118 results about "Tail rotor" patented technology

An aircraft includes a rotary wing comprising a main rotor and a tail rotor, a drive unit configured to rotate the rotary wing, an information acquirer for acquiring position information of an obstacle relative to the aircraft, a display, and a controller including at least one processor and at least one memory coupled to the processor. The processor executes processing including deriving a direction, a horizontal distance, and a vertical distance of the obstacle with respect to the aircraft, based on the position information of the obstacle, determining, for one or both of the main rotor and the tail rotor, whether the direction, the horizontal distance, and the vertical distance satisfy a first distance condition, and displaying an index corresponding to the vertical distance at a display position on the display when the first distance condition is satisfied, the display position being identified by the direction and the horizontal distance.

The invention discloses a four-degree-of-freedom helicopter tail rotor load measuring device in a wind tunnel, belongs to the technical field of helicopter tail rotor load measurement, and aims to solve the problem that an existing helicopter tail rotor wind tunnel test lacks Z-direction accurate displacement control. Comprising a tail rotor assembly and a Z-direction adjusting mechanism, the Z-direction adjusting mechanism comprises a barrel, a Z-direction bearing seat and a Z-direction motor mounting seat are arranged at the bottom of the barrel, a Z-direction lead screw is in running fit with the Z-direction bearing seat, an output shaft of a Z-direction driving motor is coaxially connected with the lower end of the Z-direction lead screw through a Z-direction speed reducer, a Z-direction nut is in threaded fit with the Z-direction lead screw, and a sliding sleeve is connected with the Z-direction nut. The upper end of the sliding sleeve is connected with the lower end of the tail rotor supporting rod, the middle of the tail rotor assembly is hinged to the upper end of the tail rotor supporting rod, the yaw adjusting mechanism is arranged on the rear side of the tail rotor supporting rod, the output end of the yaw adjusting mechanism is hinged to the rear portion of the tail rotor assembly, and the tail rotor stand column is fixed to the output end of the XY-direction adjusting mechanism. The requirement for Z-direction movement of the helicopter tail rotor can be met.

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 application belongs to the technical field of hydraulic systems, and discloses a helicopter hydraulic system joint simulation method based on multi-discipline, wherein step one is to acquire hydraulic system demand; step two is to analyze the hydraulic system demand to obtain a hydraulic system use case diagram; step three is to perform function analysis to obtain an activity diagram, a sequence diagram and a black box state diagram; step four is to perform system design synthesis based on the function analysis to obtain a white box state diagram and build a hydraulic system function logic model based on Rhapsody; step five is to build a hydraulic system physical performance model based on AMESim based on the hydraulic system physical function model based on Rhapsody; the cabin display panel of the hydraulic system is simulated in AMESim based on the sequence diagram; step six is to build a main tail rotor dynamics three-dimensional visual model based on CAT IA; step seven is to combine the hydraulic system function logic model, the complete hydraulic system physical prototype model and the main tail rotor dynamics three-dimensional visual model; and step eight is to set simulation conditions and perform joint simulation.

The application discloses a helicopter flexible beam actual load substitution analysis method, comprising the following steps: considering that the flexible beam root actual measured load data is effective in the initial flight test, using the effective time period corresponding to the effective flight test data of the flexible beam root actual measured load, performing spectrum analysis on the flexible beam root profile actual measured load and the airfoil section profile actual measured load on the same piece of tail rotor blade at preset interval time in the effective time period of the effective flight test data, so as to obtain the amplitude of the flexible beam root profile actual measured load and the airfoil section profile actual measured load corresponding to the tail rotor rotating speed frequency respectively; combining the amplitude of each time point in the effective time period to form two groups of data columns; a scatter plot of the two groups of data columns can be drawn, a plurality of correlation formulas of the two groups of data columns are obtained through linear fitting, the correlation formula with the optimal correlation between the airfoil section profile actual measured load and the flexible beam root profile actual measured load on the same piece of tail rotor blade is selected as the basis for calculating the flexible beam root profile actual measured load, and the flexible beam root profile actual measured load in the whole flight process can be obtained based on the airfoil section profile actual measured load of the tail rotor blade.

This invention discloses a machining device for a helicopter tail rotor control fork, wherein: a zero-point positioning system is used to achieve positioning and clamping between the zero-point positioning system subplate and the machine tool; a universal quick-change tooling is installed on the upper part of the zero-point positioning system subplate, the universal quick-change tooling includes a tooling plate, the upper surface of which is distributed with multiple fixing grooves; the tooling plate of the universal quick-change tooling is installed onto the zero-point positioning system subplate by means of connecting positioning pins engaging with connecting positioning holes opened on the surface of the zero-point positioning system subplate; a clamping special tooling includes a special tooling base plate, the bottom of which is provided with a connecting positioning block that can be engaged into the fixing groove on the universal quick-change tooling; the clamping special tooling also includes a pressure plate, a pressure plate nut, and a pressure plate bolt; a positioning base is installed on the special tooling base plate, the positioning base being provided with positioning screw holes for installing the helicopter tail rotor control fork.

