123 results about "Helicopter rotor" patented technology

The invention relates to a helicopter rotor crack fault identification method and device based on video semantic segmentation. The method comprises the steps that video collection and denoising and stability enhancement processing are carried out through an unmanned aerial vehicle; through designing a global-local feature interaction double-branch network, parallel computing and bidirectional fusion are carried out on context global feature extraction branches and detail mining branches, so that rotor wing crack feature extraction is realized; designing a hierarchical crack feature reconstruction decoder to realize cross-scale fusion and spatial precise alignment of crack features; and designing a multi-frequency-domain boundary sensing enhanced training head, and realizing multi-scale feature extraction and dynamic weight distribution of the crack edge through the synergistic effect of a multi-stage wavelet frequency domain decomposition module and a self-adaptive boundary weighting supervision module. Through the innovative modules, high-precision and robust crack detection is realized aiming at the problems of view limitation, global-local feature fusion difficulty, detail loss caused by down-sampling, edge blur, class imbalance and the like in helicopter rotor crack detection.

The invention relates to the technical field of helicopter aerodynamic force simulation, and discloses a sea rescue helicopter rotor aerodynamic force simulation method and system. The method comprises the following steps: acquiring a rotor wing downwash airflow historical speed data set, a historical sea surface wave data set and a historical aerodynamic pressure data set monitored within a preset historical time range; executing a dynamic volatility analysis process on the three types of historical data sets to generate a speed volatility index, a sea surface wave volatility index and an aerodynamic pressure volatility index; taking the three types of volatility indexes as guide parameters, performing adaptive search operation in a rotor aerodynamic force simulation parameter database, and determining an optimal simulation time interval; and performing numerical calculation simulation on pneumatic coupling between the rotor wing and the sea surface according to the optimal simulation time interval to generate a simulation data sequence, performing feature extraction and classification on the simulation data sequence by using an aerodynamic force recognition model, and outputting a rotor wing aerodynamic force simulation result.

The invention belongs to the technical field of helicopter rotor calibration, and discloses a rotor total pitch calibration device and method.The rotor total pitch calibration device comprises a variable pitch pull rod, one end of the variable pitch pull rod is connected with a rotor fork lug, the other end of the variable pitch pull rod is connected with a steering engine rocker arm fork lug, and rotor blades are calibrated by adjusting the axial length of the variable pitch pull rod and the relative angle of the end faces of knuckle bearings at the two ends; the auxiliary calibration device is provided with an angle measurement mechanism and a distance measurement mechanism, the distance measurement mechanism is connected to the angle measurement mechanism and can move relatively, and the angle measurement mechanism and the distance measurement mechanism are connected with the knuckle bearings at the two ends of the variable-pitch pull rod through positioning connectors respectively; the gradienter is used for measuring the mounting angle of the rotor blade at the reference position and feeding back the mounting angle to the ground detection equipment; the ground inspection equipment converts the mounting angle into a variable pitch amount and sends the variable pitch amount to the flight control system; and the flight control system generates a control instruction according to the variable pitch quantity and sends the control instruction to a steering engine, and the steering engine manipulates a variable pitch pull rod to achieve total pitch calibration of the rotor wings. According to the method, the problem that existing helicopter rotors are difficult to calibrate is solved, and the total pitch of the rotors can be calibrated with high precision and high efficiency.

The utility model discloses a mounting structure of a helicopter rotor shaft dynamic sealing ring dismounting tool, which comprises a base, a dismounting head is arranged at the top of the base, the dismounting head comprises a push-and-pull disc, vertical rods are fixedly connected to the top of the push-and-pull disc in an annular array, a top disc is fixedly connected to the tops of the vertical rods, and the top disc is fixedly connected to the top of the push-and-pull disc. Hinge grooves penetrating to the bottom are formed in the top of the top disc in an annular array mode, hydraulic push rod sleeve connecting rods are fixedly connected to the bottoms of the inner sides of the multiple vertical rods, hydraulic push rod sleeves are fixedly connected to the inner sides of the multiple hydraulic push rod sleeve connecting rods, and second hydraulic push rods are fixedly connected to the inner walls of the hydraulic push rod sleeves. Through the arrangement of the dismounting head, the dismounting process of the sealing ring and the rotor shaft is simplified through the design of the dismounting tool, the dismounting efficiency and safety are remarkably improved, through accurate control of a hydraulic system, it is ensured that dismounting force is evenly applied, and the problem that the sealing ring is damaged or is not thoroughly dismounted due to the fact that the force is too large or too small is solved.

