A helicopter

Abstract

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.

Description

Technical Field

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

[0002] Traditional helicopter tail rotor systems mostly use mechanical drive control, a technical approach characterized by complex construction and high maintenance costs. For small and medium-sized helicopters, the industry has proposed a direct-drive, variable-speed tail rotor system with an electric motor. This system directly drives the tail rotor, and adjusts the tail rotor speed to control thrust or pull. This is then balanced with the main rotor torque to achieve anti-torque balance for the helicopter.

[0003] Currently, while the aforementioned variable-speed tail rotor systems employing direct motor drive have simplified their structure, during speed changes, not only does aerodynamic drag significantly change, but torque also varies accordingly. This torque variation, in extreme cases, can adversely affect the helicopter's pitch attitude, increasing the helicopter's yaw-pitch coupling and posing challenges to flight control. Further explaining this, when existing direct motor-driven variable-speed tail rotor systems change speed, the corresponding change in torque can cause the helicopter's pitch attitude to change, further increasing the yaw-pitch coupling. This coupling complicates helicopter flight control and requires precise control algorithms to compensate and offset this effect, making implementation extremely difficult and costly. Furthermore, the increased coupling can affect the helicopter's flight stability and maneuverability, and in some cases, may even lead to flight accidents. In summary, given the technical difficulties that remain in the application of existing direct motor-driven variable-speed tail rotor systems, there is an urgent need to develop a helicopter employing a new electric-driven tail rotor. Summary of the Invention

[0004] The object of the present invention is to provide a helicopter that solves the technical problem of increasing the yaw-pitch coupling effect of the helicopter in the application of the existing variable speed electric drive tail rotor.

[0005] In order to achieve the above object, the present invention adopts the following technical solutions:

[0006] The present invention provides 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; wherein:

[0007] 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 mounted on the rear of the fuselage, and the electric drive system includes a right-side heading electric drive and a left-side heading electric drive; the forward 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 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 pull in the same direction and perpendicular to the longitudinal axis of the fuselage;

[0008] The control system is used to control the rotational speeds of the heading right electric drive and the heading left electric drive in the electric drive system according to a preset control algorithm.

[0009] A further improvement of the technical solution of the present invention is that the central installation axes of the positive blade assembly and the reverse blade assembly remain consistent.

[0010] A further improvement of the technical solution of the present invention is that the right-hand side electric drive and the left-hand side electric drive are both outer rotor motors or inner rotor motors.

[0011] A further improvement of the technical solution of the present invention is that, in the step of controlling the rotational speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system according to a preset control algorithm, the control system

[0012] When the helicopter is in hovering control, the rotation speeds of the positive blade assembly and the reverse blade assembly are kept consistent, and the reverse torque generated by the rotational motion of the main rotor is balanced.

[0013] A further improvement of the technical solution of the present invention is that, in the step of controlling the rotational speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system according to a preset control algorithm, the control system

[0014] When the helicopter performs yaw control, the rotational speeds of the positive blade assembly and the reverse blade assembly remain consistent, and an additional yaw moment is generated by simultaneously increasing the rotational speeds of the positive blade assembly and the reverse blade assembly.

[0015] A further improvement of the technical solution of the present invention is that, in the process of forming the additional yaw moment by simultaneously increasing the rotation speeds of the positive blade assembly and the reverse blade assembly, the following conditions are satisfied:

[0016] Ωt=(k1+k2+k3)·k·Ωm;

[0017] Where Ωt is the tail rotor speed; k is the theoretical balance value between the main rotor speed and the tail rotor speed; k1 is the main rotor speed fluctuation factor; k2 is the gust interference factor; k3 is the yaw command requirement; Ωm is the main rotor speed.

[0018] A further improvement of the technical solution of the present invention is that, in the step of controlling the rotational speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system according to a preset control algorithm, the control system

[0019] When the helicopter is performing roll control, the positive blade assembly and the reverse blade assembly are controlled by speed differential to prioritize balancing the yaw moment of the helicopter; then, additional pitch moment is generated by differential fine-tuning to offset the pitch moment generated by the helicopter's roll.

