571results about "Undercarriages" patented technology

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.

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 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 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.

A flying robot includes a body portion, a propulsion portion including a plurality of propulsion units configured to generate propulsion force by driving rotor blades, the plurality of propulsion units being provided at the body portion, a plurality of leg portions configured to support the body portion, each leg portion of the plurality of leg portions including at least one joint and being configured to be able to change a posture of the leg portion, and a controller configured to control the plurality of leg portions when landing on a landing surface from a flying state, and the controller controls part or all of at least one leg portion among the plurality of leg portions to adjust a tilt of the body portion from when the at least one leg portion comes into contact with the landing surface until when landing on the landing surface is completed.

A seaplane float frame apparatus with extra fuel storage, and a method of installing, comprising: 1) a middle section housing at least one auxiliary fuel storage tank; and 2) two opposing side panels affixed to the middle section, extending outward from about 45-60-degrees from the parallel, each of the two opposing side panels comprising a bottom flat edge fit-able to a float (e.g., carbon composite). The middle section further comprises a fuel pump and a fuel line attached to the auxiliary fuel storage tank(s) and to a main fuel tank within an airplane fuselage-engine. Electrical components enable the pilot, or automatedly as needed, to pump fuel from the frame's tanks. The two opposing side panels comprise triangular shaped cutouts to reduce the float frame's weight and drag, and a built-in steps, a bar, and non-skid coating on the floats to assist in climbing safely into / out of the cockpit.

An aircraft system is disclosed including a hydraulic actuator, an inlet line for supplying hydraulic fluid to the hydraulic actuator, an outlet line for removing hydraulic fluid from the hydraulic actuator, a first one-way valve configured to allow fluid flow along the inlet line in a direction toward the hydraulic actuator, and a second one-way valve configured to allow fluid flow along the outlet line in a direction away from the hydraulic actuator.

Systems for activating and deactivating an emergency landing gear system in an aircraft include a rotary toggle mechanism that actuates a switch. When the switch is actuated, the emergency landing gear system is activated and the operation of one or more other aircraft systems may be modified accordingly. The rotary toggle mechanism may be placed in three different states: a ready state, an active state, and a reset state. In the ready state, the switch is not actuated and the emergency landing gear system is not activated. In the active state the switch is actuated and the emergency landing gear system is activated. While the mechanism is in the active state the mechanism cannot be returned to the ready state without first passing through the reset state.

The application belongs to the field of electromechanical control, and relates to electric pump working control technology, and discloses an electric pump working comprehensive control method and system.The working comprehensive control method comprises the following steps: collecting machine state signals, flight control system state signals and hydraulic pressure of each sub-hydraulic system in a hydraulic system in real time; controlling the opening and closing of the electric pump connected to the sub-hydraulic system according to the machine state signals through a first electric pump control logic and an error prevention working mechanism; and automatically controlling the electric pump connected to each sub-hydraulic system according to the flight control system state signals, the hydraulic pressure and the rest of the electric pump control instructions through a second electric pump control logic, the error prevention working mechanism and a staggered working mechanism.The application can ensure the accuracy of the electric pump working, avoid system damage caused by simultaneous starting of the electric pump, and improve system safety.

The application provides an aircraft, comprising a fuselage, a bumper, a connecting mechanism and a landing gear, the cylinder body of the bumper is assembled to the fuselage; the connecting mechanism is hinged to the movable rod of the bumper and the fuselage; and the landing gear is foldably connected to the connecting mechanism. In this way, when the landing gear vibrates, the landing gear can transmit the vibration to the connecting mechanism to drive the connecting mechanism to rotate relative to the fuselage, and then drive the movable rod to slide in the cylinder body to achieve the effect of energy absorption and vibration reduction through the bumper. In addition, compared with the technical solution in the prior art that the bumper is synchronously folded with the landing gear, the aircraft provided by the embodiment of the application does not need to consider the influence of the weight of the bumper on the folding of the landing gear, the weight of the folding part is reduced, and no folding space needs to be reserved for the bumper after the landing gear is folded, which helps to reduce the volume of the aircraft.

