2145 results about "Landing gear" patented technology

An undercarriage, a middle large rotor and a middle large motor are connected to the lower portion of a fuselage body and sequentially connected to the top of a middle small tower, and the right front portion, the left front portion, the left rear portion and the right rear portion of the fuselage body are correspondingly connected with a right front fuselage arm, a left front fuselage arm, a left rear fuselage arm, a right rear fuselage arm and a power assembly. The connecting line of the rotating centers of the rotors of the four power assemblies is square or rectangular, the rotating directions of the rotors of the power assemblies at the same opposite angles are the same, the rotating directions of the rotors of the power assemblies at different opposite angles are opposite, and the rotating faces of the two rotors opposite to the middle large rotor in rotating direction incline around the axes of the respective arms. The included angle between the rotating surfaces of the two rotors and the horizontal plane is an acute angle beta, the rotating surfaces of the two rotors with the same rotating direction as the middle large rotor incline around respective arm axes, the included angle between the rotating surfaces of the two rotors and the horizontal plane is an acute angle theta, beta is larger than theta, the reaction torque and the horizontal component torque of the rotors jointly drive the aircraft to steer, and the aircraft has the advantages of being long in endurance time, large in load and high in wind resistance. And a universal flight platform is formed.

The utility model discloses a landing lever locking device and an unmanned aerial vehicle. The landing lever locking device comprises an unmanned aerial vehicle body. The foot stool assembly comprises a sliding rail seat, a lifting rod and a locking device, and the lifting rod is connected to the sliding rail seat in a sliding mode. The lock catch device comprises a lock box, a sliding block and a driver, the lock box is provided with a vertically-through box lock hole, the sliding block can be connected into the lock cavity in a left-right sliding mode, the sliding block is provided with a vertically-through lock catch hole, the lock catch hole comprises a sliding lock hole and a fastening hole communicated with the sliding lock hole, at least one clamping section is arranged on the side wall of the fastening hole, and the output end of the driver is connected to the sliding block. And the inner wall of the fastening hole is matched with the outer wall of the lifting rod. The lock box is arranged on the unmanned aerial vehicle body, the rising and falling rod is slidably connected to the lock box lock hole and the sliding lock hole, the inner wall of the fastening hole is matched with the outer wall of the rising and falling rod, and the sliding block can be inserted into the rising and falling rod in the radial direction under the driving condition of the driver and is inserted into and clamped to the outer wall of the rising and falling rod through the clamping section. The purpose of clamping the lifting rod can be achieved only by driving the sliding block to move rightwards through the driver.

The invention relates to a tilting rotor unmanned aerial vehicle, in particular to an unmanned aerial vehicle combining a flying wing type structure and a multi-shaft tilting rotor. A completely fused flying wing type fuselage is formed by a low-speed wing shape, and no other fuselage characteristics exist except that bilaterally symmetrical wings are directly connected and fused in the middle; a left tilting rotor frame and a right tilting rotor frame are mounted at the left end and the right end of the flying wing type fuselage respectively; at least one to at most six brackets extend from the front part and the rear part respectively, and symmetrical tilting rotor frames are mounted on the brackets; motors and paddles are mounted on all the tilting rotor frames; and an undercarriage is mounted below the flying wing type fuselage. During tilting, a single motor and a single blade are not tilted, but a plurality of power units formed by one or more motors and blades are tilted in groups through a tilting rotor frame. Therefore, the structure is simplified, and the safety and the maneuverability are improved. The landing gear has two types, one type is a foot type landing gear suitable for vertical take-off and landing in a soft site, and the other type is a wheel type landing gear suitable for short-track take-off and landing in a hardened site.

The invention discloses an aircraft landing gear shimmy test bed and a test method, the aircraft landing gear shimmy test bed comprises a main body rack, a hanging basket system, a large flywheel, a flywheel driving system, an excitation device and a follow-up displacement measurement system.The aircraft landing gear shimmy test bed can simulate the shimmy of an aircraft after undercarriages of different design configurations are subjected to interference excitation force or interference torque in the process of takeoff, taxiing, landing and the like of the aircraft; the control effect of the shimmy damping system on the stability of the undercarriage has short excitation time so as to simulate real working conditions, the shimmy stability performance of the undercarriage can be accurately detected, excitation is continuously carried out under the condition that the test state is continuously changed, and the working efficiency is greatly improved. The follow-up displacement measuring system designed by the invention can enable the exciting force and the center of the axle to be always kept on the same horizontal plane, can apply torque control and torque excitation control, is suitable for undercarriage shimmy tests with different axle heights, different tire diameters and different excitation modes, and has certain universality.

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.