[0020] A further improvement of the technical solution of the present invention is that, in the step of controlling the rotational speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system according to a preset control algorithm, the control system

[0021] When the helicopter is performing pitch control, the pulling direction of the main rotor is adjusted to make the helicopter pitch. In addition, the electric tail rotor is controlled by speed differential to compensate for the moment in the pitch direction of the helicopter while giving priority to balancing the moment in the yaw direction of the helicopter.

[0022] A further improvement of the technical solution of the present invention is that, in the step of controlling the rotational speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system according to a preset control algorithm, the control system

[0023] When only the right-hand electric drive fails, the rotation speed of the reverse blade assembly is increased by the left-hand electric drive to generate aerodynamic force to balance the yaw moment of the helicopter; the additional pitch moment generated by the electric tail rotor is compensated by the main rotor controlled by the control system;

[0024] Alternatively, when only the left-side electric drive fails, the rotation speed of the positive blade assembly is increased by the right-side electric drive to generate aerodynamic pull to balance the yaw moment of the helicopter; wherein, the additional pitch moment generated by the electric tail rotor is compensated by the main rotor controlled by the control system.

[0025] A further improvement of the technical solution of the present invention is that, in the step of controlling the rotational speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system according to a preset control algorithm, the control system

[0026] When both the right and left heading electric drives fail, the helicopter first lowers the main rotor collective pitch to a preset low position, then performs sideslip flight through forward or lateral control, and finally performs autorotation descent.

[0027] Compared with the prior art, the present invention has the following beneficial effects:

[0028] In the technical solution disclosed in the present invention, the main function of the fuselage with a longitudinal axis is to provide installation space or interfaces 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 part of the fuselage is to provide aerodynamic pulling force for the entire aircraft, which can overcome the gravity of the entire aircraft to achieve flight in the air, accelerate and decelerate the entire aircraft by adjusting the magnitude of the pulling force, and achieve longitudinal and lateral movement of the attitude by adjusting the direction of the pulling force; the main function of the power system arranged inside the fuselage is to provide a power source for the entire aircraft, converting chemical energy into mechanical energy and transmitting it to the main rotor and electric-driven tail rotor, as well as providing 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 of the entire aircraft and the execution of missions, and to achieve functions such as attitude stabilization, 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 balancing function for the heading attitude of the entire aircraft, and to achieve functions such as hovering and yaw movement by adjusting the magnitude of the tail rotor pulling force. In the technical solution disclosed in 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 part 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 changes, according to the torque balance principle in mechanics, the torques generated by the electric drives on both sides can offset each other; this symmetrical structure and torque balance mechanism enable the torque on one side to increase during the speed change process and the torque on the other side to 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; explanatoryly, because when the speed of the electric drive on one side changes, resulting in a torque change, the electric drive on the other side can cooperate through reasonable control to avoid undesirable pitch changes of the fuselage due to excessive torque on one side.

[0029] In a further preferred embodiment of the present invention, the control system can independently control the rotational speed 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 principle of dynamics, the aerodynamic pull and torque generated by the blades are closely related to the rotational speed of the electric drive. By independently adjusting the rotational speed of the electric drives on both sides, the aerodynamic pull and torque generated by the positive and negative blade assemblies can be accurately adjusted. For example, 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 rotational speed 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 needs; for example, if it is detected that the pitch attitude change is about to be caused by the speed change, the control system can promptly adjust the speed of the electric drive on one side or both sides to generate opposite torque to compensate, thereby reducing the impact of the yaw-pitch coupling effect. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] In order to more clearly illustrate the technical solutions in the present invention or the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below; obviously, the drawings described below are some embodiments of the present invention, and for ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0031] Figure 1 1 is a schematic structural diagram of a helicopter according to an embodiment of the present invention;