The invention discloses a hybrid power coaxial reversal helicopter transmission system, which belongs to the field of gear transmission and comprises an engine, a generator, a motor, a reversing subsystem, a coaxial reversal subsystem, a tail transmission subsystem and the like. The engine and the generator are connected through the parallel shaft speed reducer and connected to the first reversing subsystem through the friction clutch, and the reversing subsystem is connected to the coaxial reversing subsystem through the duplicate gear. The coaxial reversal subsystem adopts coaxial face gear transmission and comprises an upper face gear, a lower face gear, a left input straight gear, a right input straight gear and an idle gear. The motor is connected to the second reversing subsystem through a parallel shaft speed reducer and then connected to the coaxial reverse subsystem through a duplicate gear. And the power of the engine and the motor is converged through the coaxial reversal subsystem and is supplied to the coaxial reversal rotor wing for flight. And meanwhile, the output power of the motor is connected to the tail rotor through the reversing subsystem. The coaxial face gear transmission is adopted, torque from different power sources can be effectively transmitted to the main rotor wing or the tail rotor wing according to requirements, and better torque distribution is provided.

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.

A multi-rotor aircraft with vertical and horizontal tail rotors is characterized in that an undercarriage is connected below a fuselage, a small wing-shaped platform, a medium-large motor and a medium-large rotor are sequentially connected above the gravity center of the fuselage, the fuselage on the right side of the small wing-shaped platform is connected with a right arm, and a right small motor mounting seat, a right small motor and a right small rotor are sequentially connected to the right end of the right arm; the portion, on the left side of the small wing-shaped platform, of the fuselage is connected with a left arm, a left small motor mounting base, a left small motor and a left small rotor are sequentially connected to the left end of the left arm, the rear portion of the fuselage is longitudinally connected with a rear arm, and a near-tail small motor mounting base, a near-tail small motor and a near-tail small rotor are sequentially connected to the middle section of the rear arm. The tail end small motor mounting seat, the tail end small motor and the tail end small rotor are sequentially connected to the rear end of the rear arm; the lift moment of the tail-end small rotor enables the rotation direction of the aircraft to be opposite to the reaction torque of the middle-large rotor enables the rotation direction of the aircraft, the lift change of the tail-end small rotor controls the course of the aircraft, and the aircraft has the advantage of being high in wind resistance and becomes a universal vertical lifting flight platform.

The application provides a test platform for distributed electric tail rotor of unmanned helicopter and a yaw test method, wherein the test platform comprises a main shaft device, a distributed electric tail rotor device and a test platform circuit hardware; the main shaft device is rotatable and is used for carrying a main body of a body of the unmanned helicopter; the distributed electric tail rotor device is installed on the main shaft device and is used for carrying the distributed electric tail rotor; the test platform provided by the application has simple mechanical structure and is easy to realize; the test platform is safe and convenient to use through tests on the test platform; the test platform has wide application range, greatly expands the existing single tail rotor test device, can meet the test requirements of the distributed electric tail rotor to the maximum extent, fills the blank in the single tail rotor test field, and reduces the research and development cost of the distributed electric tail rotor; the yaw test method truly reflects the data of a yaw channel of the unmanned helicopter, has comprehensiveness and verifiability, can solve the failure caused by the single tail rotor, and improves the safety and reliability of the helicopter.

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 invention relates to the field of machining clamps, and discloses a helicopter tail rotor system bearing fork type part machining clamp which comprises a base, a transition plate, a conical positioning shaft, a rhombic plug pin and a plurality of clamping assemblies. The transition plate is arranged at the bottom of the base; the conical positioning shaft is arranged in the middle of the upper end of the base; the rhombic bolt is inserted into the transition plate at the upper end of the base, penetrates through the base and extends into the transition plate; the multiple clamping assemblies are distributed at the upper end of the base in the circumferential direction. Reliable center positioning is achieved through stable supporting of the base and the transition plate, accurate matching of the conical positioning shaft and a center conical hole of the fork-shaped part, accurate angular positioning is achieved through a rhombic plug pin, the supporting arm is stably fixed through the multiple sets of circumferential clamping assemblies, the problems of positioning deviation, inaccurate angular direction and unreliable clamping in the prior art are solved, the machining stability and precision are guaranteed, and the machining efficiency is improved. And the process design requirements are met.

