238 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 helicopter rotor head structure comprises a propeller hub, a propeller clamp, a bearing and a variable pitch mechanism, the propeller hub is arranged at the top of a rotor main shaft, a connecting shaft is arranged in the propeller hub, the end of the connecting shaft extends out of the propeller hub, the end of the connecting shaft is sleeved with the bearing, and the end, away from blades, of the propeller clamp is sleeved with the bearing and rotationally connected with the connecting shaft; the bearings comprise an angular contact bearing, a thrust roller bearing and a deep groove ball bearing, and the angular contact bearing, the thrust roller bearing and the deep groove ball bearing are sequentially arranged in the length direction of the connecting shaft; and the variable-pitch mechanism is arranged on the rotor wing main shaft in a sleeving manner and is connected with the propeller clamp. The bearing has the technical effects that the angular contact bearing, the thrust roller bearing and the deep groove ball bearing are combined for use, so that the axial and radial bearing capacities of the whole bearing are obviously improved, the bearing is not easy to damage, and the whole service life is longer; and by arranging the variable pitch mechanism, the angle of the paddle can be flexibly adjusted, so that the lift force and the resistance generated by the rotor wing are adjusted.

The invention belongs to the technical field of helicopters, and relates to a helicopter rotor blade-vortex interference noise suppression effect analysis method and system. The method comprises the following steps: establishing a CFD model for solving a helicopter rotor paddle-vortex interference flow field; establishing a CSD model for solving the helicopter rotor paddle-vortex interference flow field; establishing a coupling model; carrying out helicopter rotor paddle-vortex interference flow field simulation; establishing an FW-H model, and performing BVI noise calculation on the rotor blade; establishing an HHC numerical calculation model; and evaluating the inhibition effect of the HHC numerical calculation model on the BVI noise. According to the method, the unsteady effect of the blade can be fully considered, the unsteady air load on the surface of the rotor blade can be accurately simulated, the blade elastic deformation and the aeroelastic coupling effect caused by the HHC numerical calculation model are further deeply considered, the actual dynamic behavior of the blade is simulated, and then the prediction precision of the HHC numerical calculation model on the BVI noise suppression effect is enhanced.

The invention relates to the field of helicopter wing anti-icing and deicing, in particular to a helicopter anti-icing and deicing system applying rotor wing kinetic energy recovery and an operation method. The helicopter anti-icing and deicing system applying rotor wing kinetic energy recovery comprises an electric heating deicing device, a kinetic energy recovery device and a helicopter rotor wing key assembly. The electric heating deicing device is composed of a high-resistivity metal heating wire and an icing detector, and the device transfers heat to the surface of the wing for deicing through heating of the high-resistivity heating wire. The kinetic energy recovery device is composed of an electric energy rectification, voltage stabilization and energy storage device, a permanent magnet and a magnetic induction coil. Kinetic energy generated when the rotor rotates is converted into electric energy through the electromagnetic induction technology and stored in a battery pack. The helicopter rotor key assembly is composed of a helicopter blade and a helicopter swash plate and is a key component of the helicopter rotor. According to the invention, the kinetic energy of the helicopter rotor is recovered to prevent and remove ice, so that the overall energy utilization efficiency is improved, and the environmental adaptability and safety of the system are also improved.

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 present application belongs to the field of helicopter rotor system strength design, and particularly relates to a structural characteristic stress optimization calculation method under multi-load different phase loading. The periodic load containing exact phase information is divided into analysis steps containing certain time interval points, a numerical fitting method is used to find the maximum stress of the structure in the period, and then the structural characteristic stress is obtained. The method can quickly and accurately extract the structural characteristic stress, avoid the phenomenon that the characteristic stress is too small due to unreasonable step setting, and has important significance for the strength design of the helicopter rotor system. Compared with the traditional different phase load loading, the method of loading all load maximums, the analysis is more accurate, and the stress change of the structure in the load loading period can be effectively simulated.