[0032] Figure 2 is a schematic diagram of an electric-driven tail rotor in a helicopter according to an embodiment of the present invention;

[0033] Figure 3 : is a schematic diagram of the principle of speed control in an electric-driven tail rotor in an embodiment of the present invention; 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 theoretical balance value of the main rotor speed and the tail rotor speed, k1 is the main rotor speed fluctuation influence factor, k2 is the gust interference influence factor, and k3 is the yaw command requirement;

[0034] The explanations of the reference numerals in the figures are as follows: 1. fuselage; 2. main rotor; 3. power system; 4. control system; 5. electric-driven tail rotor; 6. forward blade assembly; 7. reverse blade assembly; 8. electric drive system. DETAILED DESCRIPTION

[0035] In order to make the purpose, technical solutions and advantages of the present invention clearer, the technical solutions of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention; it is obvious that the described embodiments and technical solutions are only part of the embodiments of the present invention, not all of the embodiments.

[0036] All other embodiments obtained by persons of ordinary skill in the art based on the technical solutions disclosed in the embodiments of the present invention without creative effort shall fall within the scope of protection of the present invention. In addition, the terms "including" and "having" and any variations thereof are intended to cover non-exclusive inclusions. For example, a process, method, system, product, or apparatus that includes a series of steps or units is not necessarily limited to those steps or units explicitly listed, but may include other steps or units that are not explicitly listed or that are inherent to these processes, methods, products, or apparatuses.

[0037] See also Figures 1 to 3An embodiment of the present invention provides a helicopter, comprising: a fuselage having a longitudinal axis, a main rotor 2 disposed above the middle portion of the fuselage, a power system 3 disposed inside the fuselage, a control system 4 disposed inside the fuselage, and an electric tail rotor 5 disposed at the rear of the fuselage;

[0038] The electric-driven tail rotor 5 specifically includes: an electric drive system 8, a positive blade assembly 6 and a reverse blade assembly 7; wherein 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 of the tail boom), and the electric drive system 8 includes: a right-side heading electric drive and a left-side heading electric drive, the positive blade assembly 6 is installed on the output shaft of the right-side heading electric drive, and the reverse blade assembly 7 is installed on the output shaft of the left-side heading electric drive (specifically, the reverse blade assembly 7 and the positive blade assembly 6 are symmetrically arranged on the left and right sides of the rear of the fuselage, and the aerodynamic pull directions they generate are all pointing in the same direction, It is perpendicular to the longitudinal axis of the fuselage and is mainly used to balance the counter-torque generated by the rotational movement of the main rotor 2; in the exemplary technical solution, the positive blade assembly 6, when viewed from the right side of the heading, performs counterclockwise rotation, and the direction of the aerodynamic pull generated is directed to the left side of the heading; the reverse blade assembly 7, when viewed from the right side of the heading, performs clockwise rotation, and the direction of the aerodynamic pull generated is also directed to the left side of the heading); the control system 4 is used to control the rotation conditions 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; in a further preferred specific technical solution, the central installation axes of the positive blade assembly 6 and the reverse blade assembly 7 are kept consistent.

[0039] In the technical solutions provided by the embodiments of the present invention, the fuselage is the main frame of the helicopter, providing a mounting base and structural support for other components, ensuring the stability and strength of the entire helicopter. Specifically, the fuselage typically contains a payload compartment for carrying mission-required equipment or cargo, such as sensors, communications equipment, cameras, and cargo racks. The fuselage's outer shell protects key components, such as the power system, from environmental damage (e.g., wind, rain, and dust), ensuring the helicopter's normal operation.

