Helicopter

Abstract

The invention belongs to the technical field of helicopters and flight control, and discloses a helicopter which comprises a helicopter body with a longitudinal axis, a main rotor arranged above the middle part of the helicopter body, a power system arranged in the helicopter body, a control system arranged in the helicopter body and an electrically-driven tail rotor arranged at the rear part of the helicopter body, the electrically-driven tail rotor comprises an electrically-driven system, a forward blade assembly and a backward blade assembly. The electric drive system comprises a heading right side electric drive and a heading left side electric drive; a forward paddle assembly is installed on an output shaft of the course right side electric drive, and a backward paddle assembly is installed on an output shaft of the course left side electric drive. And the control system is used for controlling the rotating speeds of the heading right-side electric drive and the heading left-side electric drive in the electric drive system according to a preset control algorithm. According to the technical scheme, the technical problem that the yawing-pitching coupling effect of a helicopter is increased in application of an existing variable-speed configuration electric drive tail rotor is solved.

Description

Technical Field

[0001] The present invention belongs to the technical field of helicopters and flight control, and particularly relates to a helicopter. Background Art

[0002] Most of the existing traditional helicopter tail rotor systems adopt mechanical drive for drive control. This technical means has the characteristics of complex structure and high maintenance cost. For medium and small helicopters, the industry has proposed a technical means of using a direct-drive variable-speed tail rotor system with an electric motor, which directly drives the tail rotor to rotate through the electric motor, and controls the thrust or pull by adjusting the tail rotor speed, and then balances the torque with the main rotor torque to achieve anti-torque balance control of the helicopter.

[0003] At present, although the above-mentioned variable-speed tail rotor system using direct-drive electric motor means has been simplified in structure, during the speed change process, in addition to the significant change in aerodynamic pull, the torque will also change accordingly. This torque change may have an adverse impact on the pitch attitude of the helicopter in extreme cases, and will increase the yaw-pitch coupling effect of the helicopter, bringing challenges to flight control. Further explanatory, when the existing direct-drive variable-speed tail rotor system changes the speed, the corresponding change in torque may cause the pitch attitude of the helicopter to change, thereby increasing the yaw-pitch coupling effect. This coupling effect makes the flight control of the helicopter more complex, and requires precise control algorithms to compensate and offset this effect, which is very difficult and costly to implement; in addition, the increase in the coupling effect may affect the flight stability and maneuverability of the helicopter, and may even lead to flight accidents in some cases. In summary, in view of the technical problems existing in the application of the existing direct-drive variable-speed tail rotor system, it is urgent to develop a helicopter using a new electric-driven tail rotor. Summary of the Invention

[0004] The purpose of the present invention is to provide a helicopter, which solves the technical problem of increasing the yaw-pitch coupling effect in the application of the existing variable-speed configuration electric-driven tail rotor.

[0005] To achieve the above object, the present invention adopts the following technical solutions: The present invention provides a helicopter, including: a fuselage having a longitudinal axis, a main rotor disposed above the middle of the fuselage, a power system disposed inside the fuselage, a control system disposed inside the fuselage, and an electric-driven tail rotor disposed at the rear of the fuselage; wherein, The electric-driven tail rotor includes: an electric drive system, a forward blade assembly, and a reverse blade assembly; wherein, the electric drive system is fixedly installed at the rear of the fuselage, and the electric drive system includes a right-heading electric drive and a left-heading electric drive; the forward blade assembly is installed on the output shaft of the right-heading electric drive, and the reverse blade assembly is installed on the output shaft of the left-heading electric drive; the reverse blade assembly and the forward blade assembly are symmetrically arranged on both sides of the rear of the fuselage, and are used to generate aerodynamic pulling forces with the same direction and perpendicular to the longitudinal axis of the fuselage. The control system is used to control the rotational speeds of the right-heading electric drive and the left-heading electric drive in the electric drive system according to a preset control algorithm.

[0006] A further improvement of the technical solution of the present invention is that the central installation axes of the forward blade assembly and the reverse blade assembly are kept consistent.

[0007] A further improvement of the technical solution of the present invention is that both the right-heading electric drive and the left-heading electric drive are outer-rotor motors or inner-rotor motors.

