86results about "Undercarriages" patented technology

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 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.

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.

Aircraft landing gear system (100, 200, 300), comprising actuators associated with the landing gear and supplied with pressurized hydraulic fluid by a hydraulic circuit (10; 15), including at least hatch and landing gear attachment boxes (106, 109, 206, 209, 306, 309), the aircraft comprising emergency extension means (20) capable of ensuring the release of the hatches and landing gear to cause the landing gear to extend under the effect of gravity in the event of failure of the hydraulic circuit. According to the invention, the emergency extension means comprise a pyrotechnic device (50) adapted to generate, in a controlled manner, pressurized gas acting on a first face of a piston (40) sliding with a seal in a chamber (30) such that a second face of the piston, in contact with the hydraulic fluid, transmits the pressure of the gas thus produced to at least the attachment boxes via the hydraulic fluid. FIGURE IN ABRIDGED FORM: Fig. 2

An aircraft having a landing gear provided with a stand and a pivoting assembly, the landing gear having a pivot locking system comprising a movable locking finger and a passage provided in a base. A monitoring system is configured to determine the presence of a risk of sudden yaw movement of the aircraft in the event of a take-off or an unlocking of the landing gear when the locking finger is in the locked mode on the ground, the monitoring system being configured to emit a risk signal to a processing system following a determination of the presence of the risk of sudden yaw movement.

A method for extending a landing gear system for a vehicle is disclosed. The landing gear system includes a first landing gear assembly and a second landing gear assembly. The method includes the steps of sensing a first load on the first landing gear assembly after the first landing gear assembly has reached a wheels-on-ground state and comparing the first load to a first target value. The method further includes the step of controlling an extension speed of the first landing gear assembly according to a difference between the first load and the first target value.

The present application relates to the technical field of aircraft landing gears, and provides an aircraft landing gear stay lock. The landing gear stay lock comprises: a front lock stay, a rear lock stay, an unlocking rocker arm, an unlocking actuator cylinder, locking damping actuator cylinders, and locking springs. One end of the front lock stay is connected to a landing gear stay, and the other end of the front lock stay is hingedly connected to the rear lock stay. The other end of the rear lock stay is hingedly connected to a landing gear strut and is hingedly connected to the locking damping actuator cylinders. The unlocking rocker arm is arranged at the position where the front lock stay is hingedly connected to the rear lock stay. One end of the unlocking actuator cylinder is connected to the front lock stay, and the other end thereof is connected to the unlocking rocker arm. One end of each locking damping actuator cylinder is hingedly connected to the rear lock stay, and the other end thereof is connected to the corresponding locking spring. The other end of the locking spring is connected to the front lock stay. Respective ends of each locking spring and the corresponding locking damping actuator cylinder which are connected are arranged on the landing gear strut.

The invention relates to a method for powering actuators of an aircraft comprising at least one generator (500), the actuators being associated with landing gear of the aircraft and comprising at least one actuator for providing at least one of the following functions: A / extending / retracting the landing gear, B / steering the steerable wheels of the landing gear, C / normal braking of the aircraft; and one actuator for providing at least one of the following functions: D / applying the parking brake of the aircraft, E / extending the landing gear in emergency mode, F / braking in emergency mode. The method of the invention comprises the steps of: • - powering the actuator providing at least one of the functions A / , B / , C / directly with the generator; • - powering the actuator providing at least one of the functions D / , E / , F / by means of an accumulator (501) connected to the generator by a power line (502) comprising a normally closed switching member (503) for isolating the accumulator from the generator so as to have two independent and segregated power sources, each powering actuators that perform separate functions, the switching member (503) being controllable to open and allow the accumulator to be charged by the generator.

