77 results about "Tail rotor" patented technology

The application provides a helicopter transmission test bench safety early warning method, including the following steps: compiling a user customized formula table, the user customized formula table including fault detection conditions, physical channel values and logical channel values; obtaining the customized formula corresponding to the fault detection conditions, physical channel values and logical channel values in the user customized formula table; obtaining the channel address of the customized formula. The application proposes a helicopter transmission test bench alarm method based on multiple detection conditions, wherein the fault detection conditions include conditions such as the rotation speed reaching the lower limit value, the rotor loading device being effective, the tail rotor loading device being effective, the accessory loading device being effective and the like. The alarm method can effectively monitor and alarm the transmission system test piece parameters under the limited conditions, so that the fault positioning is more effective and accurate, and some unnecessary alarm information can be prevented from occurring, so as to avoid unnecessary interference to the test operators.

The invention belongs to the technical field of synchronizers, and discloses a helicopter synchronizer and a using method thereof.The helicopter synchronizer comprises a driving device and a synchronous sleeve, a shifting fork is fixedly connected to the surface of the outer surface of the synchronous sleeve, a planet carrier is arranged on one side of the synchronous sleeve, a machine box shell is arranged on the other side of the synchronous sleeve, and the outer surface of the planet carrier is sleeved with combination teeth; the driving device is rotationally connected with the shifting fork through a second rotating shaft, the lower end of the synchronous sleeve is fixedly connected with a support, the lower end of the support is provided with a switch unit, the switch unit is electrically connected with the driving device, and the switch unit controls the synchronous sleeve to be meshed with or separated from the combination teeth, so that the helicopter tail rotor is driven to operate. The electric push rod is used for replacing hydraulic drive, hydraulic leakage is avoided, and the switch unit comprises a first switch and a second switch; electric signals are sent to the driving device through the switch unit, so that the working state of the driving device is controlled.

The present invention describes a detachable connection between a pitch control unit arm (5) and a pitch angle piece (8) of a blade holder (7) in a tail rotor head (3) of a tail rotor (1) of a rotary wing aircraft, said detachable connection comprising a pin connecting said pitch angle piece (8) attached to the blade holder (7) with said pitch control unit arm (5) of a pitch control unit (4), thereby reducing material wear and shortening the maintenance cycle. This is achieved because said pin is designed as a sliding pin (9) protruding between the pitch angle piece (8) and the pitch control unit arm (5), which in its length direction passes through a spherical bearing in the form of a ball (6) with a central ball through hole (61), wherein the ball (6) and the part of the sliding pin (9) are surrounded by an outer ring (60) of the spherical bearing, thus allowing a sliding and pivoting movement of the sliding pin (9) inside the spherical bearing.

An ultra-high-speed helicopter with stowed single-blade coaxial rotors is provided and relates to the technical field of aircrafts. The ultra-high-speed helicopter includes a fuselage. Wings and tails can be arranged on the fuselage, respectively. Tail blades can be arranged at a tail end of the fuselage. A rotor mechanism can be arranged on a top of the fuselage. The rotor mechanism includes rotors. The rotor mechanism can drive the rotors to be stored in a rotor storage cabin of the fuselage. When the helicopter flies at high speed, the rotors stop working, and tail blades provide propulsive force and the wings provide lift force, so that the defects in the aerodynamic characteristics of the rotors are avoided during high-speed flight. In addition, the rotors can be retracted and stowed in a fuselage to further reduce drag and increase the maximum flight speed.

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 application provides an aircraft tail rotor failure early warning method and device and a storage medium, and relates to the technical field of aircraft data processing. The method determines a dynamic threshold value based on current flight parameters of an aircraft and current tail rotor characteristic data in a wind tunnel test database, and determines a failure risk level of the tail rotor of the aircraft based on the dynamic threshold value and a preset risk division strategy. The dynamic threshold value is determined based on relevant parameters of the current flight state of the aircraft, is more in line with the flight state of the aircraft, can improve the rationality of the dynamic threshold value, and further improves the accuracy of the aircraft tail rotor failure early warning.