The present invention relates to the field of rotor test technology, and in particular to a helicopter rotor test bench, comprising a mounting frame, wherein the mounting frame is divided into two parts, a bottom plate and a top plate, the top of the mounting frame bottom plate is rotatably connected with two groups of waterproof covers, the two groups of waterproof covers together form a closed space to wrap the top of the mounting frame bottom plate, the top of the waterproof cover is fixedly connected with a vertical baffle, the side of the waterproof cover is fixedly connected with a side baffle, and further comprises: a cooling bin, the cooling bin is fixedly connected to the top of the mounting frame bottom plate and is located inside the waterproof cover, the top of the cooling bin is fixedly connected with a steam bin, when a nozzle moves to the top of a blade, after a movable pipe is connected with a water outlet pipe, water flows to the nozzle through the connecting pipe, and the nozzle sprays water onto a rotor, simulating the working state of the rotor on a rainy day, thereby obtaining more comprehensive data of the rotor, and at the same time testing the pressure resistance of the rotor during operation.

A metal component made of a base metal and a coating system thereon is characterized in that the coating system comprises a conductive layer on the base metal and a resin-based layer including conductive pigments on the conductive layer. The conductive pigments form electrically conductive 3D-networks in the resin, with the networks being randomly distributed in the resin. Further, a method for coating a metal component is disclosed, and may include depositing a conductive layer on a surface of the component, depositing a resin with electrically conductive pigments on the conductive layer and drying the resin. The coating system may be applied to metal components, including aerospace fasteners such as pins, bolts, collars, nuts and nut plates, and washers, as well as studs, latches, helicopter rotors, and landing gear structures.

The invention provides a device and a method for controlling folding and unfolding of a rotor wing and an oblique beam of a helicopter, and the device comprises a folding / unfolding electromagnetic valve which is used for controlling high-pressure hydraulic oil to flow into a folding / unfolding hydraulic system of the rotor wing and the oblique beam; the rotor brake electromagnetic valve is connected with the folding / unfolding electromagnetic valve, and the rotor brake electromagnetic valve is used for controlling the high-pressure hydraulic oil to flow into a rotor brake device through a rotor brake pressure supply port; the rotor wing folding electromagnetic valve is connected with the folding / unfolding electromagnetic valve, and the rotor wing folding electromagnetic valve is used for controlling high-pressure hydraulic oil to flow into a rotor wing folding pressure supply port; the rotor wing unfolding electromagnetic valve is connected with the folding / unfolding electromagnetic valve, and the rotor wing unfolding electromagnetic valve is used for controlling the high-pressure hydraulic oil to flow into a rotor wing unfolding pressure supply port; wherein the folding / unfolding electromagnetic valve and the rotor wing folding electromagnetic valve form a rotor wing folding pressure supply oil way; and the folding / unfolding electromagnetic valve and the rotor wing unfolding electromagnetic valve form a rotor wing unfolding pressure supply oil way.

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 invention relates to a helicopter rotor attitude and vibration control method and system. The control method comprises the following steps that displacement signals of a helicopter rotor hub, vibration signals of the hub and roots of blades and rotating speed signals of a rotor are collected in real time; the real-time value of the attitude angle of the rotor wing is calculated according to the displacement signal of the rotor wing propeller hub, and then the attitude angle error between the real-time value and the target value is calculated; and calculating a PID control quantity based on the attitude angle error and the vibration amplitude of the rotor, and further generating a control instruction to drive a steering engine to control the helicopter rotor to operate according to the target attitude and reduce the noise level of the rotor at the same time.

The invention discloses a shaft surface ultrasonic rolling-laser shock composite strengthening device and method. The device comprises a laser shock system, an ultrasonic rolling system, a part clamping system and a machining working platform. The laser shock system can realize laser shock and automatic replacement of an absorption layer; the ultrasonic rolling system can realize ultrasonic rolling processing; and the part clamping system can realize the rotary motion and the axial motion of the shaft part. By means of the device, full-automatic machining of ultrasonic rolling and laser shock on the surface of the shaft is achieved. By means of the device and method, laser shock and ultrasonic rolling are distributed on the two sides of the shaft piece, the ultrasonic-laser-workpiece space position analysis relation is clarified, and synchronous machining of ultrasonic rolling and laser shock is achieved. In addition, the relation among laser-ultrasonic-workpiece machining parameters is established, and efficient and high-quality machining of ultrasonic rolling-laser shock on the surfaces of shafts such as aero-engine spindles, helicopter rotor shafts and gas turbine rotors is achieved.