[0008] A further improvement of the technical solution of the present invention is that in the step of the control system executing to control the rotational speeds of the right-heading electric drive and the left-heading electric drive in the electric drive system according to a preset control algorithm, When the helicopter performs hovering control, the rotational speeds of the forward blade assembly and the reverse blade assembly are kept consistent, and the anti-torque formed by the rotational motion of the main rotor is trimmed.

[0009] A further improvement of the technical solution of the present invention is that in the step of the control system executing to control the rotational speeds of the right-heading electric drive and the left-heading electric drive in the electric drive system according to a preset control algorithm, When the helicopter performs yaw control, the rotational speeds of the forward blade assembly and the reverse blade assembly are kept consistent, and an additional yaw moment is formed by simultaneously increasing the rotational speeds of the forward blade assembly and the reverse blade assembly.

[0010] A further improvement of the technical solution of the present invention is that in the process of forming an additional yaw moment by simultaneously increasing the rotational speeds of the forward blade assembly and the reverse blade assembly, the satisfied condition is: Ωt = (k1 + k2 + k3) · k · Ωm; In the formula, Ωt is the rotational speed of the tail rotor; k is the theoretically balanced value of the rotational speed of the main rotor and the rotational speed of the tail rotor; k1 is the influence factor of the rotational speed fluctuation of the main rotor; k2 is the influence factor of gust interference; k3 is the yaw command requirement; Ωm is the rotational speed of the main rotor.

[0011] A further improvement of the technical solution of the present invention lies in that, in the step of the control system executing to control the rotational speeds of the right-side electric drive and the left-side electric drive in the electric drive system according to a preset control algorithm, When the helicopter performs roll control, the positive blade assembly and the negative blade assembly are controlled by speed differential to preferentially balance the yaw-direction moment of the helicopter; then, an additional pitch moment is generated by differential fine-tuning to offset the pitch moment generated by the helicopter due to rolling.

[0012] A further improvement of the technical solution of the present invention lies in that, in the step of the control system executing to control the rotational speeds of the right-side electric drive and the left-side electric drive in the electric drive system according to a preset control algorithm, When the helicopter performs pitch control, the pulling direction of the main rotor is adjusted and controlled to make the helicopter perform pitch motion; in addition, the electric drive tail rotor is controlled by speed differential to compensate and control the moment in the pitch direction of the helicopter on the premise of preferentially balancing the yaw-direction moment of the helicopter.

[0013] A further improvement of the technical solution of the present invention lies in that, in the step of the control system executing to control the rotational speeds of the right-side electric drive and the left-side electric drive in the electric drive system according to a preset control algorithm, When only the right-side electric drive in the heading fails, the rotational speed of the negative blade assembly is increased by the left-side electric drive in the heading to generate aerodynamic pull to balance the yaw-direction moment of the helicopter; wherein, the additional pitch moment generated by the electric drive tail rotor is compensated and controlled by the control system controlling the main rotor; Or, when only the left-side electric drive in the heading fails, the rotational speed of the positive blade assembly is increased by the right-side electric drive in the heading to generate aerodynamic pull to balance the yaw-direction moment of the helicopter; wherein, the additional pitch moment generated by the electric drive tail rotor is compensated and controlled by the control system controlling the main rotor.

[0014] A further improvement of the technical solution of the present invention lies in that, in the step of the control system executing to control the rotational speeds of the right-side electric drive and the left-side electric drive in the electric drive system according to a preset control algorithm, When both the right-side electric drive and the left-side electric drive in the heading fail, the helicopter first reduces the total pitch of the main rotor to a preset low level, then performs sideslip flight by forward or lateral control, and finally performs autorotative descent.