Landing gear (10) for an aircraft (100), comprising a wheel (20) or a skid, a shock absorber (30) with an absorber connection (32) to connect to a fuselage (110) of the aircraft (100) and at least one attachment lever (40) with a lever connection (42) to connect to the fuselage (110) of the aircraft (100) spaced apart from the absorber connection (32), the wheel (20) or skid being attached to the attachment lever (40) spaced apart from the lever connection (42), wherein the shock absorber (30) is connected to the wheel (20) or skid directly or indirectly via a wheel or skid connection (34) to provide a load path (LP) for vertical load from the wheel (20) or skid to the absorber connection (32), and wherein the shock absorber (30) comprises a fuse pin (36) at the absorber connection (32) and / or the wheel or skid connection (34) having a predetermined breaking load to break and release the connection at the absorber connection (32) and / or the wheel or skid connection (34), and wherein the attachment lever (40) comprises a rotary bearing (41) providing a folding movement for the wheel (20) or skid from an extended landing position (ELP) into a crash position (CP) when the fuse pin (36) breaks.

A door fairing for a retractable landing gear leg of an aircraft is disclosed having an outer member which covers a portion of a landing gear bay of the aircraft when the landing gear leg is retracted. An inner member of the door fairing is coupled to the outer member via a number of configurable couplings which allow adjustment of the position of the outer member with respect to the inner member. The inner member has mounting points for mounting the door fairing to the landing gear leg. Prior to fitting the door fairing at the aircraft, measured or derived dimensional parameters of the landing gear bay are used to adjust the configurable couplings such that the door fairing is preconfigured to fit the landing gear bay of the aircraft.

A landing gear door retraction device is disclosed, which can be used to open and close the landing gear door of a reusable aircraft. It is characterized by comprising: a door actuator and multiple sets of linkage mechanisms. The device uses an independent door actuator to drive multiple sets of four-bar linkages, making it suitable for aircraft with high flight speeds, limited cabin space, and multi-leaf landing gear doors. It offers significant advantages such as light weight, small size, simple timing, and good synchronization. When the landing gear door is fully open, all three sets of four-bar linkages are near their dead center positions, preventing damage to the door structure caused by actuator overtravel. The door retraction device is located near the side and top walls of the landing gear bay, requiring minimal space and fully utilizing the remaining cabin space after the landing gear is retracted.

Aircraft landing gear comprising: a strut mounted on an aircraft structure such that it can move between an extended position and a retracted position; a bracing member for holding the strut in the extended position, comprising a first connecting rod hinged to the aircraft structure and a second connecting rod hinged to the first connecting rod and to the strut; a stabilising member for maintaining the first and second connecting rods in an aligned position, comprising a first connecting rod and a second connecting rod which are hinged to each other and at least one of which is hinged to the bracing member; and a spring for resiliently returning the hinges of the first and second connecting rods to a generally aligned position. The spring comprises at least one leaf spring designed to be subject to an axial compression force when the first and second connecting rods move out of the generally aligned position.

The application discloses a kind of anti-jam emergency unlocking electromechanical actuators, realize the excess of different working medium emergency anti-jam emergency unlocking.The application is realized by the following technical solutions: the piston head tail part is designed with the steel ball guide hole of assembly rigid lock ball, in emergency unlocking, high-pressure emergency medium enters electromechanical actuator outer cylinder cavity from the emergency energy inlet on the tail part of outer cylinder, emergency sliding sleeve is pushed by high-pressure medium of emergency release circuit, compresses the upper locking force spring, the upper locking sliding sleeve of ladder cylinder is pushed from the inner ring of rigid lock ball by high-pressure medium and is separated, rigid lock ball rolls along steel ball guide hole, the upper locking sliding sleeve of ladder cylinder reaches limit position, is separated from lock groove, realizes the mechanical lock unlocking of rigid lock ball, after unlocking, high-pressure medium pushes piston cylinder to extend, and simultaneously, sliding block is radially slid along ring cloth lock groove and is separated from neck lock groove, separates from the inner surface of emergency sliding sleeve, in emergency sliding sleeve reaches limit position, realizes the anti-jam mechanism emergency unlocking in the hollow cavity of piston head.

