298 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 field of helicopter rotor blade motion parameter measurement, and relates to a blade deformation measurement method based on rotation-following vision. The method comprises the steps that LED mark points are installed on the upper surfaces of the paddles; a follow-up rotating binocular stereoscopic vision system used for shooting the upper surfaces of the blades obliquely and downwards is mounted above the propeller hub; images of the LED mark points are collected through a rotation-following binocular stereoscopic vision system according to the test state; identifying pixel coordinates of the LED mark points in the image; according to an internal reference matrix, a rotation matrix and a translation matrix, obtained through pre-calibration, of the following-rotation binocular stereoscopic vision system, the pixel coordinates of the LED mark points are converted into three-dimensional coordinates under a coordinate system of the following-rotation binocular stereoscopic vision system; and converting the three-dimensional coordinates of the LED mark points in the coordinate system of the follow-up rotation binocular stereoscopic vision system into the three-dimensional coordinates of the LED mark points in the coordinate system of the propeller hub, thereby calculating the deformation of the propeller blade.

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 discloses matched bearing test equipment for a helicopter rotor system, and belongs to the technical field of bearing test, one side of a driven swing shaft is provided with a swing gear engaged with a driving fan, and the middle part of the driven swing shaft is provided with a detection plate; a gear ring is fixedly arranged on the inner side of the rotating shaft, and an inner gear is fixedly arranged in the middle of the gear ring; an overturning gear which is intermittently meshed with the gear ring and the inner gear is mounted on the inner side of the overturning shaft, and an overturning cylinder is fixedly arranged on the other side of the overturning shaft; toothed plates matched with each other are arranged on the detection plate in a sliding mode, a positioning frame is fixedly arranged at the bottom of each toothed plate, and bearings to be detected are installed in the overturning cylinders and the positioning frames. The to-be-tested bearing can be clamped and rotated through the overturning cylinder and the sealing cover, when the rotating motor drives the gear ring and the inner gear to rotate, the outer meshing teeth and the inner meshing teeth drive the overturning gear and the to-be-tested bearing to intermittently rotate in a reversing mode, and therefore the sealing effect in a high-strength rotating mode is simulated; the design improves the automation level and the detection efficiency.

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 invention discloses a helicopter rotor system dynamics simulation efficient solving method based on a three-dimensional entity unit, and belongs to the technical field of helicopter rotor system dynamics simulation analysis. Comprising the following steps: dividing a helicopter rotor geometric model into three-dimensional entity finite element models, and endowing unit attributes; and according to the finite element model, dividing co-rotation analysis components, and realizing effective butt joint among the components. Establishing a dynamic differential algebraic equation of the rotor system, and discretizing the dynamic differential algebraic equation; two key items in the efficient solving process are calculated, calculation of iteration increments is completed, and the dynamic time-history response of the rotor system is obtained. According to the thought of a component-level co-rotation coordinate method, extracting the elastic deformation of the rotor component, completing the stress tensor calculation of the rotor structure, and outputting a stress nephogram. According to the method, each iteration step linearized equation set is solved by adopting a Schel complement method, so that the solving efficiency of the large-scale linear equation set is improved; on the premise of keeping high precision, the dynamic simulation analysis of the finite element model of the large-scale rotor system can be efficiently completed.

The invention discloses a rotor wing shielding fading self-adaptive tracking method and device for helicopter satellite communication. The method specifically comprises the following steps: establishing a rotor shielding mathematical model of helicopter satellite communication, acquiring a signal after rotor shielding fading detected by a helicopter airborne communication terminal, and adding anti-noise filtering to improve a detection signal-to-noise ratio; a rotor shielding waveform is generated according to the measured rotor shielding fading parameters, coarse synchronization of the generated waveform and the detected waveform is carried out, and the system synchronization speed is increased; comparing a parameter error between the rotor shielding waveform generated by the local periodic pulse generator and the measured rotor shielding waveform to obtain an error detection result; and correcting the period, the phase and the width of a rotor shielding waveform generated by a local periodic pulse generator by using an error detection result, so that the rotor shielding waveform can adaptively track the change of rotor shielding. According to the invention, changes of shielding fading depth and shielding time of the helicopter rotor can be adaptively tracked, and the stability of satellite communication of the helicopter is improved.