The invention discloses an aircraft front wheel electrical zero calibration method, device and equipment and a medium, relates to the technical field of aircraft front wheel control, is applied to ground comprehensive maintenance equipment, and comprises the steps that a target angle instruction is used for controlling a scale line of a rotating sleeve in a nose landing gear to be aligned with a scale line of a front supporting column outer cylinder; a front wheel shimmy damping instruction is sent to a front wheel operation control box of the aircraft, so that the front wheel operation control box controls front wheels to keep stable; the current angle information of the front wheel angle sensor is determined as current electrical zero position reference information; the front wheel and the rotating sleeve are controlled to rotate through a front wheel rotating instruction, and if the deflection angle of the scale mark of the rotating sleeve is matched with the front wheel rotating instruction, it is judged that front wheel electrical zero calibration of the aircraft is completed. And the accuracy and simplicity of the electrical zero calibration of the airplane front wheel are improved.

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.

The invention relates to the field of aviation noise control, and provides an aircraft landing gear noise reduction device and method based on micro metal wire spiral winding. The device part comprises a metal wire and a metal wire fixing buckle, the metal wire is fixed through the metal wire fixing buckle, and meanwhile the size and material parameters of the metal wire are designed in a targeted mode. In the method part, the designed metal wire is wound on a cylindrical bluff body part, possibly generating vortex shedding and vortex-fixing interference, on the nose gear in a spiral winding manner, and the two ends and the axial interval position of the metal wire are fixed through the metal wire fixing buckles, so that the metal wire is ensured to be fastened. Therefore, through the wound metal wires, single-frequency noise in the over-top direction caused by falling of a towing rod vortex can be effectively eliminated, broadband noise is reduced, and the noise control capability of an airplane is improved. The method is suitable for various civil aircraft types, and has the advantages of being simple in structure, remarkable in noise reduction effect, high in practicability, high in adaptability and the like.

A trailer landing gear deployment device for mechanically raising and lowering a trailer includes a trailer and a landing gear that can be raised and lowered relative to the trailer. A handle is mechanically coupled to the landing gear, which is selectively raised and lowered by rotation of the handle. A disk is rotatably coupled to a housing. An arm is coupled to the disk. The arm is couplable to the handle and rotates the handle when the disk is rotated. A motor is operably coupled to the disk wherein the motor is actuatable to rotate the disk. A power source is electrically coupled to the motor. A control panel is electrically coupled to the power source. The control panel is manipulatable to actuate the motor.

A dual-mode landing gear adapter enables rapid interchangeability between manual and powered operation for trailer landing gear extension and retraction. The adapter includes a gear shaft adapter having a first coupling end structured to securely engage the trailer's landing gear gearbox shaft, and an opposite second coupling end with axial slots configured for receiving a two-pin anchor. A manual crank and a drill head attachment each include a corresponding two-pin anchor for secure, selective coupling with the adapter. The drill head attachment further comprises a cylindrical collar structured to securely receive a standard electrically powered drill / driver, secured via a cotter pin or similar fastening device. The invention facilitates quick, secure switching between powered and manual operation without tools, ensuring convenience, safety, and operational flexibility. This adapter system reduces user effort, increases efficiency, and provides robust reliability under varying operational conditions common in commercial trucking and trailer applications.

The invention provides an undercarriage system scheme design method based on heterogeneous design simulation data management, which comprises the following steps of: firstly, analyzing, abstracting and standardizing different types of heterogeneous design and simulation data in a development process of an undercarriage system, and constructing a uniform abstract data model; secondly, according to the defined semantic rules, an undercarriage semantic ontology model is constructed, then the semantic rules are loaded through a semantic reasoning engine, and processed data are associated; according to the research and development process of the undercarriage system and the semantic association relationship between the data, data storage and display are realized; after data abstract analysis, semantic association and unified storage are completed, functional information, design information and simulation information are further derived based on an undercarriage semantic ontology and a reasoning result, so that an undercarriage system development scheme driven by requirements is constructed. According to the method, multi-source heterogeneous design and simulation data generated in the research and development process of the undercarriage system can be effectively managed, and a design scheme of the undercarriage system is constructed on the basis of the multi-source heterogeneous design and simulation data.