The present invention relates to a fibrous preform 1 of a core part of a fibrous reinforcement for a composite material part, wherein the preform is formed by three-dimensional weaving of a first thread extending along the longitudinal direction and a second thread crossing the first thread. In the fibrous preform, the preform comprises two longitudinal ends 3a; 3b for articulation with other parts and an intermediate region ZM located between the two longitudinal ends. Each longitudinal end has a thickness greater than the thickness of the intermediate region, the intermediate region has a first volume ratio of the first thread to the second thread greater than 1, and each longitudinal end has a second volume ratio of the first thread to the second thread smaller than the first volume ratio and closer to 1 than this first volume ratio.

A low-power consumption leg type landing gear and helicopter based on tether coupling relates to the technical field of helicopter landing gear. The existing helicopter landing gear has high energy consumption in the process of ground parking or air hovering and attitude locking, which additionally increases the power consumption of the helicopter and leads to the decline of the endurance. The present application comprises a support frame and at least one set of landing support unit; each set of landing support unit contains a support leg, a differential mechanism, a hip joint rope drive mechanism and a knee joint drive mechanism. Two support legs are symmetrically arranged on both sides of the support frame for helicopter landing support; the differential mechanism corresponds to the support leg one by one to regulate the hip joint angle; the output power of each hip joint rope drive mechanism is divided into two ways and transmitted to the differential mechanism, and after differential adjustment, the support leg is driven to realize angle adjustment; the knee joint drive mechanism adjusts the knee joint angle of the support leg. The tether coupling and differential adjustment structure are adopted to reduce the energy consumption while meeting the landing support and terrain adaptation, and to improve the endurance of the helicopter.

The present invention relates to an aircraft having a landing gear (10) equipped with a base (11) and a pivoting assembly (12), said landing gear (10) having a pivoting locking system (20) comprising a movable locking finger (25) and a passage (22) formed in a base (21). A monitoring system (50) is configured to determine the presence of a risk of sudden yaw of the aircraft in the event of takeoff or unlocking of the landing gear (10) when the locking finger (25) is in the ground-locked mode, the monitoring system (50) being configured to transmit a risk signal to a processing system (60) following a determination of the presence of said risk of sudden yaw. (Shortcut figure: Figure 3)

A shroud assembly for an aircraft landing gear includes a pneumatic shroud, a first support bracket assembly, and a second support bracket assembly. The first support bracket assembly is configured to couple with a structural member of the aircraft landing gear, support a first end of the pneumatic shroud, and have a first position fixed relative to the structural member in an x-direction, a y-direction, and a z-direction. The first support bracket assembly has a first clamp configured to fix the first support bracket assembly relative to the structural member in the x-direction. The second support bracket assembly is configured to support a second end of the pneumatic shroud and have a second position fixed relative to the structural member in the y-direction and the z-direction.

The invention relates to an aircraft landing-gear system (100, 200, 300), comprising actuators associated with the landing gear and supplied with pressurized hydraulic fluid by a hydraulic circuit (10; 15), the hydraulic circuit including at least uplock boxes (106, 109, 206, 209, 306, 309) for the hatches and the landing gear, the aircraft comprising emergency extension means (20) capable of ensuring release of the hatches and of the landing gear in order to cause the landing gear to extend under the effect of gravity in the event of failure of the hydraulic circuit. According to the invention, the emergency extension means comprise a pyrotechnic device (50) suitable for generating, in a controlled manner, pressurized gas that acts on a first face of a piston (40) which slides in a sealed manner inside a chamber (30) such that a second face of the piston, in contact with the hydraulic fluid, transmits the pressure of the gas thus produced at least to the uplock boxes via the hydraulic fluid.