The disclosed propulsion apparatus according to various aspects of the subject technology may provide improved directional control of an aircraft, eliminating the need for a tail rotor. The propulsion apparatus may include a frame connected between a body of the aircraft and a main rotor having a torque effect on the body. The propulsion apparatus may also include a first motor mounted to the frame and configured to spin in a first rotational direction. The propulsion apparatus may further include a second motor mounted to the frame and configured to spin in a second rotational direction opposite the first rotational direction to create a force opposite the motion imparted to the body by the torque effect of powering the main rotor, wherein the second motor and the first motor drive the main rotor in the first rotational direction, and wherein the aircraft is steered around a vertical axis by adjusting the speed of the first motor and the second motor.

The present application belongs to the field of tail rotor hub design, and relates to a tail rotor central part test device. The tail rotor hub central part with special shape and complex load condition lacks fatigue test method and tooling design experience. The tail rotor central part test device has a central part fixing structure, a plurality of groups of pitching force loading structures, a flapping force loading structure based on a force transmission assembly and a centrifugal force loading structure. The central part fixing structure comprises a simulation main shaft fixed on a base. The pitching force loading structure comprises a pitching steel wire rope, a movable pulley and a pitching fork lug. The force transmission assembly comprises a centrifugal force joint, a flapping force joint, an elastic bearing and a fixing pin. The flapping force joint comprises a flange head at the front end, a cross rod at the middle and a vertical rod at the rear end. The flapping force loading structure comprises a flapping loading oil cylinder. The centrifugal force loading structure comprises a double fork lug, a centrifugal fork lug and a pair of centrifugal steel wire ropes. The device is simple in structure, convenient to operate and capable of meeting the test requirements of the central part of a certain type of helicopter tail rotor.

The invention discloses a bearing-free type tail rotor blade integral drilling positioning device and method, and relates to the technical field of positioning devices.The bearing-free type tail rotor blade integral drilling positioning device comprises a base, a beam positioning module and a blade positioning module, the beam positioning module is arranged at one end of the base, the blade positioning module is arranged at the other end of the base, and the beam positioning module is connected with a mounting hole in a flexible beam of a tail rotor blade; the blade positioning module is attached to the lower surface of a mold pressing blade of the tail blade for positioning, the tail blade is connected to the beam positioning module and the blade positioning module in a suspended mode, guide holes for drilling are formed in the blade positioning module, and the guide holes are matched according to the set drilling position and size of the mold pressing blade. The drilling positioning device can effectively guarantee drilling positioning in the overall state of tail rotor blade assembly, the structure is simple, operation is convenient, reliability is high, overall clamping is conducted after the tail rotor blade is assembled, multiple sizes of the tail rotor blade are positioned and then combined for use, and high positioning precision and low operation difficulty of all the sizes during drilling are effectively guaranteed.

The invention belongs to the field of helicopter structural strength design, and relates to a seesaw type tail rotor hub fatigue test method. The method comprises the steps that test pieces of a seesaw type tail rotor hub shaft assembly and a shaft sleeve assembly are pasted and calibrated; determining a fatigue test static load and a dynamic load; the static load is centrifugal force, and the dynamic load is flapping bending moment and shimmy bending moment; determining a flapping load and a shimmy load applied to the process joint on the blade false piece through loading debugging, and determining a centrifugal force applied to the process joint on the blade false piece; and carrying out a teeterboard type tail rotor hub fatigue test. According to the seesaw type tail rotor hub fatigue test method, the seesaw type tail rotor hub fatigue test can be guided to be completed, the fatigue dangerous part and the fatigue failure mode of the seesaw type tail rotor hub are determined, and the blank in the field of seesaw type tail rotor hub fatigue test in China is filled.

The invention provides a helicopter tail rotor and tail beam coupling modeling method based on a time-varying mixed interface modal synthesis method, and the method comprises the steps: firstly considering the rotation influence of a blade, and building a blade structure kinetic equation under a physical coordinate system; a blade structure kinetic equation under the modal coordinate system is obtained through modal coordinate transformation, tail beam modal information is obtained, and a tail beam structure kinetic equation under the modal coordinate system is established. Thirdly, a time-varying effect brought by rotation of the blade is considered, a quadratic coordinate transformation formula of modal speed and modal acceleration is obtained through conversion according to the displacement boundary condition between the blade and a tail beam position structure, and quadratic coordinate transformation is conducted on a blade structure kinetic equation under a modal coordinate system based on the quadratic coordinate transformation formula; and finally, assembling the blades and the tail beam according to a number-to-number seating principle to obtain a tail rotor and tail beam coupling time-varying structure kinetic equation under the modal coordinate system. The method is suitable for the coupling dynamics problem of the uniformly and non-uniformly arranged tail rotor and the tail beam, and the calculation efficiency in the modeling process is remarkably improved.