[0015] Compared with the prior art, the present invention has the following beneficial effects: In the technical solution disclosed by the present invention, the main function of the fuselage with a longitudinal axis is to provide an installation space or interface for airborne equipment and mission payloads, maintain the theoretical aerodynamic shape, and maintain the integrity of its own structure; the main function of the main rotor arranged above the middle of the fuselage is to provide aerodynamic lift for the whole aircraft, be able to overcome the gravity of the whole aircraft to do work and achieve flight in the air, be able to realize the acceleration and deceleration of the whole aircraft by adjusting the lift size, and be able to realize the longitudinal and lateral movement of the attitude by adjusting the lift direction; the main function of the power system arranged inside the fuselage is to provide a power source for the whole aircraft, convert chemical energy into mechanical energy, and transmit it to the main rotor and the electric-driven tail rotor, and provide working energy for other airborne equipment and mission payloads; the main function of the control system arranged inside the fuselage is to provide control for the movement and mission execution of the whole aircraft, and realize functions such as attitude stability, positioning and navigation, path planning, and mission execution through closed-loop control; the main function of the electric-driven tail rotor arranged at the rear of the fuselage is to provide a trimming function for the heading attitude of the whole aircraft, and realize functions such as hovering and yaw movement by adjusting the tail rotor lift size. In the technical solution disclosed by the present invention, the positive blade assembly and the reverse blade assembly in the electric-driven tail rotor are symmetrically arranged on both sides of the rear of the fuselage. The electric drive on the right side of the heading drives the positive blade assembly, and the electric drive on the left side of the heading drives the reverse blade assembly. During normal operation and speed change, according to the torque balance principle in mechanics, the torques generated by the two electric drives on both sides can cancel each other out; this symmetrical structure and torque balance mechanism enable, during the speed change process, that when the torque on one side increases, the torque on the other side can be adjusted accordingly to balance, thereby reducing the interference with the pitch attitude of the fuselage and reducing the basic influence of the yaw-pitch coupling effect; explanatorily, because when the speed of one electric drive changes and causes a torque change, the other electric drive can cooperate through reasonable control to avoid the fuselage from having an undesired pitch change due to excessive unilateral torque.

[0016] In a further preferred solution of the present invention, the control system can independently control the speeds of the electric drive on the right side of the heading and the electric drive on the left side of the heading according to a preset control algorithm. According to the dynamics principle, the aerodynamic lift and torque generated by the blades are closely related to the electric drive speed. By independently adjusting the speeds of the two electric drives on both sides, the aerodynamic lift and torque sizes generated by the positive and reverse blade assemblies can be precisely adjusted. Exemplarily, when the helicopter needs to perform yaw control or other flight actions that cause the tail rotor speed to change, the control system can optimize and adjust the speeds of the electric drive on the right side of the heading and the electric drive on the left side of the heading according to actual requirements; for example, if it is detected that the pitch attitude change is about to occur due to the speed change, the control system can timely adjust the speed of one or both electric drives to generate an opposite torque for compensation, thereby reducing the influence of the yaw-pitch coupling effect. Description of the Drawings

[0017] To more clearly illustrate the technical solutions in the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art; obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0018] Figure 1 It is a schematic structural diagram of a helicopter in an embodiment of the present invention; Figure 2 It is a schematic diagram of an electric-driven tail rotor in a helicopter in an embodiment of the present invention; Figure 3 It is a schematic diagram of the principle of speed control in the electric-driven tail rotor; in the figure, r(t) is the main rotor speed, e(t) is the error signal, u(t) is the tail rotor speed command, and y(t) is the tail rotor speed; k is the theoretically balanced value between the main rotor speed and the tail rotor speed, k1 is the influence factor of the main rotor speed fluctuation, k2 is the influence factor of the gust interference, and k3 is the yaw command requirement; The explanations of the reference numerals in the figure are as follows: 1, fuselage; 2, main rotor; 3, power system; 4, control system; 5, electric-driven tail rotor; 6, positive blade assembly; 7, counter-rotating blade assembly; 8, electric drive system. Detailed implementation manners

[0019] To make the objectives, technical solutions, and advantages of the present invention clearer, the following will clearly and completely describe the technical solutions in the present invention in conjunction with the drawings in the embodiments of the present invention; obviously, the described embodiments of the technical solutions are some embodiments of the present invention, not all of the embodiments.

[0020] Based on the technical solutions disclosed in the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention. In addition, the terms "include" and "have" and any variations thereof are intended to cover non-exclusive inclusion. For example, a process, method, system, product, or device that includes a series of steps or units does not necessarily have to be limited to those clearly listed steps or units, but may include other steps or units that are not clearly listed or are inherent to these processes, methods, products, or devices.