The invention discloses a high subsonic velocity tilt rotor wing type and wing based on a backward blade theory, and belongs to the field of tilt rotor wing helicopter rotor wing type design. The camber of the front edge of the upper surface of the airfoil profile is 0.202-0.242, and the camber of the rear edge of the upper surface of the airfoil profile is 0.079-0.099; the camber of the front edge of the lower surface of the airfoil profile is 0.146-0.176, and the camber of the rear edge of the lower surface of the airfoil profile is 0.026-0.046; the slope of the rear edge of the upper surface of the airfoil profile is-0.09 to-0.11, and the slope of the rear edge of the lower surface of the airfoil profile is 0.0057 to 0.0077; the airfoil is arranged at the position of 60%-85% of the spanwise direction of a blade of the tilt-rotor helicopter, and can relieve the large angle-of-attack separation phenomenon at subsonic velocity, improve the stall characteristic, weaken the shock wave on the upper surface of the airfoil at transonic velocity, delay the divergence Mach number and improve the stability of the airfoil. Compared with an NACA64212 trailing blade airfoil profile, the trailing blade airfoil profile has better lift force, resistance and torque characteristics in subsonic velocity and transonic velocity ranges.

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 relates to the technical field of unmanned helicopter rotor brake systems, in particular to a rotor electric brake device which comprises an electromechanical actuating mechanism and a controller, the electromechanical actuating mechanism comprises a brake shell, a motor, a speed reducer, a lead screw, a lead screw nut, a brake pad and a force transducer, the motor is mounted on the brake shell, and the motor is mounted on the brake shell. The motor drives the speed reducer, the speed reducer is connected with the lead screw through a gear, the lead screw is sleeved with the lead screw nut, the lead screw nut is connected with the brake pads, the brake pads are installed on the two sides of the brake disc, the brake shell is further provided with the force measuring sensor, and the controller is connected with the motor and the force measuring sensor. According to the rotor wing electric brake device, the electromechanical actuating mechanism is controlled through the controller to achieve braking and brake loosening, the rotor wing electric brake device is suitable for the unmanned helicopter, traditional mechanical control type, hydraulic and pneumatic actuating type braking is replaced, full-electric automation and high-integration development of the unmanned helicopter is facilitated, the installation space is small, and maintenance is convenient.

The invention belongs to the field of helicopter tilting mechanisms, and relates to an electric helicopter rotor tilting mechanism. Comprising a motor support (1), a tilting rocker arm (2), a tilting rotating shaft assembly (7), a universal ring assembly (8) and a lead screw speed reduction structure (5), the lead screw speed reduction structure (5) comprises a tilting motor (6), a primary gear reducer and a lead screw (4), and the lead screw speed reduction structure (5) is connected with a machine body structure (3) through the universal joint assembly (8); the output end of the tilting motor (6) is connected with the input end of the primary gear reducer; the output end of the primary gear reducer is connected with one end of a lead screw (4); a rotor wing driving motor and a rotor wing are mounted on the motor support (1); one end of the tilting rocker arm (2) is fixedly connected with the motor support (1) through a bolt; the other end of the tilting rocker arm (2) is connected with the other end of the lead screw (4) through a joint bearing (17); and the motor support (1) is hinged with the machine body structure (3) through a tilting rotating shaft assembly (7).