The invention belongs to the technical field of aviation accessory production, and particularly relates to an automatic stamping system for undercarriage supporting rod piece production, the system comprises an outer supporting pipe stamping processing assembly, an inner supporting pipe stamping processing assembly, a lug stamping processing assembly, a connecting seat stamping processing assembly and a stamping assembling assembly which are sequentially arranged along a production line; the outer supporting pipe stamping machining assembly is used for stamping machining of an outer supporting pipe component, the inner supporting pipe stamping machining assembly is used for stamping machining of an inner supporting pipe component, the lug stamping machining assembly is used for stamping machining of a lug component, and the connecting base stamping machining assembly is used for stamping machining of a connecting base component. And the stamping and assembling assembly is used for stamping, riveting and assembling the outer supporting pipe component, the inner supporting pipe component, the lug base component, the rectangular spring component and the connecting base component through rivets to obtain the combined undercarriage supporting rod piece. The punching and polishing device can meet the requirements for automatic punching and punching, punching and polishing of main components in the side supporting rod of the combined undercarriage and overall automatic punching and assembling.

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 utility model is applicable to the technical field of undercarriages, and provides an undercarriage angle measuring device, which comprises a laser displacement sensor and an angle sensor, the angle sensor is used for measuring the rotation angle of a rocker arm around a hollow rotating shaft, the angle measured by the angle sensor is alpha, and the rotation angle of an undercarriage piston rod is alpha / 2. According to the utility model, the laser displacement sensor and the angle sensor are integrated, so that the problem of undercarriage measurement in a complex shape and in a dynamic process is ingeniously solved. The device can accurately measure the rotation angle and the compression amount of the undercarriage piston rod at the same time, does not need complex installation, and is high in adaptability. The high-precision measurement of the laser displacement sensor ensures the accuracy of the expansion amount, and the angle sensor realizes the accurate measurement of the rotation angle through non-contact induction and isosceles triangle geometrical relationship calculation. The whole device is compact in structure and easy to install and maintain, and provides a reliable and accurate measurement means for the dynamic performance test of the undercarriage.

The invention provides an undercarriage system modeling method based on a natural language demand text, and relates to the technical field of model-based computer aided design, and the method comprises the steps: S1, constructing a SysML meta-model knowledge base and an undercarriage system model knowledge base; s2, analyzing the natural language demand text, and generating a structured demand text of the undercarriage system; s3, mapping a structured demand text entity of the undercarriage system into a node type of a semantic graph; and S4, in combination with a historical model template and real-time retrieval, generating an undercarriage system SysML model according to the semantic map obtained in the step S3, and performing undercarriage system engineering analysis through the undercarriage system SysML model, thereby improving the efficiency and quality of aviation equipment development. On the basis of a retrieval enhancement generation technology and an intention routing mechanism, the modeling efficiency and accuracy of the undercarriage system are remarkably improved; a graph neural network and a semantic vector space model are adopted to realize deep fusion and conflict resolution of cross-domain knowledge; and through an incremental learning mechanism, model continuous evolution and design experience closed-loop iteration are supported.

The invention relates to the field of unmanned aerial vehicles, and discloses an intelligent image recognition device based on an unmanned aerial vehicle and a use method thereof.The intelligent image recognition device comprises an unmanned aerial vehicle body, vehicle arms arranged on the periphery of the unmanned aerial vehicle body and rotor wings arranged at the ends of the vehicle arms; stretching assemblies used for protecting the rotor wings in the landing state are arranged at the ends of the vehicle arms, each stretching assembly comprises a sliding rod movably connected to the inner side of the corresponding vehicle arm, and one end of each sliding rod extends to the exterior of the corresponding vehicle arm and is provided with a second protection plate. According to the intelligent image recognition device based on the unmanned aerial vehicle, the arranged extension assembly is matched with the telescopic rod and the pressure sensor which are arranged on the undercarriage, the telescopic rod can be contracted through self gravity when the device falls to the ground, then the pressure sensor is pressed, the driving mechanism is triggered to operate, and therefore the extension assembly is opened; the safety during take-off and landing is improved, the stretching assembly is retracted and hidden after take-off, the size is reduced, and the influence on airflow is reduced.

The invention belongs to the technical field of unmanned aerial vehicles, and discloses a foldable multi-rotor hybrid power unmanned aerial vehicle which comprises lift rotors (10), attitude rotors (20), vehicle arms (30), a cabin (40), a photoelectric pod (50), a cargo hold (60) and an undercarriage (70). The lift rotors and the attitude rotors are combined, the attitude control difficulty is reduced while the large load is guaranteed, and compared with a traditional helicopter, the mechanical structure complexity is reduced, and maintenance and replacement are convenient; and when the power system is redundant, the reliability of the method is ensured through a system fault reconstruction scheme.