[0021] Please refer to Figures 1 to 3 , a helicopter provided in an embodiment of the present invention includes: a fuselage having a longitudinal axis, a main rotor 2 provided above the middle of the fuselage, a power system 3 provided inside the fuselage, a control system 4 provided inside the fuselage, and an electric-driven tail rotor 5 provided at the rear of the fuselage; Among them, the electric-driven tail rotor 5 specifically includes: an electric drive system 8, a forward rotor blade assembly 6, and a reverse rotor blade assembly 7. Among them, the electric drive system 8 is fixedly installed at the rear of the fuselage (in an exemplary technical solution, it can be installed at a preset position on the tail boom). The electric drive system 8 includes: a right-heading electric drive and a left-heading electric drive. The forward rotor blade assembly 6 is installed on the output shaft of the right-heading electric drive, and the reverse rotor blade assembly 7 is installed on the output shaft of the left-heading electric drive (specifically explained, the reverse rotor blade assembly 7 and the forward rotor blade assembly 6 are symmetrically arranged on the left and right sides of the rear of the fuselage, and the directions of the aerodynamic forces generated by them both point in the same direction, perpendicular to the longitudinal axis of the fuselage, mainly used to balance the anti-torque generated by the rotation of the main rotor 2; in an exemplary technical solution, for the forward rotor blade assembly 6, when observed from the right-heading side, it rotates counterclockwise, and the direction of the aerodynamic force generated points to the left-heading side; for the reverse rotor blade assembly 7, when observed from the right-heading side, it rotates clockwise, and the direction of the aerodynamic force generated also points to the left-heading side); the control system 4 is used to control the rotation conditions of the right-heading electric drive and the left-heading electric drive according to a preset control algorithm; in a further preferably specific technical solution, the central installation axes of the forward rotor blade assembly 6 and the reverse rotor blade assembly 7 are kept consistent.

[0022] In the technical solution provided by the embodiment of the present invention, the fuselage is the main frame of the helicopter, providing an installation foundation and structural support for other components, and ensuring the stability and strength of the entire helicopter. Specifically, by way of example, a load compartment is usually designed inside the fuselage for carrying equipment or goods required for tasks, such as sensors, communication equipment, cameras, cargo racks, etc.; the fuselage shell can protect key components such as the internal power system from the damage of the external environment (such as wind, rain, sand and dust, etc.), ensuring the normal operation of the helicopter.

[0023] In the technical solution provided by the embodiment of the present invention, the main rotor is the main lift source of the helicopter. By rotating at high speed, the air flow rate above the wing increases, and the pressure decreases, thereby generating an upward lift force, enabling the helicopter to take off and land vertically and hover. By changing the pitch of the main rotor (i.e., the inclination angle of the rotor blade), the flight attitudes such as pitch, roll, and yaw of the helicopter can be controlled; for example, increasing the leading-edge pitch can make the helicopter pitch up, and increasing the trailing-edge pitch can make the helicopter pitch down. When flying forward, the main rotor generates a forward thrust through cyclic pitch control (i.e., changing the pitch of the rotor blade in different azimuths), enabling the helicopter to fly horizontally.

[0024] In the technical solution provided by the embodiment of the present invention, the power system provides energy for the main rotor and the electric-driven tail rotor, usually using an electric or fuel power system, and the specific selection depends on the design requirements and application scenarios of the helicopter. The power system needs to have a power regulation function to adapt to the changes in power requirements in different flight phases (such as takeoff, hover, cruise, landing).

[0025] In the technical solution provided by the embodiment of the present invention, while the main rotor generates lift, it also generates a counter-torque that causes the fuselage to rotate. The electric-driven tail rotor balances this counter-torque by generating a thrust in the opposite direction of the counter-torque to maintain the stable flight attitude of the helicopter. By changing the rotational speed of the electric-driven tail rotor, the heading of the helicopter can be controlled. For example, increasing the rotational speed of the tail rotor can cause the helicopter to yaw to one side, and decreasing the rotational speed can cause the helicopter to yaw to the other side. The electric-driven tail rotor has the characteristics of fast response and precise control, which can significantly improve the maneuverability and flexibility of the helicopter, enabling it to better adapt to the complex and changing flight environment. Further explanatorily, in the embodiment of the present invention, a new electric-driven tail rotor is provided, which can not only reduce the yaw-pitch coupling effect of the variable-speed tail rotor, but also shorten the length of the helicopter tail boom, improving the spatial efficiency of the helicopter structure.