[0040] In the technical solutions provided by the embodiments of the present invention, the main rotor is the primary source of lift for a helicopter. Its high-speed rotation accelerates the airflow above the wing, reducing its pressure and generating upward lift, enabling the helicopter to take off and land vertically and hover. By varying the main rotor's pitch (i.e., the tilt angle of the blades), the helicopter's flight attitude, including pitch, roll, and yaw, can be controlled. For example, increasing the leading-edge pitch causes the helicopter to tilt upward, while increasing the trailing-edge pitch causes it to tilt downward. During forward flight, the main rotor generates forward thrust through cyclic pitch control (i.e., varying the blade pitch in different orientations), enabling the helicopter to fly horizontally.

[0041] In the technical solutions provided by the embodiments of this invention, the power system provides energy for the main rotor and electrically driven tail rotor. It typically uses either an electric or fuel-powered system, with the specific choice depending on the helicopter's design requirements and application scenario. The power system must be able to adjust power to accommodate varying power requirements during different flight phases (e.g., takeoff, hovering, cruising, and landing).

[0042] 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 rotates the fuselage. The electric-driven tail rotor balances this counter-torque by generating thrust in the opposite direction of the counter-torque, thereby maintaining the stable flight attitude of the helicopter. By changing the rotation speed of the electric-driven tail rotor, the heading of the helicopter can be controlled; for example, increasing the tail rotor rotation speed can make the helicopter yaw to one side, and reducing the rotation speed can make the helicopter 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 complex and changing flight environments. Further explaining, the embodiment of the present invention provides a new electric-driven tail rotor, 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, thereby improving the spatial efficiency of the helicopter structure.

[0043] In the technical solution disclosed in the embodiment of the present invention, the preset control algorithm includes: when the helicopter is hovering, the rotation speeds of the positive blade assembly 6 and the reverse blade assembly 7 are kept consistent, and the counter-torque generated by the rotation of the main rotor 2 is balanced; when the helicopter is performing left yaw control, the rotation speeds of the positive blade assembly 6 and the reverse blade assembly 7 are kept consistent, and are implemented by increasing the motor speed; explanatoryally, when the helicopter is hovering or yaw control, the rotation speeds of the positive blade assembly 6 and the reverse blade assembly 7 are kept consistent, balancing the torque of the main rotor 2 or forming an additional yaw moment without generating an additional pitch moment.

[0044] The technical solution of the present invention uses a preset control algorithm to precisely control the rotational speeds of the forward and reverse blade assemblies in a helicopter's electric-driven tail rotor, thereby achieving stable flight in different flight states (hover, left yaw control), while also avoiding the generation of additional pitching moments, reducing yaw-pitch coupling, and improving flight control stability and reliability. Explanatory note: Existing electric-driven tail rotor solutions only control one blade assembly, adjusting the tail rotor speed to balance the counter-torque generated by the rotation of the helicopter's main rotor. During hover or yaw control, the steady-state speed of the main rotor fluctuates within a certain range (typically 3% to 5%) during actual flight, and the external environment is also subject to certain wind disturbances. Therefore, in hover / yaw states, existing electric-driven tail rotor solutions adjust the tail rotor speed based on the fluctuation range of the main rotor speed, the magnitude of wind disturbances, and the required yaw angle. However, the change in tail rotor speed also introduces additional pitching moments to the entire helicopter, thereby affecting the pitch attitude control of the entire aircraft. In contrast, the technical solution disclosed in the embodiments of the present invention incorporates a new electrically driven tail rotor that balances the anti-torque of the helicopter's main rotor by controlling the forward and reverse blade assemblies. When the main rotor speed fluctuates, gusts interfere, or yaw control occurs, the forward and reverse blade assemblies are synchronously adjusted and controlled by the electric drive system. This only alters the magnitude of the aerodynamic pull, which is always perpendicular to the longitudinal axis of the fuselage. The additional pitching moments generated by the individual blade assemblies are equal in magnitude and opposite in direction, and their combined effects cancel each other out, without adding any additional pitching moment to the entire aircraft.