The invention relates to a measurement and inversion reconstruction method for a helicopter rotor sound radiation ball in an acoustic wind tunnel, and relates to the technical field of helicopter aeroacoustics wind tunnel test technology and sound target characteristics, and the method comprises the steps: measuring the noise radiated by a helicopter rotor in the acoustic wind tunnel through a hemispherical far-field noise measurement device; noise signals of the rotor in different directions are obtained, and a rotor sound radiation ball model is constructed; based on a sound propagation theory in a flow field, constructing a propagation model from the surface of the sound radiation ball model to the position of a microphone, and based on the influence of background noise, establishing a sound propagation equation from the surface of the sound radiation ball model to the microphone in the hemispherical far-field noise measurement device; and converting the sound propagation equation into solving a least square problem, solving to obtain sound source distribution on the surface of the sound radiation ball model, and reconstructing noise at any position of the far field of the helicopter rotor. The method can be suitable for reconstructing the helicopter rotor sound radiation ball in complex environments such as acoustic wind tunnels and high background noise, and practical research and application are carried out.

The utility model relates to an unmanned helicopter and its rotor and folding structure and belongs to the technical field of unmanned plane. The folding structure comprises a hub, an end of which is provided with a paddle clamp, two upper and lower through mounting holes are arranged on the paddle clamp, a paddle blade, an end of which is provided with two connecting holes matched with the mounting holes, two stepped bolts, each of which passes through one mounting hole and one connecting hole to connect the paddle blade to the paddle clamp, and the outer peripheral wall of the cylinder part of the stepped bolt is a smooth surface, when any stepped bolt is removed, the paddle blade can rotate around the outer peripheral wall of the cylinder part of the other stepped bolt to fold the rotor. The folding structure of the rotor of the unmanned helicopter of the utility model can make the paddle blade rotate around the other stepped bolt when one stepped bolt is removed, so that the folding of the rotor can be conveniently realized, and the reconnection can be conveniently carried out after the paddle blade is rotated to the original position.

The application belongs to the field of structural strength design, and relates to a helicopter metal moving part high-cycle fatigue damage evolution method considering load sequence. The method comprises the following steps: according to a high-cycle fatigue damage model considering stress ratio and load sequence effect, based on finite element theory, evaluating a high-cycle fatigue damage evolution process of a helicopter rotor metal structure under variable amplitude loading, and predicting a high-cycle fatigue life.

The invention provides a helicopter rotor aerodynamic noise data acquisition and processing method and device, and the method comprises the steps: firstly acquiring rotor aerodynamic noise data which comprises a plurality of aerodynamic noise observation point positions and noise sound pressure levels; the noise sound pressure level of each observation point position before interpolation is converted into noise sound pressure, then the noise sound pressure of each observation point position after interpolation is calculated and converted into the noise sound pressure level, and the relation between each observation point position after interpolation and the noise sound pressure level is established and written into a specified format file; according to the method, the robustness and the accuracy of restoring the omnidirectional sound field distribution of the rotor wing by adopting discrete and non-uniformly distributed test points can be improved, the omnidirectional sound field distribution of the rotor wing can be calculated based on simulation data of a small amount of observation points, and the analysis efficiency of the aerodynamic noise of the rotor wing is improved.