The invention discloses a digital twin-driven data-enhanced aviation undercarriage laser shock peening monitoring method, and belongs to the field of aviation structural member remanufacturing, intelligent manufacturing and data driving regulation, and the method comprises the steps: building a fatigue evaluation system for an alternating load in an undercarriage service process, recognizing a high stress concentration region, and carrying out the recognition of a high stress concentration region; a multi-modal monitoring system is constructed, impact data is expanded by adopting a data enhancement method, an impact strengthening physical model is constructed in combination with finite element analysis, impact wave propagation and residual stress evolution are simulated, and collaborative optimization of data driving and mechanism modeling is realized in combination with an experimental data correction model. Impact quality is evaluated based on deep learning, multi-modal data is fused to analyze impact uniformity, and an enhanced abnormal region is identified. And the impact process parameters are optimized through reinforcement learning, and intelligent feedback regulation and control are achieved. According to the method, multi-modal monitoring, data enhancement, physical modeling and intelligent optimization are combined, the stability of laser shock peening is improved, and the fatigue life of laser shock peening is prolonged.

The invention relates to the field of aircrafts and underwater vehicles, in particular to a variable-configuration water-air amphibious cross-medium unmanned aerial vehicle which comprises a watertight electronic cabin assembly, a flight rotor component, a vehicle arm seat, a vehicle arm driving assembly, an undercarriage, a connecting rod and an underwater propelling assembly used for driving the vehicle to move underwater. The flying rotor component is mounted at the upper end of the watertight electronic cabin component and is driven by the arm driving component to be unfolded or folded, so that the unmanned aerial vehicle can be freely converted between two configurations of a four-rotor unmanned aerial vehicle and an underwater vehicle; the arm seats, the undercarriages, the flying rotor components and the connecting rods are hinged to one another to form a parallelogram connecting rod mechanism, and the undercarriages are unfolded and folded along with unfolding and folding of the flying rotor components under driving of the arm driving assemblies. The four-rotor unmanned aerial vehicle can be freely converted between two configurations of a four-rotor unmanned aerial vehicle and an underwater vehicle through the structure variant, so that the requirements of various application scenes such as water rescue, underwater detection and cross-medium monitoring are met.

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 provides a helicopter nose landing gear fatigue test device, and the device comprises a bottom plate which is connected and fixed to a ground grooved rail; the fixed connecting plate is fixed on the bottom plate through a bolt; wherein the lower end connecting structure of the test piece is of a square structure, the periphery of the lower end connecting structure is connected to the fixed connecting plate through rivets, and the airplane wheel fake part assemblies are arranged on the two sides of the upper end connecting structure of the test piece. One end of the loading double-fork lug is connected with the airplane wheel fake part assembly; the torsion-proof bolt is connected with the airplane wheel fake part assembly; one end of the loading single fork ear is connected with the other end of the loading double fork ear; the servo actuation loading mechanism is connected with the other end of the loading single fork lug; wherein the servo actuation loading mechanism is used for realizing loading in the heading direction, the lateral direction and the vertical direction; meanwhile, the invention further provides a helicopter nose landing gear fatigue test method.

The invention relates to the technical field of aircraft maintenance, and particularly discloses a guide supporting structure for replacing large components of an aircraft. The longitudinal supporting seat can be movably arranged on the vehicle body in the horizontal direction; the linkage operation adjusting mechanism is arranged on the side wall of the longitudinal supporting seat and is used for providing two independent transmission powers with variable speeds for the longitudinal supporting seat; the lifting mechanism can be movably arranged on the longitudinal supporting seat in a sleeving mode in a lifting mode, the lifting mechanism is further provided with a rotatable retainer, the retainer is used for fixing an aircraft landing gear in a disassembly and assembly state, and lifting power of the lifting mechanism and rotating power of the retainer both come from the longitudinal supporting seat. According to the invention, multi-degree-of-freedom adjustment in a limited space is realized, and meanwhile, a guide supporting structure with high positioning precision and operation convenience is taken into account, so that the safety and efficiency of replacement operation of large components of an aircraft are improved.

Provided are an in-situ detection device and method for carbon emissions from farmland. The in-situ detection device includes an unmanned aerial vehicle, and landing gear. A retraction and release mechanism is arranged between the unmanned aerial vehicle and the landing gear, and the retraction and release mechanism is configured for retracting or lowering the landing gear. An enrichment hood is installed on the landing gear through a folding mechanism, and the folding mechanism can drive the enrichment hood to be unfolded or folded on the landing gear. When the unmanned aerial vehicle lowers the landing gear to a farmland gas collection point through the retraction and release mechanism, the enrichment hood is unfolded to enrich a gas at the farmland gas collection point. The landing gear is also provided with a greenhouse gas detection mechanism for detecting the gas enriched in the enrichment hood.