This utility model relates to the field of variable pitch structure technology and discloses a tail drive variable pitch structure for a plant protection UAV. It includes a fixed column, an adjustment mechanism fixedly connected inside the fixed column, a variable pitch mechanism inside the adjustment mechanism, and a hydraulic cylinder. A sliding column is fixedly connected to the driving end of the hydraulic cylinder. Two connecting rods are rotatably connected to the outside of the sliding column. A connecting block is fixedly connected to the outside of the connecting rods. A rotating plate is fixedly connected to the outside of the connecting block. A tail rotor is fixedly connected to the outside of the rotating plate. A drive assembly is fixedly connected to the outside of the fixed column. In this utility model, the hydraulic cylinder drives the sliding column to slide on the inner wall of the support column. The sliding column drives the connecting rods to rotate, the connecting rods drive the connecting block to move, and the connecting block drives the rotating plate to rotate on the inner wall of the fixed column. The tail rotor on the rotating plate is driven.

The application discloses a compound helicopter management system and management method with tilting rotors, and the management system comprises a power management subsystem and a thermal management subsystem, and the power management subsystem comprises a fuel engine, a start-launch integrated motor, a main speed reducer, a motor I matched with each tilting rotor, and a motor II matched with a tail rotor; wherein, a high-voltage three-phase power cable built in the start-launch integrated motor is connected to a high-voltage direct-current bus of the helicopter through a bidirectional start-launch controller; the high-voltage direct-current bus is also connected with a battery management module containing a power battery pack; and each motor is connected to the high-voltage direct-current bus through a corresponding multi-channel electric drive controller. The fuel engine can be started directly by the start-launch integrated motor with the main rotor, so that the clutch, the starting motor assembly and the starting gear disc can be cancelled, the structure is greatly simplified, the structural weight is reduced, and the reliability of the system is improved.

The invention belongs to the technical field of helicopter tail rotor blade testing, and particularly relates to a helicopter tail rotor blade fatigue test loading device which comprises a flange base, a loading mechanism and a loading mechanism. The tail rotor blade is arranged in the mounting mechanism; the interior of the fixing mechanism is connected with the other end of the tail rotor blade; the load testing mechanism is installed at the end, away from the installation mechanism, of the fixing mechanism, the load testing mechanism is arranged, a tail rotor blade test piece is tested through three different forces, centrifugal and aerodynamic loads borne by the helicopter tail rotor blade in the flight state are simulated, accurate and reliable test data are obtained, and the test efficiency is improved. Three load loading points of centrifugation, waving and shimmy are kept at the same point and have sufficient degree of freedom, so that the loaded distribution of the tail rotor blade better simulates a flight state, the angle of the mounting base is adjustable, and different mounting angle requirements can be met.

This application relates to the field of aircraft structural design technology, and more particularly to a folding tail boom structure for a single-rotor helicopter with a tail rotor, including a fuselage, a tail boom, and a folding assembly disposed at the connection between the fuselage and the tail boom. The folding assembly includes a front folding frame and a rear folding frame, with the front folding frame fixedly connected to the tail of the fuselage and the rear folding frame fixedly connected to the front end of the tail boom. The front and rear folding frames are rotatably connected by at least one set of shoulder screws, forming a folding rotating pair. This invention provides a porous oxide film layer formed by phosphoric acid anodizing at the connection interface between the front and rear folding frames and the fuselage / tail boom, combined with structural adhesive bonding, riveting mechanical fixation, and end face screw locking, forming an interconnected four-fold fixing structure. At the same time, the porous oxide film layer significantly improves the roughness and chemical activity of the bonding interface, greatly enhancing the bonding strength and effectively solving the problem of loosening of traditional quick-release pin connections under heavy loads.