A six-arm rotor system for a heavy helicopter in rotor model wind tunnel tests, belonging to the field of heavy helicopter rotor systems. It includes a hub central component assembly, a blade grip assembly, a swashplate assembly, a damper, a pitch change rocker arm, a pitch change pull rod, and a rotating part. An upward flapping limiter is provided at the top of the hub central component, and a downward flapping limiter is installed on the rotor shaft and is arranged at the bottom of the hub central component. The blade grip assembly is installed on the hub central component. A swashplate assembly is sleeved on the rotor shaft below the downward flapping limiter. Both ends of the damper are connected to the blade grip assembly and the hub central component respectively. The pitch change rocker arm is installed on the blade grip assembly. Both ends of the pitch change pull rod are connected to the pitch change rocker arm and the swashplate assembly respectively. Both ends of the rotating part are connected to the upward flapping limiter and the swashplate assembly respectively. The present invention meets the requirements of the large takeoff weight of heavy helicopters and can conduct rotor model tests in a wind tunnel, improving the safety and reliability of the rotor system.

The invention relates to a piezoelectric type negative stiffness collaborative vibration absorption supporting rod for a helicopter rotor dynamic load. The piezoelectric type negative stiffness collaborative vibration absorption supporting rod comprises an upper supporting rod body, a lower supporting rod body and a vibration absorption assembly which is arranged between the upper supporting rod body and the lower supporting rod body and used for vibration isolation. The vibration absorption assembly comprises an auxiliary circular ring used for bearing a static load, a lever spring and a disc spring used for providing negative stiffness and adjustable pre-compression are arranged in the auxiliary circular ring, and a mass block used for absorbing and dissipating vibration energy is arranged below the disc spring; the two sides of the lever spring make contact with the piezoelectric ceramic stack, the piezoelectric ceramic stack is connected with the multi-mode external circuit module, the multi-mode piezoelectric shunt damping effect is formed, and meanwhile a plurality of mode components of the system are controlled. When vibration is large, electric energy is stored, and when vibration is small, the mode is switched to a multi-mode RLC circuit dissipation mode, and mechanical vibration energy is converted into electric energy which is consumed through the circuit. Negative stiffness dynamic low-frequency broadband vibration absorption and piezoelectric shunt multi-mode suppression are coupled, and the effect of efficient vibration isolation is achieved.

The invention belongs to the technical field of helicopter rotor tests, and discloses a telemetering system and method with TOP trigger signals and capable of achieving cascade synchronization. The system comprises a wireless module, a Beidou signal switching module, an acquisition module, a circuit board, a battery, an upper computer, a sensor assembly, a non-rotating table body and an equipment shell. The rotor system is arranged on the non-rotating table body, and the equipment shell is mounted above a propeller hub of the rotor system; the circuit board is arranged at the opening of the equipment shell; the battery is connected below the circuit board; a wireless module, a Beidou signal switching module and an acquisition module are connected above the circuit board; the sensor assembly is mounted on the rotor system and is connected with the acquisition module through a cable; the Beidou signal switching module is used for receiving Beidou signals; and the wireless module is used for communicating with an upper computer.

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 present invention discloses a method for extracting features of helicopter rotors from external radiation source radar based on flicker correlation. It includes the following steps: Step 1: Obtain the slow time dimension echo of the distance unit where the target is located; Step 2: Time-frequency analysis, pixel value accumulation and cross-correlation processing to determine the parity of the number of rotor blades; Step 3: Determine the reference frequency of the echo flicker; Step 4: Perform phase compensation on the echo; Step 5: Time-frequency analysis of the echo after phase compensation; Step 6: Save the center frequency of each flicker at each search node, calculate the Euclidean distance, find the minimum position, estimate the number of rotor blades and the initial phase of a single blade; Step 7: Verify the estimated value of the number of blades; Step 8: Estimate the blade length and the initial phase of all blades. The present invention overcomes the defects of the existing phase compensation method, which has a large amount of computation in the iterative traversal compensation process and is easy to fall into the local optimal solution; and has the advantage of being able to complete the feature extraction of helicopter targets.

