192 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 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 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 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 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 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.

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 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 invention provides a helicopter rotor control load acquisition method and device, and provides a calculation method for building a rotor control load analogy database of a reference helicopter based on a test flight actual measurement load and calculating rotor control loads of helicopters of the same series according to the analogy database for the first time, and the method is high in precision, high in credibility and high in reliability. And the structural design of a rotor control system of the helicopter in a scientific research stage can be supported.

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 invention belongs to the technical field of helicopter rotor systems and discloses an automatic tilter torque arm anti-torsion arm structure, comprising a rotating shaft assembly, a rod assembly, and a spherical bearing. One end of the rotating shaft assembly is a rotating shaft, and the other end is connected to the rod assembly. The rod assembly is provided with a light rod, and the spherical bearing can be tilted and moved up and down and sleeved on the light rod of the rod assembly. When used as an anti-torsion arm, the rotating shaft assembly is rotatably connected to the casing, and the outer ring of the spherical bearing is fixedly connected to the fixed ring assembly. When used as a torque arm, the rotating shaft assembly is rotatably connected to the rotor shaft, and the outer ring of the spherical bearing is fixedly connected to the dynamic ring assembly. The present invention functionally realizes the role of an automatic tilter anti-torsion arm / torsion arm, simplifies the structural composition of the anti-torsion arm assembly / torsion arm assembly, greatly reduces the weight of the assembly, and eliminates the influence of space limitations on the typical configuration torque arm anti-torsion arm structure.

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.

The invention relates to the technical field of helicopters, in particular to a helicopter rotor shaft dynamic sealing ring dismounting tool which comprises a grip, the top of the grip is fixedly connected with a reinforcing mechanism, the top of the reinforcing mechanism is provided with a pulling hook, the grip comprises a chassis, and the bottom of the chassis is fixedly connected with a grip body. A hydraulic push rod connector is fixedly connected to the top of the chassis, connecting rods are fixedly connected to the four edges of the top of the chassis, the reinforcing mechanism comprises a hydraulic push-pull assembly, a reinforcing assembly is arranged at the top of the hydraulic push-pull assembly, the pulling hook comprises a hinged disc, and a grabbing and clamping assembly is arranged at the bottom of the hinged disc. According to the helicopter rotor shaft dynamic sealing ring dismounting device, the grip, the reinforcing mechanism and the pulling hook are arranged, rapid and safe dismounting of a helicopter rotor shaft dynamic sealing ring is achieved, the problems that in a traditional dismounting mode, operation is complex, efficiency is low, and the sealing ring is prone to being damaged are effectively solved, and the dismounting process becomes simple and visual through the innovative mechanical structure design.