This invention proposes an anti-torque assembly for helicopters, relating to the field of helicopter technology. It includes a fuselage, with a support shell fixed to one side. The support shell has a storage groove inside, and a connecting frame slides within the storage groove. A power mechanism for moving the connecting frame is located inside the support shell. This invention allows the connecting frame to slide within the storage groove via a moving rod and a connecting rod. Simultaneously, the cooperation between the sliding rod and the moving groove corrects the angle of movement of the connecting frame, allowing it to move obliquely upwards after exiting the storage groove. When the connecting frame is stored, a transmission sleeve can be retracted into a placement slot. This allows the connecting frame to be stored during helicopter placement, reducing the space occupied by the helicopter. Furthermore, a pull rod and a transmission plate can drive the transmission shaft to rotate within the transmission frame, simultaneously rotating the tail rotor blades, allowing adjustment of the tail rotor blade angle according to the required anti-torque force.

The invention belongs to the technical field of aerodynamic design and verification of helicopters, and particularly relates to a method for designing a safe course control instruction of a helicopter. Comprising the steps that S1, a relation curve of heading control and tail rotor torque is obtained through a tail rotor table test, the slope relation between a power coefficient Cp and heading control Lat is determined, and the power coefficient Cp corresponds to the tail rotor torque; s2, according to a course control limit value and the slope relationship between the power coefficient Cp and the course control Lat, determining the slope relationship between the power coefficient Cp and the course control Lat corresponding to the course control limit value; and S3, obtaining actually-measured course control, actually-measured tail rotor power and tail rotor power limitation, and calculating safe course control according to a slope relation between a power coefficient Cp corresponding to a course control limitation value and course control Lat.

The invention relates to the technical field of helicopter tail rotor system part machining, and discloses a drill jig for helicopter tail rotor system part fairing drilling, which comprises a clamp body, the upper end of the clamp body is provided with a pressing plate and a drill plate, the two sides of the pressing plate are both horizontally provided with third bolts, and the lower end of the drill plate is provided with second bolts. The third bolts penetrate through the pressing plate and extend into the clamp body, second bolts are vertically mounted at the ends of the two sides of the pressing plate, and the second bolts penetrate through the pressing plate and extend into the clamp body; pins are arranged on the two sides of the end, away from the pressing plate, in the drill plate. A first bolt is arranged in the middle of the end, away from the pressing plate, in the drill plate, penetrates through the drill plate and extends into the clamp body. Through the design of multi-direction pressing, precise positioning and guiding and convenient assembling and disassembling, the problems of fairing drilling shaking, hole position deviation and tedious operation are solved, and the machining precision and efficiency are guaranteed.

The invention discloses an aircraft combining a small unmanned aerial vehicle module and a large multi-axis module, and belongs to the technical field of aircrafts. The small unmanned aerial vehicle module is hinged to the upper portion of the large multi-axis module and is responsible for controlling the flight attitude, and flexible deflection is achieved by adjusting the rotating speed of an electric small rotor. The large multi-shaft module comprises an electric middle rotor wing, a flight control system and a battery pack, mainly provides lift force, supports various configurations such as four shafts and four propellers, and can be precisely fixed in height through a high-precision air pressure sensor. The large multi-axis module can be selectively provided with self-rotating rotors to improve the safety of forced landing, universal bearings are adopted at the hinged positions to optimize posture adjustment, and yaw control is achieved by adjusting the rotating speed of a tail rotor or a middle rotor. The aircraft can carry goods or personnel, is adaptive to a wireless / wired control mode, has the advantages of large lift force, flexible attitude control, simple structure and high safety, and is suitable for the scenes of goods transportation, low-altitude tourism, emergency rescue and the like.

The application relates to the technical field of flying saucers, and provides a three-dimensional integrated magnetic fluid driving flying saucer flying device, which comprises a central stationary cabin section serving as an internal core support and carrying platform of the flying saucer, an upper power assembly arranged above the central stationary cabin section, the upper power assembly comprising an upper saucer shell, an upper disc ultra-thin motor and an upper disc screw type disc driven to rotate by the upper disc ultra-thin motor, and a lower power assembly arranged below the central stationary cabin section. The coaxial reverse rotation design of the upper and lower power assemblies effectively offsets the spin torque generated by the rotor, realizes the balance of the body yaw without a tail rotor, and ensures the stability and safety of the central stationary cabin section during flight. In combination with the spiral flow tidal inlet groove on the upper disc surface and the lower central exhaust outlet, the overall aerodynamic layout is optimized, and the vortex lift and vertical reverse thrust of the flying saucer are obviously improved.

