201 results about "Tail rotor" patented technology

A rotorcraft with at least one main rotor and a fuselage, comprising: a tail boom connected to the fuselage, the tail boom extending from the fuselage toward a duct-type portion; a shroud provided at the duct-type portion and forming a transverse duct; at least one ducted tail rotor rotatably arranged in the transverse duct, wherein the at least one ducted tail rotor comprises a plurality of rotor blades mounted to a rotatable rotor hub; and a stator with a gearbox fairing mounted to the shroud for supporting the rotatable rotor hub in the transverse duct, wherein the stator comprises at least one rotatable stator blade connecting the gearbox fairing to the shroud.

The invention provides a component dynamic load estimation device and method considering rotor wing lower washing flow interference, the device comprises a full-aircraft module, and the full-aircraft module comprises a main rotor wing module used for providing rotor wing blade airfoil parameters, rotor wing blade aerodynamic parameters and rotor wing blade structure parameters; the tail rotor module is used for providing tail rotor blade airfoil parameters, tail rotor blade aerodynamic parameters and tail rotor blade structure parameters; the horizontal tail module is used for providing horizontal tail dynamic characteristic data; the fuselage module is used for providing fuselage aerodynamic characteristic data, overall parameters and rotor wing downwash velocity distribution; the balancing analysis module is used for calculating to obtain a tail rotor dynamic load or a horizontal tail dynamic load; the method can effectively guide the strength check and life evaluation work of each part of the helicopter, can effectively guide the weight reduction and life prolonging of each part of the helicopter, reduces the flight accident rate, and improves the overall performance of the helicopter.

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 provides a direct comprehensive modeling method for a helicopter tail rotor / tail beam coupled substructure, which comprises the following steps of: constructing a blade model, and dispersing the blade model to obtain a mass matrix, a rigidity matrix, a damping matrix and a generalized force vector expression of a single blade; establishing a structural kinetic equation of the tail beam; and finally, according to the displacement boundary condition of the joint of each paddle and the tail beam, converting the paddle structure dynamic equation, and assembling the rigidity matrix, the damping matrix, the mass matrix and the generalized force vector of each paddle after conversion with the rigidity matrix and the mass matrix of the tail beam respectively to obtain the structure dynamic equation of the tail rotor / tail beam coupling system. And establishing a structural dynamic model of the tail rotor / tail beam coupling system. The method is simple and reliable in modeling process and high in universality, is not limited by the distribution type of the tail rotor, can be applied to the non-uniformly distributed tail rotor, and provides a new thought for coupling analysis of the tail rotor / tail rotor and rotor wing / fuselage of the helicopter, dynamic characteristic solving of the tail rotor and analysis of stability.

The invention provides a multi-configuration helicopter quality evaluation simulator and a flight simulation method, and belongs to the field of flight simulators, the flight simulator comprises: a hardware system, including a six-degree-of-freedom motion platform, a joystick system, a plurality of computers and a display; and the software system comprises a flight simulation system, a comprehensive simulation management platform, a virtual central control console, a visual display system and a flight quality evaluation system. The simulator provided by the invention covers a traditional single-rotor helicopter with a tail rotor to an innovative coaxial rigid rotor helicopter, a tilt rotor aircraft, a tandem helicopter, an X-3 combined high-speed helicopter and the like, and supports two modes of offline and human in-loop to realize rapid flight simulation of various flight missions; flight quality evaluation can be completed by using simulation data in time, and reference suggestions are provided for flight quality specification formulation and flight task setting of the newly-developed configuration helicopter.

The invention discloses a helicopter ducted tail rotor noise reduction lining for low-frequency line spectrum noise suppression, and belongs to the technical field of helicopter ducted tail rotor noise reduction. Ducted tail rotor noise belongs to multi-line spectrum harmonic noise, and noise energy is reduced along with increase of frequency, that is, the noise energy is concentrated in first two-order harmonic waves. And on the other hand, the ducted tail rotor is low in rotating speed and small in blade number, so that the noise frequency is low. The micro-perforated plate composite folding back cavity is used as the sound absorption unit, and two sound absorption resonance frequencies of the sound absorption unit coincide with the first two orders of harmonic noise of the tail rotor through parameter design, so that the thickness of the whole sound absorption unit is reduced, and meanwhile, the low-frequency multi-line-spectrum noise reduction performance is achieved; the weight cost is reduced, and the tail rotor noise is effectively restrained.

One embodiment is a rotorcraft comprising a fuselage; an empennage at an aft end of the fuselage; a tail rotor connected to the empennage; and a tail rotor guard (TRG) around the tail rotor, the TRG shrouding to the tail rotor; and a light emitting diode (LED) lighting element integrated into at least one surface of the TRG.

