1455results about "Undercarriages" patented technology

A landing system suitable for receiving an unmanned aerial vehicle (UAV) comprises an autonomous ground vehicle (AGV). A landing surface is disposed on the AGV, and the landing system comprises a loading channel suitable for passing an object delivered by the UAV through a first loading channel opening in the landing surface. The object passes within the loading channel through to a second loading channel opening at a bottom aspect of the AGV. In this way, a UAV can land on the landing surface, and the AGV positions the object in line with a target delivery location, where the object is delivered. Aspects of the landing system comprise an electromagnet or vacuum chamber for securing the UAV to the landing surface, thereby enhancing stability of the UAV during movement of the landing system.

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 belongs to the field of aviation technology undercarriage design, and relates to an energy conversion valve used for an undercarriage folding and unfolding system. When the main system works normally, the pilot electromagnetic valve in the energy conversion valve is not electrified, the energy conversion valve is located at the working position of the main system, at the moment, the load subsystem communicates with the main system and isolates the standby system, and the landing gear cabin door / landing gear actuator cylinder is normally retracted and released; when the main system breaks down, the standby system supplies pressure, a pilot electromagnetic valve of the energy conversion valve is powered on, the load subsystem communicates with the standby system, the main system is isolated, and an undercarriage cabin door / undercarriage actuator cylinder is released emergently. The two positions of the valve element are reliably isolated, and a single function fault is not leaked to another system.

The invention provides a main landing gear cabin door mechanism, an aircraft and a folding and unfolding method. The main undercarriage cabin door mechanism comprises a left cabin door and a right cabin door, one end of the left cabin door is provided with a first retractable driving mechanism, one end of the right cabin door is provided with a second retractable driving mechanism, and a small cabin door is hinged to the side, close to the right cabin door, of the left cabin door; the small cabin door and the left cabin door are provided with third folding and unfolding driving mechanisms which can respectively drive the small cabin door and the left cabin door to rotate; in the first position, the left cabin door and the right cabin door are both in a vertical state, and the small cabin door and the left cabin door are folded and overlapped; and at the second position, the left cabin door, the small cabin door and the right cabin door are horizontally positioned on the same plane. The small cabin door can be independently retracted and folded, after the cabin door is opened, the folded small cabin door can reduce the windward area of the cabin door in the put-down state, the pneumatic load borne by the cabin door is reduced, and the design of a driving mechanism for retracting and releasing the cabin door is lighter.

A landing gear assembly for rotorcraft includes a shock absorber, a ground support assembly and an airframe linkage assembly. The shock absorber includes a force sensor to detect a pressure of an incompressible fluid within the shock absorber and a temperature sensor to detect a temperature of the fluid. The ground support assembly is linked to the shock absorber and includes an axle and a ground contact element movable in relation to the axle. The linkage assembly is linked to the shock absorber and / or the ground support assembly and includes a position sensor to detect movement of the axle between compressed and uncompressed conditions of the shock absorber. Various rotorcraft include one, two, three or more landing gear assemblies and an attitude control system. Various rotorcraft methods process sensor signals from one, two, three or more landing gear assemblies at the attitude control system to provide various rotorcraft functions.

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 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.

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.

A method for filling a hydraulic circuit of an electro-hydrostatic system includes connecting first and third shut-off valves to a discharge valve, and connecting a second shut-off valve to a filling valve, while the second and third shut-off valves are closed, opening the filling valve, the discharge valve and the first shut-off valve and operating a vacuum generator, once a vacuum has been established in the hydraulic circuit, opening the second shut-off valve in order to fill the hydraulic circuit with pressurized hydraulic fluid while the vacuum generator is still in operation, once the hydraulic circuit has been filled with fluid, closing the first shut-off valve and then opening the third shut-off valve in order to bypass the vacuum generator and circulate the fluid through the electro-hydrostatic system until the fluid contained in the hydraulic circuit reaches a predetermined pressure, and closing the discharge valve and the filling valve.

According to the modularized traffic tool with multiple take-off modes, through the innovative split modularized design, the form of the traffic tool can be flexibly changed between an automobile and an aircraft, the flight state and the land running state are switched, and more novel passenger and freight application scenes can be expanded through cooperation of the independent carrying cabin modules. By utilizing the tiltable ducted fan mechanism and by virtue of boosting of the autonomous traveling chassis, the composite wing flight module can realize vertical take-off and landing, and wing lift force can be provided for short-distance taxiing take-off and landing to reduce energy consumption, so that the purposes of large voyage and high load are finally realized. The foldable fixed wing, the horizontal tail and the telescopic tail wing structure ensure that the traffic tool is small in size when running and parking on the land, so that the requirements of existing urban and rural roads and traffic safety laws and regulations can be met, and actual popularization is more facilitated.