A motorized landing gear including: a motor housing that encloses a motor; a gearbox; a controller; and a power source powering the motor and the controller. The gearbox attaches to an axle having two ends, with a first end of the axle connecting to a first landing gear leg and a second end of the axle connecting to a second landing gear leg. The motor connects to said gearbox and the controller instructs the motor to rotate. The rotation of the motor transfers power to the gearbox, which rotates the axle. The rotation of the axle in a first direction simultaneously telescopes the landing gear first leg and the landing gear second leg; and the rotation of the axle in a second direction simultaneously retracts the landing gear first leg and the landing gear second leg.

A landing gear assembly configured to support a trailer includes a leg assembly including a first leg member and a second leg member telescopingly received within the first leg member and adjustable between a retracted position and an extended position, a gear assembly operably coupled to the leg assembly and configured to move the first leg member between the retracted and extended positions, wherein the gear assembly includes a first shaft member, and a first bushing member rotatably supporting a first end of the first shaft member within the first leg member, wherein the first bushing member is attached to the first leg member via at least one first mechanical fastener, and wherein the first bushing member comprises a powdered metal.

The invention provides an anti-offset magnetic coupling mechanism for a wireless charging system of an unmanned aerial vehicle, the magnetic coupling mechanism comprises a transmitting end arranged on a landing platform and a receiving end mounted on the unmanned aerial vehicle, and a transmitting coil is a double-layer circular coil wound at gradient intervals; an axial magnetic field which is uniformly distributed is formed through a non-uniform distribution structure of coil turn spacing; the receiving end is composed of four sets of same magnetic core-coil structures, each set is composed of a lightweight cylindrical novel soft magnetic composite material magnetic core and an externally-wound solenoid coil of a vertical landing gear of the unmanned aerial vehicle, the magnetic cores are located in the vertical landing gear of the unmanned aerial vehicle, and the coils of the four magnetic core-coil structures are closely wound in the axial direction of the outer surface of the landing gear and connected in series. And coil and magnetic core parameters are designed by using an evolutionary algorithm. The problems that in a traditional scheme, the anti-offset capacity is weak, the magnetic core weight is large, and the receiving end is not matched with the special-shaped structure of the unmanned aerial vehicle are solved, and the method is suitable for unmanned aerial vehicle wireless charging scenes.

The invention belongs to the field of vertical take-off and landing aircrafts, and relates to a distributed contra-rotating rotor hybrid tilting rotorcraft. Comprising a front tilting electric contra-rotating rotor nacelle, a front cabin body, a wing end tilting electric contra-rotating rotor nacelle, a V-shaped empennage, a V-shaped empennage aileron, a vertical empennage, a retractable undercarriage, a green hybrid power energy system and an empennage tilting electric contra-rotating rotor nacelle, the front tilting electric contra-rotating rotor nacelle is hinged with the other end of the front cabin body; an empennage tilting electric contra-rotating rotor nacelle is arranged at the end part of the V-shaped empennage; wing end tilting electric contra-rotating rotor wing nacelles are arranged at the ends of the wings; and rotor wing units are arranged in the rotor wing nacelles, the six rotor wing units all adopt a coaxial reverse upper and lower two-paddle structure, and the rotating directions of the rotor wings in the six rotor wing units are that the adjacent rotating directions and the opposite rotating directions are opposite.

The invention particularly relates to a numerical control machining path optimization method for undercarriage shell machining, and relates to the technical field of numerical control machining. Generating an initial processing path and setting parameters; adjusting parameters driven by a path optimization algorithm; and performing virtual simulation verification and path correction. According to the method, Dijkstra and genetic algorithms are used for optimizing the machining sequence and path, strategies such as equal residual height control are adopted for balancing the cutting load, and the optimal cutting parameters are solved through a multi-objective function. The relation among the machining time, the surface roughness and the tool abrasion loss is fully considered in the process, and collaborative optimization of the machining efficiency, the machining quality and the tool service life is achieved. And the problems of low machining efficiency, substandard surface quality, premature wear of the cutter and the like caused by unreasonable parameters are avoided, the production cost is effectively reduced, and the production benefit is improved.

An unmanned aerial vehicle (UAV) includes a UAV body constructed and arranged to carry a payload, a set of UAV wings coupled with the UAV body, and a set of undercarriage apparatus coupled with at least one of the UAV body and the set of UAV wings. The set of UAV wings is constructed and arranged to provide lift to the UAV. At least one undercarriage apparatus of the set of undercarriage apparatus includes an elastomeric vehicle mount, a leg, and a coupling constructed and arranged to couple the leg with the elastomeric vehicle mount. The elastomeric vehicle mount is constructed and arranged to absorb energy provided from the leg through the coupling.

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.