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.

A right small rotor is connected to the upper portion of the right end of a right arm, a right large rotor is connected to the upper portion of the middle section of the right arm, a left small rotor is connected to the upper portion of the left end of a left arm, a left large rotor is connected to the upper portion of the middle section of the left arm, and a horizontal tail rotor is connected to the upper end of a vertical tail arm. The vertical right tail rotor wing is connected to the right end of the right horizontal tail machine arm, the vertical left tail rotor wing is connected to the left end of the left horizontal tail machine arm, the lifting force of the right small rotor wing, the right large rotor wing, the left small rotor wing and the left large rotor wing is upward, the lifting force of the horizontal tail rotor wings is downward, the lifting force of the vertical right tail rotor wing is leftward, and the lifting force of the vertical left tail rotor wing is rightward. The lift force differential motion of the vertical right tail rotor wing and the vertical left tail rotor wing causes torque differential motion to control the course of the aircraft; pitching, rolling and heading of the aircraft are controlled through torque changes, and the aircraft is suitable for being combined with a fixed wing to form a vertical take-off and landing fixed wing aircraft, has the advantages of being high in wind resistance and long in endurance time and becomes a universal vertical take-off and landing flight platform.

The utility model relates to a length measuring tool. The utility model relates to a helicopter tail speed reducer connecting rod extension measuring scale, which mainly comprises a reference sheet, a large sleeve, a sliding rod with scales, a small sleeve with internal threads, an upper support arm with external threads and a foldable lower support arm, and is characterized in that the reference sheet is fixed at one end of the large sleeve; the sliding rod with the scales slides and is nested in the large sleeve, and the front end of the sliding rod with the scales is connected with the small sleeve with the internal thread; the upper support arm with the external thread and the small sleeve with the internal thread are connected through threads and are locked through the locking nut; and the upper support arm with the external thread is hinged with the foldable lower support arm. The measuring scale adopts a telescopic folding design, and through the design of the telescopic scale cylinder and the folding arms, the extension amount of the helicopter tail speed reducer connecting rod can be measured under the condition that only a screw cap on the outer side of a tail rotor hub is disassembled and assembled, and a tail rotor blade, the tail rotor hub and a collector ring are not disassembled and assembled, so that the workload is greatly reduced. The device is reasonable in design, simple in structure and high in practicability.

The mooring unmanned aerial vehicle comprises an unmanned aerial vehicle body, balance rotors are symmetrically arranged on the two sides of a fuselage of the unmanned aerial vehicle body respectively, a tail rotor is arranged on the side, adjacent to the balance rotors, of the fuselage, and the balance rotors and the tail rotor are connected with the fuselage through quick release mechanisms respectively. The balance rotors are arranged on the two sides of the unmanned aerial vehicle body, so that the lateral disturbance influence of the unmanned aerial vehicle is effectively counteracted, the attitude stability in a hovering mode is greatly improved, and the unmanned aerial vehicle can effectively cope with a complex airflow environment; meanwhile, a tail rotor wing is further arranged and is matched with the balance rotor wing, so that the height stability of the unmanned aerial vehicle in various flight attitudes is ensured, and the operation precision and the execution efficiency are effectively improved; the balance rotor wing and the tail rotor wing are connected and clamped with the fuselage of the unmanned aerial vehicle body through the quick disassembly mechanism, so that the balance rotor wing and the tail rotor wing can be quickly disassembled and assembled, the balance rotor wing and the tail rotor wing can be installed according to use requirements, and use is more flexible.

The invention discloses a helicopter tail rotor damping installation bolt fatigue test device and a loading method, the device comprises a bolt fixing device and a load loading device, the bolt fixing device comprises a bolt installation table and a swing support seat, the bolt installation table is used for installing and fixing a damping installation bolt, and the swing support seat is used for supporting the damping installation bolt. The swing supporting seat is used for achieving swing simulation of the damping mounting bolt, the load loading device comprises a tensile testing machine, an axial hydraulic cylinder, a steel wire rope, a connecting piece, a rod end knuckle bearing and a centripetal knuckle bearing, and the damping mounting bolt can apply multidirectional load phases to the damping mounting bolt under the action of the load loading device; meanwhile, based on a given alternating load calculation method, the calculated alternating load is input into a control system of the tensile testing machine and the axial hydraulic cylinder, and the fatigue test requirements of multi-axis loading and complex working conditions can be met, so that the problem that a traditional testing device adopts a single-axis load loading mode, and the testing efficiency is low is solved. And the actual load bearing situation of the damping mounting bolt cannot be accurately reproduced.

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.