The invention belongs to an aircraft undercarriage driving technology, and relates to a high-safety undercarriage folding and unfolding system for a medium-sized transport plane. Comprising two sets of hydraulic sources including a first hydraulic source and a second hydraulic source, the two sets of hydraulic sources are backups for each other, the first hydraulic source is used for controlling folding and unfolding of the undercarriage when the undercarriage is normally folded and unfolded, and the second hydraulic source is used for achieving folding and unfolding of the undercarriage when the undercarriage is emergently put down; the two undercarriage control units are used for controlling electromagnetic valves in the system, the normal folding and unfolding function of the undercarriage is achieved, when the undercarriage is put down emergently, a second hydraulic source supplies pressure to control the undercarriage to be put down emergently, and when the hydraulic system fails, gravity is used for driving backup to put down the undercarriage.

The invention discloses an unmanned aerial vehicle-based power distribution network automatic inspection system in the technical field of power distribution network inspection, and the system comprises inspection unmanned aerial vehicles, communication base stations and an inspection master control platform, the inspection unmanned aerial vehicles and the communication base stations are arranged in a matched manner, and the inspection master control platform is equipped with and controls multiple groups of inspection unmanned aerial vehicles and the communication base stations. Mounting matching seat plates are fixedly arranged on the two sides of the inspection unmanned aerial vehicle, and undercarriages are rotationally connected to the lower portions of the mounting matching seat plates. The inspection unmanned aerial vehicle is provided with the communication base station in a matched mode, effective transmission of inspection data and images can be achieved even in a remote mountain environment, in addition, a supporting block and a landing buffering seat are arranged at the landing gear installation position of the inspection unmanned aerial vehicle in a matched mode, and the inspection unmanned aerial vehicle is convenient to use. And the landing and landing buffering seat comprises a buffering main seat body, a buffering telescopic seat body, a buffering matching guide column and a damping spring, when the unmanned aerial vehicle lands and makes contact with the ground, the landing gear can turn over towards the outer side by a certain angle, and effective landing buffering is achieved.

The invention relates to the technical field of hydraulic systems, in particular to a multistage constant-pressure variable adjusting hydraulic system, flight equipment and a method. The multi-stage constant-pressure variable adjusting hydraulic system comprises an elastic assembly, a valve seat assembly, a valve element assembly, a hydraulic pump assembly and a control valve assembly. A constant-pressure variable oil hole is formed in the outer wall of the valve seat assembly, and a first channel, a first containing cavity, a second channel and a second containing cavity which are sequentially communicated are formed in the valve seat assembly. The valve element assembly comprises a first valve element unit and a second valve element unit, and the end face area, away from the first valve element unit, of the second valve element unit is larger than the end face area, facing the second valve element unit, of the first valve element unit. The control valve assembly controls the second containing cavity to be switched and communicated between the oil outlet and the oil return opening of the hydraulic pump, and therefore the opening pressure of the constant-pressure variable oil hole is changed. Therefore, the problem that a single-stage constant-pressure variable form of a pressure regulating mechanism of a conventional flight equipment engine driving pump cannot meet various complex working conditions is solved.

The invention provides an adjustable truss type multi-supporting-column undercarriage, and belongs to the field of aircraft undercarriage.The adjustable truss type multi-supporting-column undercarriage comprises undercarriage supporting columns, lower rocker arms, rotating shafts, side supporting rods, supports, buffers, upper rocker arms and front supporting rods, the undercarriage supporting columns and the lower rocker arms are in pivot joint through the rotating shafts, machine wheels are installed on the lower rocker arms, and the side supporting rods are arranged on the lower rocker arms; the side supporting rod is of a length-adjustable structure, one end of the side supporting rod is hinged to the rotating shaft, the other end of the side supporting rod is connected to the aircraft body structure, the support is installed on the rotating shaft pivoted to the undercarriage supporting column and the lower rocker arm and used for supporting the undercarriage supporting column of the rear undercarriage, one end of the buffer is hinged to the lower rocker arm, and the other end of the buffer is hinged to the upper rocker arm. The upper rocker arm is pivoted to the undercarriage supporting column, the buffer is driven to move by driving the upper rocker arm, then the lower rocker arm can rotate relative to the rotating shaft, and folding and unfolding of the undercarriage are achieved.