The application discloses a rotor propulsion system and application thereof, and belongs to the technical field of propulsion systems. The system adopts a double-motor counter-rotating structure, driving torque is superposed and output, and reverse torques are offset to each other, so that no driving reverse torque is generated to an external fuselage, and therefore no balancing device such as a tail rotor is needed. The system realizes unidirectional output through a overrunning clutch, prevents the rotor from generating a power generation effect on the driving device, and significantly improves reliability and safety. The specific structure comprises a shell, a first motor mechanism and a second motor mechanism fixed to the shell, the two are connected through meshing of parallel gears and rotate in opposite directions, a main shaft is axially fixed to the parallel gears, a speed reducer is fixed to the main shaft, and an output end of the speed reducer is connected to the overrunning clutch. The output ports of the first and second motor mechanisms are 180° rotationally symmetrical about the intersection point of the meshing center point and the parallel gear axis.

The present application relates to a rotor hub (10) for a tail rotor (6) of a rotary wing aircraft (1) comprising a plurality of tail rotor blades (11) as well as a central mounting (12) for fastening the plurality of tail rotor blades. The central mounting is designed to be driven by means of a tail rotor drive. The rotor hub further includes a plurality of blade holders (30), each of which connects a tail rotor blade (11) to the central mounting (12). The rotor hub comprises a cup-shaped hub body (13) including at least an annular outer structure (14) and an annular inner structure (15), both the annular outer structure and the annular inner structure including a plurality of first openings (16) and second openings (17), respectively. Each of the plurality of first and second openings is configured for receiving a blade holder (30), wherein at least the first openings of the annular outer structure each comprise a first reinforcing element (18) for increasing the rigidity of the rotor hub (10).

Tail rotor control system is described for helicopters. A pedal position sensor operable by a pilot yields greater tail rotor RPM relative to the main rotor RPM, giving the pilot increased control over the vehicle. This proves especially useful in certain situations, such as high altitude, where increasing tail rotor speed from main rotor speed can give a pilot increased maneuverability and stability.

The planetary motion rotor drive 1 is to be realized in a rotor speed alternating between lower and upper velocity periods -Vp, +Vp each 180° per revolution of the rotor and can be used as a rotor blade drive or centrifugal drive with high efficiency, consisting of an input coupling (2), at least one output coupling (6) and by means of a continuously variable electromagnetic torque manipulator (5) and a force-displacement storage device (4) and at least one differential gear (3) which has a basket (k), two drive wheels (a, e) and at least one compensating wheel. This ensures ideal conditions of force symmetry without unwanted forces, without increased noise, without swashplates and hydraulics, and without a tail propeller, with high flight stability, high airspeed, and lift.

An aircraft propulsion system is provided. The system may provide improved directional control of an aircraft, eliminating the need for a tail rotor. The system may include a frame connected between a body of the aircraft and a main rotor having a torque effect on the body. The system may also include a first motor mounted to the frame and configured to spin in a first rotational direction. The system may further include a second motor mounted to the frame and configured to spin in a second rotational direction opposite the first rotational direction to create a force opposite the motion imparted to the body by the torque effect of powering the main rotor. The second motor and the first motor drive the main rotor in the first rotational direction. The aircraft is steered around a vertical axis by adjusting the speed of the first motor and the second motor.

The invention provides an aircraft tail rotor failure early warning method and device and a storage medium, and relates to the technical field of aircraft data processing, and the method comprises the steps: determining a dynamic threshold value through the current flight parameters of an aircraft and the current tail rotor characteristic data in a wind tunnel test database, and carrying out the early warning of the failure of an aircraft tail rotor based on the dynamic threshold value and a preset risk division strategy; the dynamic threshold value is determined based on the relevant parameters of the current flight state of the aircraft, the flight state of the aircraft is better met, the reasonability of the dynamic threshold value can be improved, and the accuracy of aircraft tail rotor failure early warning is further improved.