The application relates to the field of aircraft control, and discloses a landing gear actuating cylinder control device and a landing gear, which comprise a controller, a main control circuit and at least one standby control circuit, wherein the main control circuit comprises a detection circuit; the detection circuit is arranged at the landing gear actuating cylinder to generate a detection signal; the controller is used for generating an action instruction according to state information of an airplane and sending the action instruction to the main control circuit, so that the main control circuit controls the landing gear actuating cylinder according to the action instruction; the controller is also connected with the detection unit to acquire the detection signal, and when the detection signal does not satisfy a preset condition, the controller controls the standby control circuit to work and sends the action instruction to the standby control circuit. When the main control circuit fails, the standby control circuit is started to work, so that the stability of the landing gear actuating cylinder control system and the safety of the airplane are improved on the basis of preventing multiple controller interference and motor power disputes, so as to prevent the airplane from being unable to take off or land safely due to the failure of the actuating cylinder.

An aircraft hydraulic system 100 comprises common hydraulic pressure sources 11 and 12 and pressure reduction valves 21 and 22. The common hydraulic pressure sources 11 and 12 supply a shared hydraulic pressure to a low pressure operation unit 103 that is disposed in an aircraft 101 and that operates by a relatively low operating hydraulic pressure and to a high pressure operation unit 102 that is disposed separately from the low pressure operation unit 103 in the aircraft 101 and that operates by a relatively high operating hydraulic pressure. The pressure reduction valves 21 and 22 reduce the hydraulic pressure that is supplied to the low pressure operation unit 103.

This invention discloses a landing gear and a racing drone, belonging to the field of racing drones. It includes two mounting beams positioned front and rear at the bottom of the drone's fuselage; two support rods symmetrically mounted on one of the mounting beams; two telescopic components symmetrically mounted on the other mounting beam; and two base plates symmetrically mounted at the bottom ends of the support rods and telescopic components. The support rods and telescopic components are arranged in pairs, with each base plate simultaneously universally connected to the bottom end of one pair of support rods and telescopic components. The telescopic components extend and retract to adjust their axial length, changing the tilt angle of the base plate relative to the plane of the two mounting beams. This invention adjusts the forward tilt angle of the racing drone by extending and retracting the length of the rear support legs, enabling optimal launch at the forward tilt angle under different load conditions. After the racing drone leaves the ground, as the load on the landing gear is released, the landing gear automatically returns to a state of uniform front-to-back height in the air, ensuring a safe landing of the racing drone in a horizontal attitude.

The invention discloses a drive control device and an aircraft landing gear extension and retraction system.The drive control device comprises an outer cylinder, a first pipe joint and a second pipe joint, the first pipe joint and the second pipe joint are connected with the outer cylinder, a first cavity is formed in the outer cylinder, and the first pipe joint and the second pipe joint are arranged in the length direction of the outer cylinder; a piston rod is arranged in the first cavity, one end of the piston rod is slidably connected with the inner wall of the first cavity, and the outer end of the piston rod and the second pipe connector are located on the same side of the first pipe connector. A first channel is arranged in the first pipe connector, a first connector is arranged on the first channel, and the first channel is communicated with the first cavity. A pressure relief assembly is arranged on the first pipe joint, and one end of the pressure relief assembly is communicated with the first channel; a second channel is arranged in the second pipe joint, and a second interface and a third interface are arranged on the second channel; a moving assembly and a first elastic element are arranged in the second channel. The driving control device solves the technical problems that in the prior art, an undercarriage folding and unfolding system is large in weight and complex.

The application discloses an online configuration system and method for landing gear control redundancy, wherein a comprehensive control system publishes a redundancy scheme according to system working states and application scene requirements; a redundancy monitoring module analyzes redundancy configuration and scheduling strategy information according to the received redundancy scheme, realizes online redundancy configuration and recombination of a control module, carries out redundancy management on the control module according to the redundancy strategy, monitors the health state of the control module and reports the comprehensive control system, and outputs an effective driving signal to an element driving module; a landing gear control signal is provided for the control module by a landing gear control device; and the control module outputs a driving signal to the element driving module after logical calculation according to the received landing gear control signal. The application realizes online on-demand configuration of control redundancy and dynamic management of redundancy scheduling strategy.