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.

An aircraft landing gear having a male part interlocked with a female part along an interlocking direction is provided. The external contact surface of the male part, when it is observed along the interlocking direction, can be an exclusively convex closed surface having external contact surface portions in the form of rounded lobes that form obstacles to the pivoting of the female part with respect to the male part about the interlocking direction. The aircraft landing gear is suitable for use with an aircraft.

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.

The invention discloses an undercarriage and an aircraft, and relates to the technical field of aircraft structural parts, the undercarriage comprises a front undercarriage and two main undercarriages, the front undercarriage is arranged on the front side of an aircraft body, the front undercarriage comprises a front aircraft wheel and a front buffer supporting column, one end of the front buffer supporting column is connected with the aircraft body and arranged in the aircraft body, and the other end of the front buffer 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 aircraft body respectively, each main undercarriage comprises a rocker arm, a main buffer and a main wheel, the two ends of each rocker arm are a hinged end and a free end respectively, the hinged ends are hinged to the inner side of the aircraft body so that the rocker arms can only swing in the course of the aircraft, and the free ends are connected with the main wheels. A 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 to 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, and the overall performance and endurance of the aircraft are improved.

The present invention discloses an actuator cylinder that can be locked and unlocked in a full-stroke cyclic motion and can be controlled arbitrarily. The actuator cylinder is simple and controllable, and can realize full-stroke free motion and active restoration of locking by active failure. The present invention is realized by the following technical scheme: during the reciprocating motion of the piston rod, the mechanical lock provides a locking force through a spring sleeve connected to a spring to realize locking or unlocking; when the hydraulic / pneumatic oil circuit system supplies pressure to the actuator cylinder cavity, the reciprocating motion of the piston rod follows the mechanical lock locking → unlocking → piston rod movement → locking → unlocking... piston rod movement to perform logical cyclic locking and unlocking. The unlocking pipeline is depressurized, and the hydraulic / pneumatic pressure supply system supplies pressure to the unlocking pipeline. The unlocking push rod moves toward the end face of the spring sleeve, pushes the spring sleeve, and pushes it away from the locked position of the mechanical lock, and is in an unlocked state, realizing the free motion of the piston rod in the full stroke of the unlocked actuator cylinder; after the unlocking pipeline is depressurized, the actuator cylinder is restored to the serial logical cyclic stroke of arbitrary locking and unlocking.

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.

The invention provides an aircraft landing gear supporting rod lock and belongs to the technical field of aircraft landing gear, the aircraft landing gear supporting rod lock comprises a front lock supporting rod, a rear lock supporting rod, an unlocking rocker arm, an unlocking actuating cylinder, a locking damping actuating cylinder and a locking spring, one end of the front lock supporting rod is connected with the landing gear supporting rod, the other end of the front lock supporting rod is hinged to the rear lock supporting rod, and the rear lock supporting rod is hinged to the rear lock supporting rod. The other end of the rear lock supporting rod is hinged to an undercarriage supporting column and hinged to the locking damping actuating cylinder, the unlocking rocker arm is arranged at the hinged position of the front lock supporting rod and the rear lock supporting rod, one end of the unlocking actuating cylinder is connected with the front lock supporting rod, and the other end of the unlocking actuating cylinder is connected to the unlocking rocker arm. One end of the locking damping actuating cylinder is hinged to the rear lock supporting rod, the other end of the locking damping actuating cylinder is connected with the locking spring, the other end of the locking spring is connected with the front lock supporting rod, and the connecting end of the locking spring and the locking damping actuating cylinder is arranged on the landing gear supporting column.

A wheel (1) mounted for rotation on an axle and comprising: a rim (2) connected by a flange (3) to a hub (4) received for pivoting on the axle, and defining with the hub an annular space; a radial flux electric motor (10) comprising a first stator (11) and a first rotor (12) arranged to generate a driving torque for rotation of the wheel; a magnetic brake (20) comprising a second stator (21) and a second rotor (22) arranged to generate a braking torque for the wheel, the electric motor and the magnetic brake extending into each other within the annular space, one of the first and second rotors being adjacent to the rim and the other of the first and second rotors being adjacent to the hub, the first and second stators extending between the first and second rotors. (Figure from the abbreviation: Fig. 2)

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.