[0045] In a further preferred technical solution of the embodiment of the present invention, the rotational speed of the electric tail rotor is linearly related to the rotational speed of the main rotor, the wind disturbance factor and the yaw command, and the conditions satisfied are:

[0046] Ωt=(k1+k2+k3)·k·Ωm, closed-loop control is performed; illustratively and preferably, a PID controller is used for control;

[0047] Where k is the theoretical balance value between the main rotor speed and the tail rotor speed, which can be obtained through theoretical calculation; k1 is the main rotor speed fluctuation influencing factor, which can be obtained through experiments; k2 is the gust interference influencing factor, which can be corrected through experiments; k3 is the yaw command requirement, which can be input according to actual needs; Ωt is the tail rotor speed, and Ωm is the main rotor speed.

[0048] As a preferred technical solution of an embodiment of the present invention, the preset control algorithm also includes: when the helicopter is performing roll control: controlling the positive blade assembly and the reverse blade assembly through the speed differential to prioritize balancing the yaw direction torque of the helicopter, and then generating additional pitch torque through differential fine-tuning to offset the pitch torque generated by the helicopter's roll.

[0049] In an embodiment of the present invention, when a helicopter performs roll control, the fuselage rotates around its longitudinal axis. This action not only changes the roll attitude of the helicopter, but may also trigger a series of complex torque changes; among which, the changes in the yaw torque and the pitch torque are particularly critical. The change in the yaw torque may cause the helicopter to deviate from the predetermined heading, while the change in the pitch torque may cause the helicopter to raise its head or lower its head, affecting the stability and controllability of the flight. When the helicopter performs roll control, the preset control algorithm controls the positive blade assembly and the reverse blade assembly through the speed differential, that is, adjusts the speed of the right-hand heading electric drive and the left-hand heading electric drive respectively, so that the speed of the positive and reverse blade assemblies is different.

[0050] To explain the principle, the forward and reverse blade assemblies rotate in opposite directions, and when their rotational speeds are different, they generate torques of different magnitudes and directions. By properly controlling the speed difference, the helicopter's yaw torque can be prioritized. For example, when a helicopter rolls to the right, a torque may be generated that causes the helicopter to yaw to the left. At this time, by adjusting the speed difference between the forward and reverse blade assemblies, the yaw torque generated by the tail rotor and this unfavorable yaw torque are offset, thereby maintaining the helicopter's yaw stability and avoiding heading deviation caused by roll. When a helicopter is performing roll control, the change in the fuselage attitude will disrupt 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 tend to lower its head, generating additional pitch moment. If not controlled, this will affect the helicopter's pitch attitude stability. After prioritizing yaw moment, further fine-tuning the speed differential between the positive and negative blade assemblies generates an additional pitching moment. This additional pitching moment is equal in magnitude and opposite in direction to the pitching moment generated by the helicopter's roll, precisely offsetting the pitching effect of the roll. For example, if a roll causes the helicopter to pitch upward, the additional pitching moment generated by differential fine-tuning will cause the helicopter to pitch downward. The two forces counteract each other, maintaining a stable pitch attitude.

[0051] Explanatory note: Existing electric tail rotor solutions only adjust and control the direction of the main rotor's pulling force to cause the helicopter to roll. However, as the helicopter rolls, the main rotor is constantly rotating during helicopter flight. The superposition of the two rotational motions generates an additional pitching moment in the pitch direction. A single-blade tail rotor typically only trims the yaw moment and therefore cannot eliminate the pitching moment caused by the helicopter's roll. In the new technical solution of the present invention, the forward and reverse blade assemblies can be controlled by speed differential to prioritize trimming the helicopter's yaw moment. Then, through differential fine-tuning, an additional pitching moment is generated to offset the pitching moment caused by the helicopter's roll, thereby reducing the helicopter's roll-pitch coupling and addressing the shortcomings of existing technical solutions. In summary, by prioritizing balancing the yaw moment and offsetting the pitching moment caused by roll, the present invention's control strategy can effectively reduce the helicopter's attitude fluctuations during roll and improve flight stability. By differentially controlling the rotational speeds of the forward and reverse blade assemblies, yaw and pitch moments are balanced, eliminating the need for complex control mechanisms or algorithms and simplifying control logic. By prioritizing balancing the yaw moment and then fine-tuning to offset the pitch moment, the coupling between yaw and pitch is reduced, making the helicopter's flight control more stable and precise.