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 application belongs to the field of helicopter strength design, and relates to a rotor shaft cross section moment time domain correlation graph analysis method of a helicopter. The method comprises the following steps: extracting measured data of rotor shaft moment measuring strain gauges on all cross sections under a specific working condition, constructing a moment data matrix M of each cross section according to the geometric arrangement of the rotor shaft moment measuring strain gauges m ; wherein each cross section has two-channel orthogonal measured data; forming data of the same time in two groups of measured data of the same direction channels of different cross sections into a measuring point horizontal coordinate and a vertical coordinate of a Cartesian coordinate system, or forming data of the same time in two groups of measured data of different direction channels of different cross sections into a measuring point horizontal coordinate and a vertical coordinate of a Cartesian coordinate system; drawing positions of all measuring points in the time domain to obtain a time domain correlation graph; obtaining a rotor shaft moment correct measurement data time domain correlation graph paradigm of a helicopter with the same working condition and the same number of rotor arms; and comparing the time domain correlation graph with the time domain correlation graph paradigm.

A multi-harmonic, multi-channel, global active vibration control method for helicopters involves constructing independent bandpass tracking filter banks to separate the harmonic components in the acceleration response error signal at each control point on the helicopter's fuselage. Each harmonic component is independently controlled by selecting a convergence factor. During the control process, the reference input signal is reconstructed while updating the weights of the Fx-LMS adaptive controller with global vibration minimization as the control objective. This method achieves significant attenuation of the acceleration response at all measurement points on the helicopter's fuselage under multi-harmonic interference from the helicopter's rotor blades, effectively improving the vibration level of the fuselage. The present invention separates and independently controls the fundamental and harmonic components of the vibration signal by constructing several independent adaptive filters, with global vibration minimization as the control objective, to update the controller weights.

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.

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 belongs to the technical field of helicopter rotor structures, and particularly relates to a bearingless tail rotor structure. Comprising M pull rod assemblies, a fork-shaped piece, a tail shaft, a clamping plate, M bolts, M nuts and M paddle assemblies, each paddle assembly comprises a flexible beam, a paddle, two supporting bearings, a crankshaft and a spherical hinge, and the two ends of each pull rod assembly are connected with the corresponding paddle and the fork-shaped piece respectively; one end of the flexible beam and the paddle are solidified into a whole through mould pressing, the other end of the flexible beam is forked into fork lugs, the left fork lug is connected with the tail shaft and the clamping plate through a left side bolt and a left side nut, and the right fork lug is connected with the tail shaft and the clamping plate through a right side bolt and a right side nut; a spherical hinge is clamped between adjacent flexible beam fork lugs, an upper supporting bearing and a lower supporting bearing which are the same are fixed to the end face of each paddle, the supporting bearings are connected through a crankshaft, and the spherical hinges penetrate through the middles of the crankshafts and are connected with middle bolts in a sleeved mode through single lug rings on the spherical hinges.

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 invention relates to the technical field of composite material structures, in particular to a composite material paddle pull rod and a forming die and method thereof, the composite material paddle pull rod comprises a rod body and two metal joints, and the two metal joints are fixedly connected to the two ends of the rod body; the rod body comprises an inner pipe, and the two metal connectors are symmetrically and fixedly connected to the two ends of the inner pipe. The inner pipe is coaxially sleeved with a middle pipe, and the two ends of the middle pipe extend out of the inner pipe and are fixedly connected to the metal connectors in a sleeving mode. The middle pipe is coaxially sleeved with an outer pipe, and the two ends of the outer pipe are flush with the two ends of the middle pipe. The diameter of the blade pull rod is not changed, the existing design is met, the connecting strength of the metal connector and the rod body made of the composite material is high, loosening is avoided in the fatigue test process, and the requirement of the helicopter blade pull rod for the fatigue performance is met.

The present invention discloses a three-servo fixed-axis variable-pitch rotor module, including a central axis, a rotary actuator, a servo mounting seat, a propeller hub, a propeller clamp, a servo, a servo rocker arm, a connecting rod, a tilting plate, a rotating tilting plate, a variable-pitch connecting rod, a propeller clamp connecting rod, a tilting plate extension rod, and a tilting plate limiter. This solution adopts a fixed-axis solution, which reduces the volume of the entire rotor system and avoids the stability, vibration, friction and other problems caused by the movement of the slender rotating shaft. The connecting rod structure and the synchronization structure of the variable-pitch control system are effectively simplified, while fully retaining the various variable-pitch characteristics of the helicopter rotor system. In addition, the present solution has a simple and compact structure and can be combined and designed as a standardized module to obtain a variety of different helicopter configurations. The present invention also discloses a variety of different types of helicopter configurations that adopt the three-servo fixed-axis variable-pitch rotor module.

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.