This invention provides a technology that allows the main rotor and shaft to be tilted relative to the landing plane in a rotary-wing aircraft using a simple configuration. [Solution] The rotary-wing aircraft comprises a fuselage having a base located at the bottom, a shaft connected to the fuselage so as to rotate around a rotation axis that is inclined at a predetermined first angle with respect to the base, a main rotor connected to the shaft, and a landing gear that contacts the object to be landed on when the rotary-wing aircraft lands. The landing gear includes a pair of skid tubes that define a landing plane, and a pair of cross tubes connecting the pair of skid tubes, each having a first portion extending along the landing plane and a pair of second portions connecting the first portion and the pair of skid tubes. The base is inclined at a predetermined second angle with respect to the first portion and connected directly or indirectly to the first portion so as to the rotation axis inclined with respect to the landing plane.

A sled landing gear and a method for using same. The sled landing gear comprises a main landing gear (100), a front steel cable (200), a rear steel cable (300), a retractable steel cable (400), a main gear sled (500), and a retraction / extension device. The main gear sled (500) is a double-board structure with an opening. An end of the front steel cable (200) is connected to the main landing gear (100). Another end of the front steel cable (200) is connected to a front end of the main gear sled (500). An end of the rear steel cable (300) is connected to the main landing gear (100). Another end of the rear steel cable (300) is connected to a rear end of the main gear sled (500). An end of the retractable steel cable (400) is connected to a middle section of the rear steel cable (300). Another end of the retractable steel cable (400) is connected to the retraction / extension device. The retraction / extension device controls retraction and extension of the sled via the retractable steel cable.

The invention discloses a multi-rotor aircraft with lift-enhanced wings, which comprises rotor power devices, a fuselage for mounting airborne equipment and connecting the rotor power devices, and a pair of independent and controllable lift-enhanced wings, the high-lift wings are mounted on the fuselage; when the multi-rotor aircraft with the high-lift wings flies, a flight control algorithm controls the tilting angles of the high-lift wings according to the attitude and the flight speed of the aircraft, so that the high-lift wings are in a proper attack angle state, and proper lift force can be provided for the aircraft. The flying speed can be increased, the cruising ability is improved, and transportation and storage are convenient.

An aircraft landing gear assembly 6 comprising; a noise-inducing component 8 which, in use, with the landing gear assembly deployed, has a front region and a back region separated by a pair of side re

An unmanned helicopter platform includes a fuselage, a tail coupled with the fuselage, a payload rail coupled with and extending along the fuselage and a main rotor assembly coupled with the fuselage. The tail includes a tail rotor and a tail rotor motor. The main rotor assembly includes a main rotor having an axis of rotation and a main rotor motor. The payload rail allows mechanical connection of payloads to the fuselage and positioning of the payloads such that a center of gravity of the payloads is alignable with the axis of rotation. A system for controlling the unmanned helicopter includes a processor and a memory for providing instructions to the processor. The processor can receive a task, dynamically determine a route for the task and autonomously perform the task including flying along at least part of the route. The route is based on the task, geography and terrain.

The utility model discloses a worm undercarriage which comprises an undercarriage shell, a driving device, a pushing device and a quick-release foot stool, a motor mounting part and a worm mounting part are sequentially arranged in the undercarriage shell from left to right, the motor mounting part is communicated with the worm mounting part, a fixing pin is arranged in the worm mounting part, the driving device is arranged in the motor mounting part, and the pushing device is arranged in the quick-release foot stool. The pushing device is arranged in the worm mounting part, one end of the driving device is connected with the upper end of the pushing device, the lower end of the pushing device is fixedly connected with the quick-release foot stool, the pushing device is divided into a crank arm part and a fixing part from top to bottom, the crank arm part and the fixing part are rotationally mounted in the worm mounting part through fixing pins, and the driving device drives the crank arm part to move left and right. And the fixing part drives the quick-release foot stool to upwarp or fall. The landing requirement of the unmanned aerial vehicle is met; the pushing device can be conveniently driven by the driving device to fold and unfold, the worm undercarriage is simply, conveniently and rapidly folded and unfolded, and the folding and unfolding efficiency of the unmanned aerial vehicle undercarriage is improved.

A retractable landing gear assembly for an aircraft has a main strut (210) arranged to pivot between a retracted position and an extended position, and a sidestay (220) having an unfolded position and a folded position. A telescopic actuator (250) is coupled between the main strut (210) and the sidestay (220), so that when the actuator (250) has a first length the sidestay (220) locks the main strut (210) in the extended position, and when the actuator (250) has a second length the sidestay (220) is in the folded position and the main strut (210) is in the retracted position.