This invention relates to a device and method for measuring helicopter rotor blade deformation, comprising: two semiconductor lasers, two sinusoidal gratings, two linear CCD arrays, and a data receiving and processing unit. The semiconductor lasers emit light beams as optical cutters. The sinusoidal gratings receive the light beams, which then fall onto the spanwise plane of the rotor blade, forming stripes. The linear CCD arrays collect all the stripes on the blade surface. The data receiving and processing unit receives the stripe information collected by the linear CCD arrays and applies Fourier transform profilometry to obtain the three-dimensional surface profile of the rotor blade under test. Furthermore, by comparing the three-dimensional surface profiles of the rotor blade under test at different rotor speeds, the amount of blade deformation is obtained. Determining the blade deformation at two different rotor speeds using the three-dimensional surface profile of the blade improves the accuracy of blade deformation measurement.

The application relates to a rotor forming technical field, in particular to a processing device for helicopter rotor forming, which comprises a lower mold core and a matched upper mold core, a gate is arranged on the upper mold core, and a flow circulation mechanism is arranged on the lower mold core and the upper mold core. The processing device for helicopter rotor forming is characterized in that high-pressure inert gas is controlled to enter a forming cavity through a gas valve, pressure is synchronously applied in a casting process, and the compactness of a casting is improved; a right-side reserved cavity serves as a metal liquid buffer area, cooperates with inert gas pressure adjustment, ensures that metal liquid uniformly fills a complex cavity, and reduces problems such as cold separation or insufficient filling; a sealing mechanism maintains a high-temperature and high-pressure environment, so that casting and HIP are completed in the same mold cavity, secondary oxidation or cooling defects caused by traditional step-by-step processes are avoided, and a production cycle is shortened. Therefore, the design of the integrated pressure die casting and HIP and inert gas dynamic pressure realizes batch production of high-precision complex connecting rods.

The application provides a helicopter rotor system, wherein a cavity in the inside of a ring-shaped shell is provided with a ring-shaped cabin, a central space in the outside is provided with a main rotor structure through a main speed reducer support frame, a rear end is provided with a tail rotor cabin through a tail rotor shell, and a tail rotor is arranged in the tail rotor cabin; an engine is arranged at the rear of the ring-shaped cabin, outputs power to the tail rotor through a power shaft in the rear direction, outputs power to the main speed reducer through a transmission shaft in the front direction, and transmits the power after reduction to a main rotor hub and main rotor blades through a main rotor shaft; a small pull rod is hinged on an outer edge fork of a moving ring through a ball bearing at the lower end and is hinged on a horizontal shaft at the outer end of the main rotor blade through a ball bearing at the upper end; a plurality of rudders are evenly distributed on the same horizontal plane and are vertically installed on the main speed reducer support frame; the application adopts a non-open, supported and limited wing tip design, so that the linear speed of the wing tip of the main rotor of the helicopter can easily break through the sound speed without strong vibration, and a new scheme is provided for improving the speed and load of the helicopter.

The utility model provides a movable Bell helicopter propeller hub frame. The movable Bell helicopter propeller hub frame comprises a placing table; the four supporting pieces are arranged around the bottom of the containing table, a bottom platform is fixedly installed at the bottoms of the supporting pieces, and a setting groove is formed in the middle of the top of the containing table. According to the movable Bell helicopter propeller hub frame, the four containing plates are installed at the top of the containing table, the Bell helicopter propeller hub is placed through the containing grooves, the propeller hub can be fixed and does not slide, the safety of the propeller hub is guaranteed, oil drops dripping from the propeller hub can be received through the grooves and the oil receiving disc, pollution is avoided, and the service life of the propeller hub is prolonged. The propeller hub frame main body is formed by processing aluminum alloy sections, so that the propeller hub frame has the characteristics of light weight, high rigidity and good stability; universal wheels and directional wheels are arranged at the bottom of the propeller hub frame, so that free movement in a plane can be realized; a fixed supporting device is arranged at the bottom of the propeller hub frame, the propeller hub frame can be fixed to the ground and does not slide at will, and the functionality of the propeller hub frame is improved.