[0026] In the technical solution disclosed in the embodiment of the present invention, the preset control algorithm includes: when the helicopter hovers, the rotational speeds of the positive blade assembly 6 and the counter blade assembly 7 are kept consistent to balance the counter-torque formed by the rotation of the trimming main rotor 2; when the helicopter performs a left yaw control, the rotational speeds of the positive blade assembly 6 and the counter blade assembly 7 are kept consistent and implemented by increasing the motor rotational speed. Explanatorily, when the helicopter hovers or performs a yaw control, the rotational speeds of the positive blade assembly 6 and the counter blade assembly 7 are kept consistent to balance the torque of the main rotor 2 or form an additional yaw moment without generating an additional pitch moment.

[0027] In the technical solution of the embodiment of the present invention, the preset control algorithm is used to precisely control the rotational speeds of the forward blade assembly and the reverse blade assembly in the electric drive tail rotor of the helicopter, so as to achieve stable flight of the helicopter in different flight states (hovering, left yaw control), avoid generating additional pitching moments at the same time, reduce the yaw-pitch coupling effect, and improve the stability and reliability of flight control. Explanatorily, in the existing electric drive tail rotor solutions, only one blade assembly is controlled, and the anti-torque formed by the rotation of the main rotor of the helicopter is trimmed by adjusting the rotational speed of the tail rotor; in the existing technical solutions, during the hovering or yaw control of the helicopter, during actual flight, the steady-state rotational speed of the main rotor will fluctuate within a certain range (usually within the range of 3% - 5%), and there will also be a certain wind disturbance in the external environment. Therefore, in the hovering / yaw state, the existing electric drive tail rotor solutions will adjust and control the rotational speed of the tail rotor according to the fluctuation range of the main rotor rotational speed, the magnitude of the wind disturbance, and the yaw angle requirement. However, with the change of the tail rotor rotational speed, an additional pitching moment is also brought to the whole machine, which further affects the pitching attitude control of the whole machine. Differently, in the technical solution disclosed in the embodiment of the present invention, a new electric drive tail rotor is designed, which trims the anti-torque of the main rotor of the helicopter by controlling the forward and reverse blade assemblies; among them, when the main rotor rotational speed fluctuates, there is a gust of wind interference, or yaw control is performed, the forward and reverse blade assemblies synchronously adjust and control their rotational speed changes through the electric drive system. They only change the magnitude of the aerodynamic pull, and the direction of the generated pull is always perpendicular to the longitudinal axis of the fuselage. The additional pitching moments generated by a single blade assembly are equal in magnitude and opposite in direction, and cancel each other out when superimposed, and will not bring additional pitching moments to the whole machine.

[0028] In a further preferred technical solution of the embodiment of the present invention, the rotational speed of the electric drive tail rotor is linearly related to the rotational speed of the main rotor, the wind disturbance factor, and the yaw command, and the condition satisfied is: Ωt = (k1 + k2 + k3) · k · Ωm, and closed-loop control is performed; preferably, an exemplary PID controller is used for control; In the formula, k is the theoretically trimmed value of the rotational speed of the main rotor and the tail rotor, which can be obtained through theoretical calculation; k1 is the influence factor of the main rotor rotational speed fluctuation, which can be obtained through experiments; k2 is the influence factor of the gust of wind interference, which can be corrected through experiments; k3 is the yaw command requirement, which can be input according to actual requirements; Ωt is the rotational speed of the tail rotor, and Ωm is the rotational speed of the main rotor.

[0029] As a preferred technical solution of the embodiment of the present invention, the preset control algorithm further includes: when the helicopter performs roll control: the forward blade assembly and the reverse blade assembly are controlled by speed differential to preferentially trim the yaw-direction moment of the helicopter, and then through differential fine-tuning, an additional pitching moment is generated to offset the pitching moment generated by the helicopter due to roll.

[0030] In the embodiments of the present invention, when the helicopter performs roll control, the fuselage rotates around its longitudinal axis. This action not only changes the roll attitude of the helicopter but also may trigger a series of complex moment changes. Among them, the changes in the yaw moment and pitch moment are particularly crucial. The change in the yaw moment may cause the helicopter to deviate from the predetermined course, while the change in the pitch moment may cause the helicopter to assume a nose-up or nose-down attitude, affecting flight stability and controllability. When the helicopter performs roll control, the preset control algorithm controls the positive blade assembly and the negative blade assembly through speed differential control, that is, adjusts the speeds of the electric drive on the right side of the course and the electric drive on the left side of the course respectively, so that there is a difference in the speeds of the positive and negative blade assemblies.

