cells, the embedded pipes penetrate through the perforated holes in the perforated plates and extend into the cavities. Cavities of the two single cells face the same direction, and the perforated plate forms a curved surface conformal with an undercarriage cabin door. The effective length of the neck of the perforated plate is increased by using the embedded pipe, the low-frequency sound absorption performance is improved, the sound absorption performance and the sound absorption bandwidth of the middle-frequency band are improved by using the perforated plate type Helmholtz cavity, meanwhile, the curved surface conformal design is highly matched with the special-shaped cabin door structure, and the sound absorption performance is improved. The low-frequency and medium-frequency sound absorption performance can be guaranteed, and meanwhile the light and thin requirements of an aircraft can be met; and a pure passive mechanical structure is adopted, so that the control complexity and stability risks of an active noise reduction system can be avoided.

A locking system for doors of a deployable aircraft landing gear system includes a spring-biased swing arm configured to be pivotally attached to a body of an aircraft and including a roller at a distal end of the swing arm. The swing arm is configured to engage with a portion of a first landing gear bay door to prevent the first landing gear bay door from opening.

Disclosed here is a landing gear support fixture for an aircraft under test having a landing gear. The disclosed landing gear support fixture includes: an axle attachment assembly that is compatibly shaped, sized, and configured to mate with and attach to at least one axle of the landing gear of the aircraft under test; a support platform to support the axle attachment assembly, wherein the support platform is in a fixed position relative to a stationary reference foundation; and a force sensor couplable between the axle attachment assembly and the support platform to provide sensor output associated with load imparted by the landing gear of the aircraft under test.

The invention relates to the technical field of unmanned aerial vehicle loads, in particular to an unmanned aerial vehicle load structure which comprises a concave supporting seat, supporting legs are symmetrically and fixedly mounted on the left side and the right side of the concave supporting seat, a clamping driving mechanism is mounted on the inner side of the concave supporting seat, and an opening of the concave supporting seat is designed downwards. T-shaped connecting plates are symmetrically mounted on the left side and the right side of the clamping driving mechanism, and clamping parts are detachably mounted at the lower ends of the two T-shaped connecting plates. The unmanned aerial vehicle has the beneficial effects that when the load body needs to be clamped, the unmanned aerial vehicle drives the load structure to sleeve the load body, and meanwhile, the clamping parts are distributed on the left side and the right side of the load body; and in the downward landing process of the unmanned aerial vehicle, the lower end of the anti-falling mechanism rotates outwards to be opened, and meanwhile, the anti-falling mechanism drives the transmission mechanism to move close to the first locking mechanism, so that the first locking mechanism loosens locking of the clamping driving mechanism.

A rotorcraft with at least one non-retractable wheel-type landing gear and a fuselage comprising a lower shell, an inner floor, and at least one lower fuselage compartment arranged between the inner floor and the lower shell, wherein the at least one non-retractable wheel-type landing gear is rigidly attached in the at least one lower fuselage compartment and comprises: at least one wheel; a shock absorber connected to the at least one wheel and comprising a shock absorber tube and a shock absorber rod telescopically mounted to the shock absorber tube; at least two mounting arms extending laterally from the shock absorber tube; and a mounting pin extending from the shock absorber tube toward a floor fitting provided at the inner floor, wherein the mounting pin is connected to the floor fitting.

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 and methods for mechanically rotating an aircraft about its center-of-gravity (CG) are disclosed. The system can enable the rear, or main, landing gear to squat, while the nose landing gear raises to generate a positive angle of attack for the aircraft for takeoff or landing. The system can also enable the nose gear and main gear to return to a relatively level fuselage attitude for ground operations. The system can include one or more hydraulically linked hydraulic cylinders to control the overall height of the nose gear and the main gear. Because the hydraulic cylinders are linked, a change on the length of the nose cylinder generates a proportional, and opposite, change in the length of the main cylinder, and vice-versa. A method and control system for monitoring and controlling the relative positions of the nose gear and main gear is also disclosed.

A rotorcraft comprising a fuselage forming at least one lower lateral fuselage compartment with at least a first fitting, a second fitting, and a third fitting; and at least one non-retractable wheel-type landing gear that comprises: at least one wheel; a shock absorber with a shock absorber tube having an open end and an axially opposed closed end, and a shock absorber rod telescopically mounted to the shock absorber tube via the open end; a first fixation provided close to the open end on the shock absorber tube and connected to the first fitting; a second fixation provided close to the closed end on the shock absorber tube and connected to the second fitting; and a third fixation provided close to the closed end on the shock absorber tube and connected to the third fitting.

An aircraft landing gear shock absorber strut including an outer cylinder having a bore defining an opening; an inner cylinder having a first end region movably coupled within the bore and a second end region which projects out of the opening. The inner cylinder being arranged to move along a longitudinal axis of the bore between a first condition and a second condition and is biased by a spring force to assume the second condition. A mechanical outstop is arranged to engage an abutment surface of the inner cylinder to limit extension of the shock absorber strut; and a rotatable member drives the out stop axially to move the inner cylinder from the second condition towards the first condition. The rotatable member is arranged such that the spring force cannot rotate the rotatable member to allow the shock absorber strut to extend.