The present application relates to a kind of acoustic wind tunnel inside helicopter rotor sound radiation sphere measurement and inversion reconstruction method, it is related to helicopter aerodynamic acoustics wind tunnel test technology and acoustic target characteristic technical field, include: in acoustic wind tunnel using hemispherical far-field noise measuring device measurement helicopter rotor radiated noise, obtain the noise signal of different azimuth of rotor and construct rotor sound radiation sphere model;Based on the sound propagation theory in flow field, construct the propagation model of the surface of sound radiation sphere model to microphone position, and based on the influence of background noise, establish the sound propagation equation of the surface of sound radiation sphere model to microphone in hemispherical far-field noise measuring device;Sound propagation equation is converted to solve least square problem, and the sound source distribution of the surface of sound radiation sphere model is obtained, and the noise of helicopter rotor far field arbitrary position is reconstructed.Can be applied to in acoustic wind tunnel, high background noise and other complex environments reconstruct helicopter rotor sound radiation sphere, carry out practical research and application.

The invention belongs to the technical field of aircraft design, and particularly relates to a helicopter rotor mechanism and an aircraft. The mechanism comprises a coaxial double-rotor assembly which comprises an upper rotor unit and a lower rotor unit; the coaxial double-rotor variable-pitch assembly is connected with the coaxial double-rotor assembly and used for adjusting the propeller pitch of the coaxial double-rotor assembly; the cycloidal propeller assembly is vertically mounted between the upper rotor wing unit and the lower rotor wing unit; and the cycloid propeller variable pitch assembly is connected with the cycloid propeller assembly and used for adjusting the propeller pitch of the cycloid propeller assembly, so that the stress of the helicopter in the horizontal direction is adjusted, the posture of the helicopter is adjusted, and control over the position in the air is achieved.

This invention discloses a flow diversion device and a helicopter that enhances helicopter performance by utilizing the coupling of internal and external flow fields. The flow diversion device includes a sliding support frame, a grooved stabilizing wing assembly, and a helical crossflow fan. The sliding support frame is disposed within the helicopter's fuselage and is slidably extendable to drive the grooved stabilizing wing assembly and the helical crossflow fan to extend through the mounting port and approach the air intake. The upper end of the grooved stabilizing wing assembly is recessed to form a mounting groove. The helical crossflow fan is rotatably mounted within the mounting groove and rotates under the action of the rotor downwash generated during helicopter rotor rotation. This rotation draws the rotor downwash and forward or side-flight turbulence into the mounting groove to form an accelerated airflow, which is then accelerated and discharged from the mounting groove before entering the air intake. This flow diversion device increases the intake volume and pressure of the engine, reduces most of the engine power loss caused by the coupling of internal and external flow fields, and improves engine output power.

The invention discloses a high-confidence helicopter rotor flow-solid-sound coupling calculation method, system and device, and relates to the technical field of helicopter design, the method comprises the following steps: transmitting a blade aerodynamic force as a dynamic load to a blade kinetic model, advancing a time step, and calculating a blade elastic response; meanwhile, a balancing model is introduced, one time step is propelled according to blade aerodynamic force distribution, and the rotor control amount is updated; when the dynamic load and the blade elastic response converge, calculating the total noise of the rotor wing according to the dynamic load distribution and the blade elastic response in a preset period; when convergence is not carried out, the blade appearance and the surface element state of the blade model are updated according to the blade elastic response, a time step is propelled, the motion state of the blade model is adjusted according to the updated rotor wing manipulated variable, new vortex particles are generated through the rotor wing wake model, blade aerodynamic force distribution is updated, and next iteration is carried out. According to the method, the precision and the efficiency can be considered in rotor wing fluid-solid-sound coupling analysis.

The application discloses a multi-mode multi-rate satellite signal transmitting device suitable for a rotorcraft, and relates to the technical field of satellite communication. The device comprises a monitoring generation module, a multi-mode control module, a service access module, a fountain coding module, an interface management module, a coding interleaving module, a data group framing module, a data retransmission module, a data modulation module and an intermediate frequency processing and control module. The multi-mode control module receives modulation coding information and transmission mode information transmitted by the monitoring generation module, and sends the transmission mode information to the fountain coding module, the coding interleaving module and the data retransmission module respectively, sends modulation mode information to the data modulation module, and sends coding mode to the coding interleaving module. The application can overcome the influence of rotorcraft rotor shielding on satellite signal transmission, adopts a multi-mode switching technology, and can select a suitable anti-rotor shielding mode according to different rotorcraft rotor speeds and different satellite conditions.