A right front machine arm, a left front machine arm and a rear machine arm are connected to the right front portion, the left front portion and the rear portion of a machine body respectively, a right front motor and a right front propeller are sequentially connected to the front end of the right front machine arm, and a left front motor and a left front propeller are sequentially connected to the front end of the left front machine arm. The rear end of the rear arm is connected with a V-shaped dual-motor mounting seat with a transverse opening, a rear upper propeller is obliquely connected to the upper motor mounting seat, and a rear lower thrust propeller is obliquely connected to the lower motor mounting seat; the lift force of the right front propeller and the left front propeller controls rolling in a differential mode, the lift force of the right front propeller, the left front propeller, the rear upper propeller and the rear lower thrust propeller controls pitching in a differential mode, the lift force of the rear upper propeller and the rear lower thrust propeller controls the course in a differential mode, and the propellers for controlling the attitude only participate in two of the three kinds of control. The available accelerator stroke of a driving motor corresponding to the propeller is increased, the capacity of controlling pitching, rolling and heading is improved, and the wind resistance of the aircraft is enhanced.

The invention discloses a bearingless tail rotor aeroelastic dynamics multi-component coupling modeling method, which comprises the following steps of: performing tail rotor pneumatic coupling modeling under the action of rotor interference airflow, performing tail beam / tail rotor structure coupling modeling, and coupling a tail rotor pneumatic model and a tail beam elastic support tail rotor structure model considering rotor interference; and dynamic modeling and analysis of the tail rotor gas bomb in a complex dynamic environment are realized. The tail rotor blade aerodynamic modeling adopts an unsteady surface element method and a free wake model, the rotor interference modeling calculates the induction influence of the space where the tail rotor and the tail rotor wake are located based on the rotor wake, and a tail rotor / tail beam coupling structure model is established based on tail rotor hub freedom degree coordination. According to the method, coupling factors in helicopter tail rotor aeroelastic dynamics analysis are fully considered, high-confidence modeling of the actual physical state of the tail rotor is achieved under the condition that the calculation scale is controllable, and the reliability of tail rotor aeroelastic dynamics modeling and load prediction is improved; the method is particularly suitable for predicting and analyzing the tail rotor load of the single-rotor helicopter with the tail rotor.

The application belongs to the technical field of helicopter aerodynamic design analysis, and particularly relates to an aerodynamic characteristic analysis method considering rotor fuselage tail rotor interference. The method comprises the following steps: S1, modeling the rotor, the fuselage and the tail rotor in a three-dimensional modeling software to obtain a rotor model, a fuselage model and a tail rotor model; S2, performing grid division on the rotor model, the fuselage model and the tail rotor model according to a set grid division strategy to obtain a rotor grid model, a fuselage grid model and a tail rotor grid model; the set grid division strategy comprises a grid type, a calculation domain size, a grid number, a grid distribution and a boundary layer setting; S3, setting a multi-stage relative motion coordinate system, which comprises a rotating coordinate system, a flapping coordinate system and a pitch coordinate system, defining the rotation, flapping and pitch motion of the rotor blade in the rotating coordinate system, the flapping coordinate system and the pitch coordinate system respectively to obtain a rotor blade displacement equation; and S4, simulating a rotor fuselage tail rotor interference flow field according to the rotor blade displacement equation and the rotor grid model, the fuselage grid model and the tail rotor grid model, and obtaining the aerodynamic characteristic considering the rotor fuselage tail rotor interference according to the rotor fuselage tail rotor interference flow field.

The application provides a helicopter tail rotor flexible beam calibration device, the device comprises: a calibration platform fixed on the ground; a connecting block arranged on the calibration platform; wherein the connecting block comprises a first surface and a second surface, the first surface and the second surface are adjacent; a clamping tool for assembling a flexible beam test piece; wherein the assembled flexible beam test piece is fixed in the connecting block together with the clamping tool; meanwhile, the application also provides a helicopter tail rotor flexible beam calibration method; the calibration method provided by the application realizes the calibration of three kinds of loads of the flexible beam test piece, that is, the edgewise bending moment, the flapwise bending moment and the torsion moment, obtains the calibration coefficient of the corresponding load of the flexible beam test piece, and can be used for measuring and debugging the test load in the fatigue test process of the flexible beam test piece, and calculating the rigidity and deformation of the flexible beam; the calibration device provided by the application adopts the overall structure design of the test platform, is stable in rigidity, has a wide application range, is convenient to install and disassemble, and has no redundant structure.