The invention discloses a real helicopter aerodynamic characteristic test ground simulation method and device, and relates to the helicopter wind tunnel test field, and the device comprises a pedestal arranged on the ground, the pedestal is provided with a rotating disc, the rotating disc is provided with a balance mechanism, and the balance mechanism is provided with a three-point supported web support type mechanism; the real helicopter can be mounted on the belly support type mechanism; the strain bridges are arranged on the main transmission shaft and the tail transmission shaft. A real helicopter is installed on a full-size wind tunnel test section, and the aerodynamic performance of the helicopter is measured by adopting a balance mechanism and a strain bridge; according to the method, the aerodynamic performance assessment test is carried out under the working conditions of the rotor, the tail rotor and the whole helicopter, and on the basis, the aerodynamic performance income data of the helicopter are obtained through the optimization of the whole-helicopter aerodynamic layout, so that a basis is provided for the comprehensive optimization design of the whole-helicopter aerodynamic layout in model development.

The invention relates to the field of power systems, and discloses a helicopter hybrid power system without tail rotor transmission and a control method. The power system includes a power battery, an engine having a mechanical power output side and an aerodynamic power output side, a first and a second drive, a turbine, and a controller, either drive being provided as a motor / generator. Through multi-path coupling of mechanical energy, pneumatic energy and electric energy, dynamic distribution of power is achieved under different flight working conditions, the power requirement difference of the helicopter in different working states is met, and the power matching performance, the working stability and the flight state safety reliability are improved. According to the actual flight state, the controller intelligently adjusts power distribution to adapt to power requirements under different atmospheric densities at high / low altitudes, power matching of a main rotor and a tail rotor for driving a helicopter is met, and the working reliability of a power system is effectively improved.

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

The invention discloses a helicopter flight control performance determination method and device, equipment, a medium and a product, and relates to the field of helicopter flight control performance evaluation.The method comprises the steps that flight data of helicopter tail rotor fault flight are collected; calculating each index value by using an index model according to the flight data; the index model comprises a process index model and a transient index model; the process index model comprises a tracking index model, an airspeed keeping index model and a safety boundary index model; the transient index model comprises a landing state index model, a landing position index model and a controller delay index model; determining a comprehensive evaluation value of the flight control performance according to each index value and the weight after weight changing; according to the method and the device, the accuracy of helicopter flight control performance evaluation under the condition of the tail rotor fault is improved.

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 present invention provides an intelligent unmanned flight system with a signal shielding function, which relates to the technical field of weaponry and equipment. It includes a mother aircraft and multiple daughter drones. The daughter drones are equipped with radio shielding devices and are carried inside the mother aircraft. The daughter drones have a tail rotor, a middle front engine, a middle rear engine, and a head. The daughter drones achieve hovering and formation functions through the tail rotor; the daughter drones achieve dive acceleration functions through the middle rear engine; the daughter drones achieve recoil deceleration functions through the middle front engine; and the mother aircraft automatically disintegrates after reaching a predetermined position. The solution provided by the present invention uses the mother missile to carry a large number of daughter drones with electromagnetic shielding functions, and can land stably on the ground for radio shielding.

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.

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

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 utility model relates to the field of ornithopters, in particular to a pure electric vertical take-off and landing high-capacity aircraft with double foldable flapping wings, which comprises an aircraft body, a cockpit is arranged at the top of the front end of the aircraft body, double rows of flapping wings are arranged on two sides of the top of the aircraft body in a front-back manner, and a plurality of side rotors are arranged on two sides of the aircraft body and positioned below the flapping wings. A plurality of tail rotors are arranged at the tail end of the fuselage, and a walking device is arranged below the fuselage. The double-row flapping wing aircraft is reasonable in structure, large in capacity, large in space and high in loading capacity, the aircraft body is made of light materials such as titanium alloy and carbon fibers, the dead weight can be effectively reduced, the loading capacity can be effectively improved, the double-row flapping wings fully utilize aerodynamic force, multiple functions of flight are improved, the flight efficiency is improved, and the double-row flapping wings are folded, so that the double-row flapping wing aircraft is more convenient to fold and unfold.

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 present invention belongs to the field of helicopter and flight control technology, and discloses a helicopter comprising: a fuselage having a longitudinal axis, a main rotor arranged above the middle portion of the fuselage, a power system arranged inside the fuselage, a control system arranged inside the fuselage, and an electric-driven tail rotor arranged at the rear of the fuselage; the electric-driven tail rotor comprises: an electric drive system, a positive blade assembly, and a reverse blade assembly; the electric drive system comprises a right-side heading electric drive and a left-side heading electric drive; the positive blade assembly is mounted on the output shaft of the right-side heading electric drive, and the reverse blade assembly is mounted on the output shaft of the left-side heading electric drive; the control system is used to control the speed of the right-side heading electric drive and the left-side heading electric drive in the electric drive system according to a preset control algorithm. The technical solution disclosed in the present invention solves the technical problem of increasing the yaw-pitch coupling effect of the helicopter in the application of existing variable-speed electric-driven tail rotors.

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.