This invention relates to the field of tail gear reducer bearing lubrication technology, and discloses a helicopter rotor pitch control shaft bearing lubrication structure and a tail gear reducer having the same. The helicopter rotor pitch control stick bearing lubrication structure and tail gear reducer having the same include a tail gear reducer main housing, a control shaft, a tail gear reducer output gear, a guide seat, and a control stick assembly. The lubrication structure for the helicopter tail gear reducer control shaft bearing provided by this invention collects splashed lubricating oil by setting a high-level oil collection groove, and reliably delivers lubricating oil to the upwardly inclined control shaft bearing by utilizing the first inner cavity, second inner cavity, third inner cavity, and corresponding oil guide holes that are sequentially connected in the guide seat, control stick assembly, and bearing housing. This improves the original grease lubrication method to oil lubrication, enhances the bearing's heat dissipation capacity, and effectively avoids the risk of seizure caused by heat accumulation.

This invention discloses an adjustable fixture for CNC machining of rotor blades of manned helicopters, relating to the field of machine tool fixture technology. It includes a positioning frame and a mounting plate. The edge of the mounting plate has a detection station and a machining station corresponding to two adjacent sets of blades. The detection station has a shape measurement mechanism, and the machining station has a clamping machining mechanism. The shape measurement mechanism, combined with a laser ranging probe, measures the shape of a standard blade. The clamping machining mechanism includes support cylinders distributed on the machining station, positioned below the blade at the machining station corresponding to the laser ranging probe. A machining motor is mounted above the blade at the machining station. The laser ranging probe measures the shape of the standard blade, the support cylinders in the clamping machining mechanism support and position the blade to be machined, and the machining motor, combined with the machining of the upper surface of the blade, achieves overall blade machining in conjunction with an attitude adjustment mechanism.

A canopy apparatus for protecting individuals from weather elements near a helicopter includes a pair of rotor blade pockets which mount to adjacent rotor blades of the helicopter at their free ends. A flexible sheet extends between the rotor blade pockets and toward a rotor head of the helicopter. A rotor head pocket is attached to the flexible sheet and is shaped to receive the rotor head of the helicopter. Additional straps are used to further secure the flexible sheet to the rotor blades. The canopy apparatus mounts to the helicopter such that the rotor blades may be rotated to adjust a position of the canopy apparatus with respect to a body of the helicopter.

The invention belongs to the field of helicopter rotor systems, and particularly relates to a high-reliability hydraulic damper. Comprising a blocking cover assembly (1), a barrel assembly (2), an oil supplementing one-way valve (3) and a piston rod assembly (7), the blocking cover assembly (1) is connected with the barrel assembly (2) through a flange plate, two working cavities (9) are formed in the two sides of a piston head of the piston rod assembly (7), oil return cavities (6) are formed in the blocking cover assembly (1) and the barrel assembly (2), the blocking cover assembly (1) and the barrel assembly (2) are connected with an oil supplementing channel through a transparent oil pipe (4), and the oil supplementing one-way valve (3) is of a three-way structure. The oil supplementing one-way valve (3) is arranged on the barrel assembly (2) and used for communicating the oil return cavity (6) with the two working cavities (9), and in a non-working state, the oil supplementing one-way valve (3) communicates the oil return cavity (6) with the two working cavities (9); in the working state, the oil return cavity (6) is directly communicated with the low-pressure cavity (9) and isolated from the high-pressure cavity (9), and therefore the oil supplementing function is achieved.