[0031] Explanatorily, due to the opposite rotation directions of the positive and negative blade assemblies, when their speeds are different, torques of different magnitudes and directions will be generated. By reasonably controlling the speed difference, the yaw moment of the helicopter can be preferentially balanced. For example, when the helicopter rolls to the right, a torque that causes the helicopter to yaw to the left may be generated. At this time, by adjusting the speed difference between the positive and negative blade assemblies, the yaw torque generated by the tail rotor is offset against this adverse yaw torque, so as to maintain the yaw stability of the helicopter and avoid course deviation caused by rolling. When the helicopter performs roll control, the change in the fuselage attitude will destroy the original aerodynamic balance, resulting in a change in the pitch moment. For example, when the helicopter rolls to the right, the aerodynamic force distribution generated by the main rotor changes, which may cause the helicopter to have a nose-down tendency and generate an additional pitch moment. If not controlled, it will affect the pitch attitude stability of the helicopter. After preferentially balancing the yaw moment, by further finely adjusting the speed difference between the positive and negative blade assemblies, an additional pitch moment can be generated. This additional pitch moment is opposite in direction and equal in magnitude to the pitch moment generated by the helicopter due to rolling, so as to accurately offset the pitch influence brought by rolling. For example, if rolling causes the helicopter to have a nose-up tendency, then the additional pitch moment generated by differential fine adjustment will cause the helicopter to have a nose-down tendency, and the two offset each other, keeping the helicopter in a stable pitch attitude.

[0032] Explanatorily, for the existing electric-driven tail rotor solution, it only adjusts and controls the pulling force direction of the main rotor to make the helicopter perform a rolling motion. However, with the rolling motion of the helicopter, since the main rotor of the helicopter is always in a rotating state, the superposition of the two rotational motions will generate an additional pitching moment in the pitching direction. And the single-blade assembly tail rotor usually only balances the yaw moment, so it cannot eliminate the pitching moment generated by the helicopter due to rolling. In the new technical solution of the embodiment of the present invention, the positive and negative blade assemblies can be controlled by speed differential. First, the yaw moment of the helicopter is preferentially balanced, and then through differential fine-tuning, an additional pitching moment is also generated to offset the pitching moment generated by the helicopter due to rolling, thereby reducing the roll-pitch coupling effect of the helicopter and solving the defects existing in the prior art solution. In summary, the present invention preferentially balances the yaw moment and offsets the pitching moment generated by rolling. This control strategy can effectively reduce the attitude fluctuation of the helicopter during rolling and improve the flight stability. By controlling the rotational speeds of the positive and negative blade assemblies through differential, the balance of the yaw and pitching moments is achieved, without the need for additional complex control mechanisms or algorithms, simplifying the control logic. By preferentially balancing the yaw moment and then fine-tuning to generate an offset pitching moment, the coupling effect between yaw and pitch is reduced, making the flight control of the helicopter more stable and precise.

[0033] As a preferred technical solution of the embodiment of the present invention, the preset control algorithm further includes: when the helicopter performs pitching control: mainly adjusting and controlling the pulling force direction of the main rotor 2 to make the helicopter perform a pitching motion, and also controlling the electric-driven tail rotor 5 through differential. On the premise of preferentially balancing the yaw moment of the helicopter, the moment in the pitching direction of the helicopter is compensated and controlled.

[0034] During the flight of a helicopter, pitch control is one of the key flight actions. It allows the helicopter to achieve head-up or head-down attitude changes to meet different flight requirements, such as climbing, descending or adjusting the flight trajectory. However, pitch control does not exist in isolation. When a helicopter is performing pitch motion, the change in the thrust direction of the main rotor will not only affect the pitch attitude, but may also cause a change in the yaw torque, causing the helicopter to deviate from the predetermined heading. At the same time, other factors may also cause additional torque disturbances in the pitch direction, affecting the stability of the pitch attitude. The existing electric drive tail rotor solution only controls the thrust direction of the main rotor to make the helicopter pitch, while the tail rotor maintains the original motion balance state. In the new technical solution disclosed in the embodiment of the present invention, it can mainly adjust and control the pulling direction of the main rotor to make the helicopter pitch. It can also control the tail rotor through differential control to compensate for the moment in the pitch direction of the helicopter under the premise of giving priority to balancing the moment in the yaw direction of the helicopter. By optimizing the control of the yaw and pitch directions, the new solution helps to improve the overall flight performance of the helicopter, which may include increasing the flight speed, increasing the range, improving the flight stability and other aspects.