[0052] As a preferred technical solution of an embodiment of the present invention, the preset control algorithm also includes: when the helicopter is pitching, it mainly adjusts the pulling direction of the main rotor 2 to make the helicopter pitch, and also controls the electric tail rotor 5 through differential control to compensate for the moment in the pitch direction of the helicopter while giving priority to balancing the moment in the yaw direction of the helicopter.

[0053] During helicopter flight, pitch control is a critical maneuver, enabling the helicopter to tilt its head up or down to meet various flight requirements, such as climbing, descending, or adjusting its flight trajectory. However, pitch control doesn't exist in isolation. Changes in the thrust direction of the main rotor during pitch motion not only affect the pitch attitude but can also trigger changes in the yaw torque, causing the helicopter to deviate from its intended course. Furthermore, other factors can also generate additional torque disturbances in the pitch direction, affecting pitch attitude stability. Existing electric tail rotor solutions only adjust the thrust direction of the main rotor to achieve pitch motion, while maintaining the tail rotor's original kinematic trim. The new technical solution disclosed in the embodiments of the present invention can mainly adjust the direction of the thrust 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 while 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 can include increasing flight speed, increasing range, improving flight stability and other aspects.

[0054] 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 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.

[0055] 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 will first quickly reduce the main rotor collective pitch to a preset low position (usually 0), perform sideslip flight through forward or lateral control, and finally perform autorotation descent.

[0056] 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 system 8 on the right side of the heading direction, it can be fixedly connected by bolts in a detachable manner, which facilitates later disassembly and maintenance. It can be composed of 2 to 3 blades. Explanatoryally, the positive blade assembly 6 generates aerodynamic pull through rotation, generating a torque on the center of mass of the helicopter, balancing the torque of the helicopter's main rotor 2 or forming an additional yaw moment. Correspondingly, when the reverse blade assembly 7 is installed on the output shaft of the electric drive system 8 on the left side of the heading direction, it can be fixedly connected by bolts in a detachable manner, which facilitates later disassembly and maintenance. It can be composed of 2 to 3 blades. Explanatoryally, the reverse blade assembly 7 generates aerodynamic pull through rotation, and its pulling direction is the same as that of the positive blade assembly 6. It also generates a torque on the center of mass of the helicopter, and together with the positive blade assembly 6, balances the anti-torque of the helicopter's main rotor 2 or forms an additional yaw moment.

[0057] 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 boom of the fuselage and fixed by bolts to facilitate later disassembly and maintenance; wherein, the types of the right-hand side heading electric drive and the left-hand side heading electric drive can be two outer rotor motors and two inner rotor motors; or, one outer rotor motor and one inner rotor motor can be used in combination.

[0058] In a specific exemplary embodiment of the present invention, a control system can transmit current and control signals through the electrical system to adjust and control the speed of the electric drive system. In this specific exemplary embodiment, the control system can utilize a 6V power supply and PWM (Pulse Width Modulation) signal control.

[0059] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and not to limit it. Although the present invention has been described in detail with reference to the above embodiments, ordinary technicians in the field should understand that the specific implementation methods of the present invention can still be modified or replaced by equivalents. Any modification or equivalent replacement that does not depart from the spirit and scope of the present invention should be covered by the scope of protection of the claims of the present invention.