The invention provides a 5G NTN waveform anti-rotor-shielding method suitable for helicopter communication, and the method comprises the steps: setting a switching threshold value of an adaptive modulation code and a fountain code, and carrying out the switching according to the channel quality; when the channel quality is good, adaptive modulation and coding are adopted, channel feedback CQI is obtained, and according to the adaptive modulation and coding technology, a proper modulation mode and coding efficiency are selected, and uplink and downlink waveforms of a 5G NTN system conforming to a target modulation mode and a target code rate are constructed; and when the channel quality is poor, using a fountain code, taking the LDPC code under the 5G NTN specification as an outer code, and taking the LT code as an inner code, and constructing uplink and downlink waveforms of the 5G NTN system conforming to the target modulation mode and the target code rate. According to the method, the adaptive modulation coding technology and the Raptor fountain code are utilized, so that the communication quality is ensured under the condition that the helicopter rotor is shielded, and the rotor shielding resistance of helicopter communication is improved.

A vortex inducer is applied to the pressure face of a propulsion rotor blade. The vortex inducer defines an aft face located forward of the trailing edge of the blade. In operation, the vortex inducer generates a primary vortex behind its apex, which induces a counterrotating secondary vortex through viscous coupling at the trailing edge. The secondary vortex accelerates and deflects the suction face wake, enhancing overall blade performance. The system is applicable to propellers, helicopter rotors, prop-rotors, fan blades, and marine propellers.

The invention relates to the technical field of manual pressure supply structures of helicopter rotor braking systems, in particular to a self-locking control device and method with position feedback, the device comprises an upper limiting plate and a lower limiting plate, the upper limiting plate and the lower limiting plate are connected to form a limiting plate, and a mounting space is formed in the limiting plate; the front end of the handle is rotatably mounted in the mounting space of the limiting plate through a rotating mandrel, a connecting hole is formed in the middle of the handle and used for being connected with an actuator cylinder piston connecting rod in the rotor brake control assembly, the rear end of the handle is connected with the self-locking functional assembly, and the self-locking functional assembly can slide at the rear end of the handle. A cylinder and a limiting block are arranged at the front end and the rear end in the limiting plate respectively to limit rotation of the handle, a first limiting groove close to the cylinder and a second limiting groove close to the limiting block are formed in the front end and the rear end of the limiting plate respectively, and the self-locking functional assembly can be matched with the first limiting groove and the second limiting groove to achieve locking of the handle at two rotation limiting positions. A cam structure is arranged at the front end of the handle and matched with a microswitch arranged at the front end in the limiting plate, and a signal indicating that the brake device has pressure is output.

This invention discloses an intelligent, controllable, and self-healing honeycomb torsion rotor and its control method, belonging to the field of rotorcraft technology, applicable to helicopters, multi-rotor UAVs, electric vertical takeoff and landing aircraft, and any aircraft whose primary power source is a rotor. The rotor of this invention includes a rotor hub and at least two blades. The blades have a multi-layered composite structure, including: an aerodynamic skin, an intelligent honeycomb sandwich structure (composed of piezoelectric composite honeycomb cells and MFC actuators, embedded with bistable cells), an internal vortex tube heat exchange channel, a self-healing functional layer (microcapsules and / or dynamic covalent bond system), a surface-programmable electronic paper unit, and a control system. This invention utilizes the rotor's centrifugal force for passive air intake, and vortex tube heat exchange to achieve cold and hot separation; the hot air can be used for de-icing and forming a Coanda effect air cushion; active blade torsion and zero-power state maintenance are achieved through piezoelectric intelligent materials and a bistable structure; and the self-healing functional unit enables automatic healing of microcracks. This invention integrates active deformation, thermal management, structural self-repair, and air cushion lift enhancement, solving the shortcomings of existing helicopter rotor structures, such as complex structure, limited efficiency, difficult thermal management, and lack of self-repair capability.

The present application relates to a kind of helicopter rotor brake hydraulic systems, the hydraulic system is by hydraulic power source subsystem, control unit, pollution isolation unit, redundancy module, load;The hydraulic power source subsystem outputs high-pressure hydraulic energy source, passes through control unit control pressure state, delivery to the entrance of pollution isolation unit, pollution isolation unit passes on pressure to redundancy module, the redundancy module responds to the control instruction issued by cockpit and realizes the on-off of pressure transmission oil circuit to load.This application has double insurance design on the rotor brake supply pressure oil circuit, avoids in the air hydraulic pressure into rotor brake device causes non-instructive rotor brake;System is designed with buffer device, avoids hydraulic pressure impact when rotor brake;Set pollution isolation device, carry out pollution key isolation to rotor brake device, improve the pollution control ability of hydraulic system.When helicopter is parked for a long time, the rotor brake pressure of helicopter can be kept by manual mode.