The utility model relates to the technical field of undercarriages, in particular to a modular undercarriage with a strong stability mechanism, which comprises a bearing frame, two adjusting plates are fixedly connected to one side of the bearing frame and are symmetrically arranged, an auxiliary bearing frame is fixedly mounted at one end of the bearing frame, and a plurality of auxiliary bearing frames are fixedly mounted at the other end of the bearing frame. The surface of the adjusting plate is rotatably connected with a mounting plate, the surface of the mounting plate is fixedly connected with a first shock absorption supporting column, the first shock absorption supporting column is hollow, and the inner wall of the first shock absorption supporting column is slidably connected with a vertical rod. Kinetic energy generated when the aircraft lands on the runway is converted into electric energy to be transmitted to the super-capacitor module to be stored, so that energy which is originally consumed and wasted in the braking process is recycled, cyclic utilization of the energy is achieved, the overall energy utilization efficiency of the aircraft is improved, dependence on external energy is reduced to a certain degree, and the operation cost is reduced.

An aircraft landing gear (100,fig.1) comprising a linear actuator 110 attached to a portion of the aircraft landing gear and operable to move the portion of the aircraft landing gear. The landing gear

The application belongs to the technical field of aircrafts, and specifically discloses a vertical take-off and landing aircraft landing system, a control method and an aircraft. The vertical take-off and landing aircraft landing system comprises a landing mechanism, an energy absorption mechanism and a braking mechanism. The landing mechanism comprises a frame, a support frame and a wheel arranged at the lower end of the support frame. The energy absorption mechanism is arranged on the frame and is used for absorbing kinetic energy when the support frame is turned outwards. The braking mechanism is used for adjusting the braking force of the wheel to adjust the turning-out amplitude of the support frame. When the aircraft lands, the braking mechanism actively adjusts the friction coefficient between the landing gear and the ground to adjust the size of the friction work of the landing gear and the work of the energy absorption system of the landing gear, thereby reducing the vertical acceleration of the fuselage of the aircraft, further reducing the fuselage load, reducing the risk of plastic deformation and damage of the fuselage and airborne equipment, reducing the load transmitted to the passengers, improving the safety and reliability of the whole machine, ensuring the safe landing of the aircraft, and reducing the risk of damage of the aircraft due to hard landing.

The invention provides a method and a device for calculating the speed of lifting a front wheel during takeoff in consideration of dynamic influence, and belongs to the technical field of aircraft pneumatic and operation stability design. The method comprises the following steps: calculating the speed of lifting the front wheel during takeoff according to the influence of a propulsion system on flow characteristics and pressure distribution outside and inside a runner during takeoff and taxiing of an aircraft; determining an external flow power influence quantity and an internal flow power influence quantity; constructing a pneumatic pitching force and torque expression including the external flow power influence quantity and the internal flow power influence quantity; an aircraft taxiing kinetic equation is constructed based on pitching moment force and moment borne by the aircraft during taxiing and taxiing; a compressed attack angle is obtained based on the compression amount of a main undercarriage strut and tires when a front wheel is lifted during takeoff, and the pneumatic pitching moment of the center of gravity of the aircraft is solved based on the compressed attack angle, a taxiing kinetic equation and the stress condition of the front wheel; based on the pneumatic pitching moment under the gravity center of the aircraft, obtaining the total pneumatic moment acting on the pneumatic reference point, and based on the total pneumatic moment, obtaining the front wheel lifting speed.

This invention discloses a landing gear and an aircraft, relating to the field of aircraft structural components. The landing gear includes a nose landing gear and two main landing gears. The nose landing gear is located at the front of the fuselage and includes a nose wheel and a nose shock strut. One end of the nose shock strut is connected to the fuselage and located inside the fuselage, while the other end extends outside the fuselage and is elastically connected to the nose wheel. The two main landing gears are respectively located on opposite sides of the fuselage. Each main landing gear includes a rocker arm, a main shock absorber, and a main wheel. The rocker arm has a hinged end and a free end, respectively. The hinged end is hinged to the inside of the fuselage so that the rocker arm can only swing along the aircraft's heading. The free end is connected to the main wheel. This invention incorporates part of the nose landing gear structure inside the fuselage and directly hinges the rocker arms of the main landing gear to the fuselage, reducing the overall frontal area of ​​the landing gear, decreasing the aerodynamic drag and overall weight of the aircraft, thereby improving the overall performance and range of the aircraft.