[0035] As a preferred technical solution of an embodiment of the present invention, the preset control algorithm also includes: when the electric drive system in the electric drive tail rotor fails but does not completely fail, the heading right electric drive or the heading left electric drive on one side can be retained, and the blade assembly is driven to generate aerodynamic thrust by increasing the rotation speed to balance the yaw direction torque of the helicopter. The additional pitch torque generated by the tail rotor is compensated and controlled by the helicopter main control system, and the current electric drive tail rotor fault information is fed back.

[0036] Explanatoryally, since the electric drive system is provided with a right-side heading electric drive and a left-side heading electric drive, when one side fails, the electric drive on the other side can continue to work; by way of example, assuming that the right-side heading electric drive fails, the left-side heading electric drive can generate greater aerodynamic pull by increasing the rotation speed of the blade assembly, and then balance the anti-torque of the helicopter to maintain the heading stability of the helicopter; when the electric drive system fails completely, the helicopter first quickly reduces the main rotor collective pitch to a preset low position (usually 0), performs sideslip flight through forward or lateral control, and finally performs autorotation descent.

[0037] In the preferred technical solution of the embodiment of the present invention, when the positive blade assembly 6 is installed on the output shaft of the electric drive on the right side of the heading of the electric drive system 8, it can be detachably fixedly connected by bolts, which is helpful for disassembly and maintenance in the later stage, and can be composed of 2 to 3 blades; explanatorily, the positive blade assembly 6 generates aerodynamic pull through rotation, generates a moment on the center of mass of the helicopter, balances the torque of the main rotor 2 of the helicopter or forms an additional yaw moment. Correspondingly, when the counter blade assembly 7 is installed on the output shaft of the electric drive on the left side of the heading of the electric drive system 8, it can be detachably fixedly connected by bolts, which is helpful for disassembly and maintenance in the later stage, and can be composed of 2 to 3 blades; explanatorily, the counter blade assembly 7 generates aerodynamic pull through rotation, and the direction of its pull is the same as that of the positive blade assembly 6, and it also generates a moment on the center of mass of the helicopter, and together with the positive blade assembly 6, it balances the counter torque of the main rotor 2 of the helicopter or forms an additional yaw moment.

[0038] In the specific exemplary technical solution of the embodiment of the present invention, the electric drive system can be installed on the mounting bracket at the rear end of the tail beam of the fuselage and fixedly connected by bolts, which is convenient for disassembly and maintenance in the later stage; among them, the types of the electric drive on the right side of the heading and the electric drive on the left side of the heading can be two outer rotor motors, two inner rotor motors; or, one outer rotor motor and one inner rotor motor are used in combination.

[0039] In the specific exemplary technical solution of the embodiment of the present invention, the control system can transmit current and control signals through the electrical system and can be used to adjust and control the speed of the electric drive system. In the specific exemplary technical solution, the control system can adopt 6V power supply and PWM (Pulse Width Modulation) signal control.

[0040] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them. Although the present invention has been described in detail with reference to the above embodiments, those of ordinary skill in the art should understand that: the specific implementation manners of the present invention can still be modified or equivalently replaced, and any modification or equivalent replacement without departing from the spirit and scope of the present invention shall be covered by the protection scope of the claims of the present invention.

Claims

1. A helicopter, characterized in that, Comprising: A fuselage (1) having a longitudinal axis, a main rotor (2) disposed above the middle of the fuselage (1), a power system (3) disposed inside the fuselage (1), a control system (4) disposed inside the fuselage (1), and an electric-driven tail rotor (5) disposed at the rear of the fuselage (1); wherein, The electric-driven tail rotor (5) comprises: an electric drive system (8), a forward blade assembly (6), and a reverse blade assembly (7); wherein, the electric drive system (8) is fixedly installed at the rear of the fuselage (1), and the electric drive system (8) comprises a right-heading electric drive and a left-heading electric drive; the forward blade assembly (6) is installed on the output shaft of the right-heading electric drive, and the reverse blade assembly (7) is installed on the output shaft of the left-heading electric drive; the reverse blade assembly (7) and the forward blade assembly (6) are symmetrically disposed on both sides of the rear of the fuselage (1) for generating aerodynamic pulling forces having the same direction and perpendicular to the longitudinal axis of the fuselage (1); The control system (4) is used for controlling the rotational speeds of the right-heading electric drive and the left-heading electric drive in the electric drive system (8) according to a preset control algorithm.