Claims

1. A helicopter, characterized in that: include: A fuselage (1) having a longitudinal axis, a main rotor (2) arranged above the middle of the fuselage (1), a power system (3) arranged inside the fuselage (1), a control system (4) arranged inside the fuselage (1), and an electric tail rotor (5) arranged at the rear of the fuselage (1); wherein, The electric drive tail rotor (5) comprises: an electric drive system (8), a positive blade assembly (6) and a reverse blade assembly (7); wherein the electric drive system (8) is fixedly mounted on the rear portion of the fuselage (1), and the electric drive system (8) comprises a right-side heading electric drive and a left-side heading electric drive; the positive blade assembly (6) is mounted on the output shaft of the right-side heading electric drive, and the reverse blade assembly (7) is mounted on the output shaft of the left-side heading electric drive; the reverse blade assembly (7) and the positive blade assembly (6) are symmetrically arranged on both sides of the rear portion of the fuselage (1), and are used to generate aerodynamic pulling forces in the same direction and perpendicular to the longitudinal axis of the fuselage (1); The control system (4) is used to control the rotation speeds of the heading right electric drive and the heading left electric drive in the electric drive system (8) according to a preset control algorithm; wherein, when the helicopter performs yaw control, the rotation speeds of the positive blade assembly (6) and the reverse blade assembly (7) are kept consistent, and an additional yaw moment is formed by simultaneously increasing the rotation speeds of the positive blade assembly (6) and the reverse blade assembly (7); in the process of forming the additional yaw moment, the conditions satisfied are: Ωt=(k1+k2+k3)·k·Ωm; Where Ωt is the tail rotor speed; k is the theoretical balance value between the main rotor speed and the tail rotor speed; k1 is the main rotor speed fluctuation factor; k2 is the gust interference factor; k3 is the yaw command requirement; Ωm is the main rotor speed.

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

3. A helicopter according to claim 1, characterized in that: The right-hand side electric drive and the left-hand side electric drive are both outer rotor motors or inner rotor motors.

4. A helicopter according to claim 1, characterized in that: The control system (4) executes the step of controlling the rotation speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system (8) according to a preset control algorithm, When the helicopter is in hovering control, the rotation speeds of the positive blade assembly (6) and the reverse blade assembly (7) are kept consistent, and the reverse torque generated by the rotational movement of the main rotor (2) is balanced.

5. The helicopter according to claim 1, characterized in that: The control system (4) executes the step of controlling the rotation speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system (8) according to a preset control algorithm, When the helicopter is performing roll control, the positive blade assembly (6) and the reverse blade assembly (7) are controlled by speed differential to prioritize balancing the yaw moment of the helicopter; then, an additional pitch moment is generated by differential fine-tuning to offset the pitch moment generated by the helicopter's roll.

6. A helicopter according to claim 1, characterized in that: The control system (4) executes the step of controlling the rotation speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system (8) according to a preset control algorithm, When the helicopter is performing 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 controlled by speed differential, and the moment in the pitch direction of the helicopter is compensated and controlled under the premise of giving priority to balancing the moment in the yaw direction of the helicopter.

7. A helicopter according to claim 1, characterized in that: The control system (4) executes the step of controlling the rotation speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system (8) according to a preset control algorithm, When only the right-hand electric drive fails, the rotation speed of the counter-blade assembly (7) is increased by the left-hand electric drive to generate aerodynamic pulling force to balance the yaw moment of the helicopter; wherein the additional pitch moment generated by the electric tail rotor (5) is controlled by the control system (4) to control the main rotor (2) for compensation control; Alternatively, when only the left-side electric drive fails, the rotation speed of the positive blade assembly (6) is increased by the right-side electric drive to generate aerodynamic pulling force to balance the yaw moment of the helicopter; wherein the additional pitch moment generated by the electric tail rotor (5) is controlled by the control system (4) to control the main rotor (2) for compensation control.

8. The helicopter according to claim 1, characterized in that: The control system (4) executes the step of controlling the rotation speeds of the right-hand electric drive and the left-hand electric drive in the electric drive system (8) according to a preset control algorithm, When both the right-hand electric drive and the left-hand electric drive fail, the helicopter first lowers the total pitch of the main rotor (2) to a preset low position, then performs sideslip flight through forward or lateral control, and finally performs autorotation descent.

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