A linear actuator may include a piston and a linear drive. The piston is slidably mounted for sliding along an axis. The linear drive includes a ball screw and a ball nut operably mounted on the ball screw, with rotation of the ball screw causing the ball nut to move axially along the ball screw. The piston is attached to the ball nut, and the piston is configured to be advanced along the axis as the ball screw is rotated.

A landing gear assembly for an aircraft, the landing gear assembly comprising: a forward brace having a first end for attachment to an airframe of the aircraft, and a second end; an aft brace having a first end for attachment to the airframe of the aircraft, and a second end, wherein the second end of the aft brace is coupled to the second end of the forward brace; an actuator piston having a first end and a second end, wherein the first end of the actuator piston is pivotally attachable to the airframe of the aircraft; a bellcrank linkage having a first end, a second end, and a pivot point between the first end and the second end, such that a first portion of the bellcrank linkage extends from the first end to the pivot point, and a second portion of the bellcrank linkage extends from the pivot point to the second end, wherein the pivot point of the bellcrank linkage is pivotally attachable to the airframe of the aircraft, and wherein the first end of the bellcrank linkage is pivotally coupled to the second end of the actuator piston; a primary strut having a first end and a second end, wherein the first end of the primary strut is pivotally coupled to the second end of the forward brace and the second end of the aft brace, and wherein one or more wheels of the aircraft are mounted at the second end of the primary strut; and a secondary strut (e.g. a shock absorber) having a first end and a second end, wherein the first end of the secondary strut is pivotally coupled to the second end of the bellcrank linkage, and wherein the second end of the secondary strut is pivotally coupled to a pivot point near to the second end of the primary strut; wherein the landing gear assembly is configured to adopt a deployed configuration when the actuator piston is in an extended configuration, and wherein the landing gear assembly is configured to adopt a retracted configuration when the actuator piston is in a retracted configuration. Also provided is an aircraft (such as a blended wing body aircraft) having such a landing gear assembly.

Shield assemblies for an aircraft landing gear include an aerodynamic shield, a first support bracket assembly, and a second support bracket assembly. The first support bracket assembly is configured to couple with a structural member of the aircraft landing gear, to support a first end of the aerodynamic shield, and to have a first position that is fixed relative to the structural member in an x-direction, a y-direction, and a z-direction. The first support bracket assembly has a first clamp that is configured to fix the first support bracket assembly relative to the structural member in the x-direction. The second support bracket assembly is configured to support a second end of the aerodynamic shield and to have a second position that is fixed relative to the structural member in the y-direction and the z-direction.

The invention relates to an electro-hydraulic system (1) for controlling the landing gear (2) of an aircraft, the landing gear (2) comprising a plurality of retractable landing gear units (21, 22), each landing gear unit being provided with at least one wheel (25, 26). The system (1) comprises two braking devices (11, 12), a moving device (13, 14, 15) for each landing gear unit, and two hydraulic units (16, 17). Each hydraulic unit comprises a reservoir (161, 171) containing a fluid, a pump (162, 172) and a distributor (164, 174) distributing the fluid to at least one braking device (11, 12) and to the moving device (13, 14, 15) through a braking valve (165, 175), an extension valve (166, 176), a retraction valve (168, 178) and a discharge valve (167, 169, 177, 179).

A method diagnoses a state of deterioration of an uplock having an unlocking actuator provided with a hydraulic cylinder. The method includes the step of supplying the cylinder with pressurized fluid with an evolution law according to which the pressure of the fluid is, for a predetermined period, equal to a predetermined pressure lower than a full pressure, then equal to the full pressure. The method further includes the steps of determining a time of deactivation of the uplock and comparing it with the time when the predetermined time period ends; and deducing therefrom a state of deterioration of the uplock.

The utility model discloses a fixed wing aircraft diagonal bracing assembly structure, including fuselage and the main wing on the fuselage, the bottom of main wing is equipped with landing gear, its characterized in that, the both sides of landing gear are equipped with diagonal bracing mechanism no.