2. A helicopter according to claim 1, characterized in that, The central mounting axes of the forward blade assembly (6) and the reverse blade assembly (7) are kept consistent.

3. A helicopter according to claim 1, characterized in that, Both the right-heading electric drive and the left-heading electric drive are external rotor motors or internal rotor motors.

4. A helicopter according to claim 1, characterized in that, In the step that the control system (4) executes controlling the rotational speeds of the right-heading electric drive and the left-heading electric drive in the electric drive system (8) according to a preset control algorithm, When the helicopter performs hovering control, the rotational speeds of the forward blade assembly (6) and the reverse blade assembly (7) are kept consistent, and the reaction torque formed by the rotational movement of the main rotor (2) is trimmed.

5. A helicopter according to claim 1, characterized in that, In the step that the control system (4) executes controlling the rotational speeds of the right-heading electric drive and the left-heading electric drive in the electric drive system (8) according to a preset control algorithm, When the helicopter performs yaw control, the rotational speeds of the forward blade assembly (6) and the reverse blade assembly (7) are kept consistent, and an additional yaw moment is formed by simultaneously increasing the rotational speeds of the forward blade assembly (6) and the reverse blade assembly (7).

6. A helicopter according to claim 5, characterized in that, In the process of forming the additional yaw moment by simultaneously increasing the rotational speeds of the forward blade assembly (6) and the reverse blade assembly (7), the satisfied condition is: Ωt = (k1 + k2 + k3)·k·Ωm; In the formula, Ωt is the rotational speed of the tail rotor; k is the theoretically trimmed value of the rotational speed of the main rotor and the rotational speed of the tail rotor; k1 is the influence factor of the rotational speed fluctuation of the main rotor; k2 is the influence factor of gust interference; k3 is the yaw command requirement; Ωm is the rotational speed of the main rotor.

7. A helicopter according to claim 1, characterized in that, In the step that the control system (4) executes controlling the rotational speeds of the right-heading electric drive and the left-heading electric drive in the electric drive system (8) according to a preset control algorithm, When the helicopter performs roll control, the positive blade assembly (6) and the negative blade assembly (7) are controlled by speed differential to preferentially trim the yaw-direction moment of the helicopter; then, an additional pitch moment is generated by differential fine-tuning to counteract the pitch moment generated by the helicopter due to roll.

8. A helicopter according to claim 1, characterized in that, In the step where the control system (4) executes controlling the rotational speeds of the right-side electric drive and the left-side electric drive in the electric drive system (8) according to a preset control algorithm, When the helicopter performs pitch control, the pulling direction of the main rotor (2) is adjusted and controlled to make the helicopter perform pitch motion; in addition, the electric drive tail rotor (5) is also controlled by speed differential to compensate and control the moment in the pitch direction of the helicopter on the premise of preferentially trimming the yaw-direction moment of the helicopter.

9. A helicopter according to claim 1, characterized in that, In the step where the control system (4) executes controlling the rotational speeds of the right-side electric drive and the left-side electric drive in the electric drive system (8) according to a preset control algorithm, When only the right-side electric drive in the heading fails, the rotational speed of the negative blade assembly (7) is increased by the left-side electric drive in the heading to generate aerodynamic pulling force to trim the yaw-direction moment of the helicopter; wherein, the additional pitch moment generated by the electric drive tail rotor (5) is compensated and controlled by the control system (4) controlling the main rotor (2). Or, when only the left-side electric drive in the heading fails, the rotational speed of the positive blade assembly (6) is increased by the right-side electric drive in the heading to generate aerodynamic pulling force to trim the yaw-direction moment of the helicopter; wherein, the additional pitch moment generated by the electric drive tail rotor (5) is compensated and controlled by the control system (4) controlling the main rotor (2).

10. A helicopter according to claim 1, characterized in that, In the step where the control system (4) executes controlling the rotational speeds of the right-side electric drive and the left-side electric drive in the electric drive system (8) according to a preset control algorithm, When both the right-side electric drive and the left-side electric drive in the heading fail, the helicopter first reduces the collective pitch of the main rotor (2) to a preset low level, then performs sideslip flight by forward or lateral control, and finally performs autorotative descent.

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