The present invention relates to a fibrous preform of a core portion of a fibrous reinforcing material of a composite material part, which has an elongated shape along the longitudinal direction X and is formed by first yarns C1 to C16 extending along the longitudinal direction and second yarns t1 to t8 crossing the first yarns. In the fibrous preform, the preform includes two longitudinal ends intended for articulation joining with other parts and an intermediate region located between the longitudinal ends. Each longitudinal end has a thickness greater than the thickness of the intermediate region. The intermediate region includes first non-woven yarns C5 to C12. Each longitudinal end includes a reinforcing segment 102c including a three-dimensional weave of a first non-woven yarn and a second yarn of the reinforcing segment.

The invention relates to the field of valve sealing structures, and particularly discloses an aircraft pneumatic electromagnetic valve sealing structure and a machining method thereof.The aircraft pneumatic electromagnetic valve sealing structure comprises a sealing structure body, a side wall groove is formed in one end of the sealing structure body, and the sealing structure body comprises any one of a first sealing structure and a second sealing structure; according to the first sealing structure, a plurality of annular bionic grooves which are arranged up and down are formed in the inner circumferential face of the side wall groove, and a plurality of concentrically-distributed annular bionic grooves are formed in the bottom end face of the side wall groove. According to the second sealing structure, a plurality of linear bionic grooves which are distributed in a circumferential array mode are formed in the inner circumferential face of the side wall groove, and a plurality of divergently-distributed linear bionic grooves are formed in the bottom end face of the side wall groove. According to the method, the micro-nano bionic sealing structure is machined on the electromagnetic valve sealing surface through the femtosecond laser micro-nano machining technology, the roughness consistency and the bonding strength of the rubber bonding surface of the pneumatic sealing element are improved, the sealing performance of the pneumatic element is guaranteed, and the reliability of an airplane emergency control system is guaranteed.

The present invention discloses a processing method for an emergency breakaway pin for an aircraft main landing gear. The pin is hollow and cylindrical, with two breakaway slots provided in the inner bore corresponding to two locations bearing shear force. A thickened section is provided between the two breakaway slots, and the inner bore diameter corresponding to the thickened section is smaller than that at other locations. The bottom of the breakaway slot is connected to the straight stepped surfaces on both sides by a rounded transition. The emergency breakaway pin of this solution has a breakaway slot design that not only controls precise breakaway but also limits the breakaway area. The rounded transition between the bottom of the breakaway slot and the straight stepped surfaces on both sides improves fatigue strength, and the thickened section between the two breakaway slots improves static strength. The measured breakaway value can be controlled within a range of ±4% of the set threshold, achieving precise breakaway.

The invention relates to a hovercar, a vehicle and an aircraft. The hovercar comprises the vehicle and the aircraft. The vehicle comprises a vehicle body and the separation and combination device, the vehicle body is fixedly provided with a land-running cabin, and the separation and combination device is arranged on the vehicle body; the aircraft comprises an aircraft body and a plurality of foldable rotors, and the aircraft body is fixedly provided with a flight cabin; the plurality of foldable rotors are arranged on the aircraft body; the aircraft can vertically take off and land on the ground, and the separation and combination device is used for pulling the aircraft to be in coupling connection with the vehicle when the aircraft is in a state of stopping on the ground. In the embodiment, the aircraft does not need to be precisely positioned with the bearing platform on the vehicle body and can rapidly and vertically stop and descend on the ground, then the separation and combination device of the vehicle is connected with the aircraft and drags the aircraft to the bearing platform, and the control difficulty of coupling of the aircraft and the vehicle is relatively low.

A landing gear includes a supporting structure, and a damping cylinder connected to a lever which is hinged to the supporting structure about a hinge axis. The lever is movable between a towards end position, in which the wheel is moved towards the hinge axis, and an away end position, in which the wheel is moved away from the hinge axis. The towards and away end positions are on either side of an intermediate position, in which the lever and the cylinder are aligned. The end away position is defined by a stop secured to the supporting structure and against which the lever in the away position is applied to take up the landing forces.

A landing gear assembly is provided. The landing gear assembly includes a shock strut. The shock strut includes a shock strut cylinder and a shock strut piston slidably disposed within the shock strut cylinder. The landing gear assembly further includes a temperature control unit assembly disposed within the shock strut cylinder. The temperature control unit assembly is configured to, responsive to a temperature within the shock strut cylinder falling below a predetermined temperature, circulate heated fluid into a chamber of the shock strut cylinder to heat gas within the shock strut cylinder.

An aircraft notification system, the aircraft notification system having a processor and a memory communicatively connected to the processor, the memory containing instructions configuring the processor to receive aircraft angle data from a blended wing body aircraft (BWB) having a main body and a wing with no clear demarcation between the wings and the main body along a leading edge of the aircraft, determine an aircraft orientation as a function of the aircraft angle data, wherein the aircraft orientation comprises at least a pitch angle of the BWB while the BWB is in contact with a ground surface, transmit the aircraft orientation to an aircraft display system and visually display the aircraft orientation through the aircraft display system.