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 rotor anti-icing and de-icing systems, and discloses a helicopter rotor anti-icing and de-icing system based on a 270V direct-current power supply and a control method, and the system comprises two icing detection sensors which are used for detecting icing signals; the icing signal processor is connected with the fly pipe computer, judges whether icing occurs or not according to the icing detection signal and transmits a result to the fly pipe computer; the flight pipe computer is connected with the atmospheric static temperature system and the man-machine interaction system, and the man-machine interaction system is used for displaying whether freezing occurs or not, receiving a control instruction of a pilot and transmitting the atmospheric static temperature and the control instruction to the anti-icing and deicing distribution box; the anti-icing and deicing distribution box is connected with the atmospheric temperature sensor, the 270 DC power supply and the main and tail rotor collector rings, the current temperature is determined according to the atmospheric temperature sensor and the atmospheric static temperature, and the deicing time is calculated according to the current temperature; and the main rotor heating assembly and the tail rotor heating assembly are respectively connected with the main rotor collector ring and the tail rotor collector ring and are used for heating and deicing by utilizing power supplies from the main rotor collector ring and the tail rotor collector ring.

The invention belongs to the field of vibration control, and relates to a self-adaptive torsion damper and an aircraft which are used for variable-speed helicopter rotors, torque transmission paths and transmission systems, and the damper is characterized in that the inner wall of an outer ring is connected with the outer wall of an inner ring of a center hole through N groups of centrifugal grooves; the center ring is installed on the rotating component through the center hole inner ring and rotates together with the rotating component. N is a natural number greater than or equal to 2; n groups of first-stage springs, N groups of second-stage springs, N groups of second-stage mass blocks and N groups of friction mass block assemblies are arranged in the N groups of centrifugal grooves respectively; one end of the secondary spring is connected with the outer wall of the inner ring of the center hole, and the other end of the secondary spring is connected with the inner side of the mass block; the outer ring is provided with an annular groove penetrating through the inner end face and the outer end face of the outer ring, and the inner friction ring is arranged in the annular groove, connected with the outer ring through a sliding unit and rotates along the annular groove in a sliding mode.

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 invention belongs to the field of helicopter strength design, and relates to a time domain correlation atlas analysis method for bending moments of different cross sections of a helicopter rotor shaft. The method comprises the following steps: extracting actual measurement data of rotor shaft bending moment measurement strain gauges on all cross sections under a specific working condition, and constructing a bending moment data matrix Mm of each cross section according to geometric arrangement of the rotor shaft bending moment measurement strain gauges; wherein each cross section has actually measured data with two orthogonal channels; forming a measuring point horizontal coordinate and a measuring point vertical coordinate of a Cartesian coordinate system by the data at the same moment in the two groups of measured data of the channels in the same direction of the different cross sections, or forming the measuring point horizontal coordinate and the measuring point vertical coordinate of the Cartesian coordinate system by the data at the same moment in the two groups of measured data of the channels in the different directions of the different cross sections; drawing positions of all measuring points on the time domain to obtain a time domain correlation graph; obtaining a helicopter rotor shaft bending moment correct measurement data time domain correlation graph normal form with the same working condition and the same number of rotor support arms; and comparing the time domain correlation graph with the time domain correlation graph normal form.

The invention relates to the technical field of aircrafts, in particular to a rotor wing speed reducer and a helicopter. The rotor wing speed reduction device comprises a rotor wing assembly, a power input assembly, a tail transmission assembly and a speed reduction assembly, the rotor wing assembly comprises a rotor wing shaft, the speed reduction assembly comprises a first non-orthogonal face gear and a second non-orthogonal face gear which are arranged on the rotor wing shaft, and the tooth surface diameter of the first non-orthogonal face gear is smaller than that of the second non-orthogonal face gear; the power input assembly comprises a first cylindrical gear connected with the second non-orthogonal face gear in a meshed mode, and the tail transmission assembly comprises a second cylindrical gear connected with the first non-orthogonal face gear in a meshed mode. The design requirement of any front inclination angle of the rotor shaft can be met, the transmission mode of the cylindrical gear and the face gear is adopted, the higher speed ratio of the whole machine is achieved, the transmission ratio is stable, and the power density is effectively improved. And pairing processing is not needed, so that the processing and maintenance cost is greatly reduced.

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.