The application belongs to the technical field of helicopter rotor system anti-icing, and particularly relates to a high-reliability blade anti-icing device and an assembling method thereof. The device comprises a nickel leading edge wrapping sheet and a heating pad, wherein the nickel leading edge wrapping sheet extends from 1R to the blade root along the span direction; the heating pad is arranged between the blade root and the nickel leading edge wrapping sheet and below the blade tip of the nickel leading edge wrapping sheet.

The invention relates to the technical field of helicopters, and discloses a double-cylinder oil-gas separation type helicopter blade shimmy damping device which comprises a first connecting ring assembly, a second connecting ring assembly, an oil storage cylinder and a working cylinder, the first connecting ring assembly is connected with the roots of helicopter blades, the second connecting ring assembly is connected with a helicopter hub, a working cavity is formed in the working cylinder, and the oil storage cylinder is connected with the working cavity. An oil storage cavity is formed in the oil storage cylinder, an air bag is arranged in the oil storage cavity, the working cavity and the oil storage cavity are filled with hydraulic oil, and the working cavity is communicated with the oil storage cavity through a bottom valve assembly. By adopting a double-cylinder oil-gas separation structure in which the working cylinder and the oil storage cylinder are matched with each other, compared with a traditional full-hydraulic single-cylinder shimmy damping device, the double-cylinder shimmy damping device has remarkable improvement, adopts a single-extension piston rod structure and is combined with a volume compensation mode of a built-in air bag of the oil storage cylinder, so that the axial mounting size of the shimmy damping device is greatly shortened, and the production efficiency is improved. And the design of a mounting interface is simplified, so that the mounting interface is easier to arrange and assemble in a limited space of a helicopter rotor system.

The present application relates to a kind of piezoelectric negative stiffness collaborative vibration-absorbing strut for helicopter rotor dynamic load, including upper strut and lower strut, and the vibration-absorbing component for vibration isolation between upper strut and lower strut;The vibration-absorbing component includes auxiliary ring for bearing static load, lever spring and disc spring for realizing providing negative stiffness, adjustable precompression are equipped in the auxiliary ring, mass for absorbing and dissipating vibration energy is equipped below the disc spring; Lever spring side is contacted with piezoelectric ceramic stack, piezoelectric ceramic stack connects multiple mode external circuit module, forms multiple mode piezoelectric shunt damping effect, while controlling multiple mode component of system;When vibration is larger, store electric energy, when vibration is smaller, switch to multiple mode RLC circuit dissipating mode, convert mechanical vibration energy into electric energy and consume by circuit.The present application couples negative stiffness dynamic low-frequency broadband vibration-absorbing and piezoelectric shunt multiple mode suppression, realizes the effect of high-efficiency vibration isolation.

The retractable hangar comprises two sets of symmetrical retractable sunroom structures which can be relatively moved to form a folded state for shielding the helicopter and an unfolded state for taking off and landing the helicopter; the symmetrical retractable sunroom structures are cooperatively designed with the lifting platform, the aerodynamic resonance problem of the traditional hangar under the action of strong wind is effectively avoided by lifting the helicopter rotor to 1.2-1.3 times the height above the top of the sunroom, and the stability of the platform lifting is ensured to meet the requirements of the building wind resistance specification by cooperating with the three sets of hydraulic lifting mechanisms in the equilateral triangle distribution. The diameter of the disc-shaped landing platform is optimized to be 2-2.5 times the diameter of the propeller, and the ground embedded installation groove design is matched, so that the equipment is flush with the ground when it is completely retracted, which not only meets the demand of maximizing the utilization rate of high space, but also builds a safe area meeting the aerodynamic characteristics of the helicopter taking off and landing through the sunroom unfolding spacing of 1-1.2 times the platform diameter.