An aircraft landing gear has a suspension mechanism which has a four-bar linkage that comprises four links rigidly extending between a corresponding four pivots. A lower link of the four-bar linkage supports front and rear wheels, and is movable relative to an upper link between lengthened and shortened positions. A shock absorber urges the four-bar linkage towards the lengthened position. A length of the upper link of the four-bar linkage is adjustable.

The present disclosure relates to a strut and a landing gear including the strut, and more particularly, to an disclosure in which a strut is formed of a curved pipe formed of a flexible body to reduce a weight of a component, and at the same time, an impact load of the strut and the landing gear is alleviated through a design of the physical properties and shape of the strut.

A hydraulic brake system for activating the wheel brakes of an aircraft landing gear assembly includes a hydraulic fluid supply line and an associated return line. There is a valve which separates a source of hydraulic fluid from a portion of the return line and an associated leak path associated with the valve that in use maintains the pressure and / or volume of hydraulic fluid in the portion of the return line above a minimum acceptable level. The leak path may be provided by a hole drilled in a part (e.g. the poppet) of a non-return check valve.

An aircraft including a blended wing body aircraft (BWB) having a main body and wings with no clear demarcation between the wings and the main body along a leading edge of the aircraft, a rear spar located within the main body, one or more gear wells located on a lower surface of the rear spar, a landing gear system having at least a main gear the at least a main gear situated in one of the one or more gear wells, and one or more propulsors located directly above the one or more gear wells, wherein at least one propulsor is located over the at least one gear well of the one or more gear wells.

The invention relates to the technical field of aerospace, and discloses a novel aircraft auxiliary landing stabilizing device and a control system thereof.The novel aircraft auxiliary landing stabilizing device comprises an aircraft body, an undercarriage is rotatably connected to the interior of the aircraft body, a fixing plate is arranged in the undercarriage, and a control center is fixedly connected to the outer wall of one side of the fixing plate; a fixing assembly is fixedly connected to the outer wall of the other side of the fixing plate, an adjusting assembly is fixed in the fixing assembly, an air inlet assembly is fixedly connected to the lower surface of the adjusting assembly, an airflow rotating assembly is fixedly connected to the interior of the air inlet assembly, and a transmission assembly is arranged in the fixing shell. And the outer wall of the transmission assembly is fixedly connected with a machine wheel. In the invention, the electric push rod is controlled by the control center, the rotation angle of the guide plate is finally controlled, and the airplane wheels are finally driven to rotate, so that the friction force at the landing moment can be obviously reduced, and the tire wear and overload risk can be reduced.

The invention discloses an aircraft nose landing gear structure and a folding and unfolding method, the aircraft nose landing gear structure comprises a fuselage connecting assembly, a strut assembly, an oil injection valve assembly, a shimmy damping assembly, a wheel assembly and a folding and unfolding control assembly, the fuselage connecting assembly is arranged at the upper section of the strut assembly, the top end of the strut assembly is connected with a fuselage, and the lower end of the strut assembly is connected with the wheel assembly; the supporting column assembly is arranged on the machine body and fixedly connected with the machine body through the machine body connecting assembly, the oil injection valve assembly is arranged on the supporting column assembly and higher than the horizontal position of the machine body connecting assembly, the upper portion of the anti-swing assembly is connected with the supporting column assembly, the lower portion of the anti-swing assembly is connected with the wheel assembly, and the two ends of the folding and unfolding control assembly are connected with the supporting column assembly and the machine body respectively. The first semicircular piece and the second semicircular piece are fixed through a tightening bolt, the friction plate is made of a carbon-carbon composite material, and vibration transmitted to a machine body is reduced. The split structure facilitates replacement of the friction plates, the whole shimmy damper does not need to be disassembled, and maintenance time is shortened.

An embodiment aircraft boom (66, 76) includes a boom member (34, 42), a vertical stabilizer (40, 50), a rear landing gear (114, 116), and a front landing gear (110, 112). The boom member (34, 42) may extend in a first direction and have a front internal structural member (154) disposed therein, where a portion of the front internal structural member (154) is a first attachment point for a first wing. The vertical stabilizer (40, 50) may extend from the boom member (34, 42), where the vertical stabilizer (40, 50) has a rear internal structural member (120) disposed therein and extends from the boom member (34, 42) and through an interior of the vertical stabilizer (40, 50), and where a portion of the rear internal structural member (120) is a second attachment point for a second wing. The rear landing gear (114, 116) may be coupled to the rear internal structural member (120). The front landing gear (110, 112) may be coupled to the front internal structural member (154) of the boom member (34, 42).

The disclosure proposes a locking device for movable components, in particular aircraft components such as an aircraft landing gear, which comprises a hook rotatable about an axis of rotation for receiving a pin portion fixedly arranged on the landing gear, wherein the hook has a U-shaped, C-shaped or V-shaped opening, which is designed receive the pin portion in a locked state in a first rotational position and to release the pin portion in an unlocked state in a second rotational position. The disclosure is characterised in that at least one of the two legs of the U-shaped, C-shaped or V-shaped opening has a non-constant contour in the width direction.