A dual-redundant electric power system suitable for light electric helicopters

Abstract

The present invention belongs to the technical field of new energy aviation electric power systems, and in particular, relates to a dual-redundant electric power system suitable for light electric helicopters. The system includes an intelligent control system, rotors, batteries, a high-voltage box, a gearbox, a rotor speed setting device, a power integrated display, and a battery management system (BMS). The intelligent control system includes electric power system 1, electric power system 2, and a speed control box. Electric power system 1 and electric power system 2 exchange information via CAN communication. The dual-redundant electric power system provided by the present invention provides a reliable and safe power solution for electric helicopters, positively impacting helicopter operations from multiple perspectives, including flight safety, reliability, and performance optimization.

Description

Technical Field

[0001] The present invention relates to the technical field of new energy aviation electric power systems, and in particular to a dual-redundant electric power system suitable for light electric helicopters. Background Art

[0002] As a green and pollution-free aircraft, electric helicopters have broad application prospects. They typically utilize a permanent magnet synchronous motor (PMSM) equipped with a vector control system to form an electric propulsion system. As a core component of an aircraft, the electric propulsion system, comprised of the motor and controller, demands high reliability. Traditional electric helicopters typically rely on a single electric propulsion system. Failures in either or both the motor and controller can lead to a loss of power, potentially resulting in a crash and jeopardizing operational safety. Summary of the Invention

[0003] In view of this, the present invention discloses a dual-redundant electric power system suitable for a light electric helicopter to solve the problem of low operating safety of traditional electric helicopters due to their reliance on a single electric power system to provide power for the aircraft.

[0004] The technical solution provided by the present invention is specifically a dual-redundant electric power system suitable for a light electric helicopter, including an intelligent control system, a rotor, a battery, a high-voltage box, a gearbox, a rotor speed setting device, a power integrated display and a BMS10; the intelligent control system includes an electric power system 1, an electric power system 2 and a speed control box, and the electric power system 1 and the electric power system 2 exchange information via CAN communication;

[0005] The first electric power system includes a first motor and a first controller connected by signals, and the second electric power system includes a second motor and a second controller connected by signals; the first motor and the second motor are coaxially connected; the first controller and the second controller are provided with two control modes: speed and torque;

[0006] The high-voltage box is connected to the battery and is used to control the battery to output DC power;

[0007] The first controller and the second controller are connected to a battery, and are used to convert the direct current output by the battery into three-phase alternating current. The battery is used to provide power to the first motor and the second motor, and the battery provides the three-phase alternating current to the first motor and the second motor;

[0008] The rotor speed setting device is installed in the pilot's cockpit and is used to adjust the rotor speed. The rotor speed setting device is connected to the speed control box signal;

[0009] The speed control box is respectively connected to the first controller and the second controller by signal;

[0010] The power integrated display is used to monitor the operating status of the dual-redundant electric power system in real time, and collect and display the operating status information of the motor, battery and controller;

[0011] The gearbox is used to drive the first motor and the second motor to rotate the rotor to generate the thrust required by the aircraft;

[0012] The BMS is used to collect and detect the single cell information and operating status of the power battery in real time;

[0013] The BMS is connected to the first controller and the second controller via CAN bus signals respectively, and the BMS is also connected to a power integrated display.

[0014] Preferably, it also includes an auxiliary power supply battery, which is used to supply power to the power integrated display, BMS, first controller, second controller, high-voltage box, rotor speed setting device and speed control box.

[0015] The first motor and the second motor are provided with motor speed sensors at their ends, which are connected to the first controller and the second controller via the motor speed sensors;

[0016] The frequency of the three-phase AC power output by the controller is used to adjust the rotor speed in real time. The motor speed sensor feeds back the actual motor speed and position information to the corresponding controller in real time, forming a closed-loop control of the motor speed, thereby controlling the operating conditions of the aircraft.

[0017] The controller uses the rotor speed setting device to set the desired rotor speed as the control target;

[0018] The motor speed sensor monitors the actual rotor speed in real time and feeds it back to the controller. The feedback speed value is compared with the given speed value, and a PI speed closed-loop control strategy is adopted to achieve precise follow-up control of the rotor speed by the controller.

[0019] The three-phase output ends of the first controller and the second controller are provided with current transformers, and the input current of the motor is monitored through the current transformers;

[0020] The current transformer of the first controller monitors the input current of the first motor. The speed controller box collects the three-phase output current of the first controller in real time via CAN. The least squares polynomial parameter identification method is used for current sharing and filtering to obtain the actual controller output three-phase current value. The three-phase current value is converted into the torque set value required by the second controller and input into the second controller via the CAN bus.

[0021] The second controller receives the torque given value from the speed controller box, performs torque closed-loop control according to the torque given value, and performs real-time adjustment according to the real-time torque given value, ultimately achieving the same output torque of the first controller and the second controller.

[0022] The rotor speed setting device outputs a speed control signal to the speed control box;

[0023] The speed control box sends the speed command to the first controller via the CAN bus, and the speed control box determines whether the two electric power systems need to operate simultaneously according to the actual speed requirement;

[0024] When the speed control box determines that the actual speed demand can be provided by the electric power system 1 alone, the speed control box will issue a shutdown command to the second controller to stop the electric power system 2 from running;

[0025] Based on the actual speed requirement of the system, the speed control box determines that two electric power systems need to operate simultaneously. The speed control box sends a controller start command to the second controller to put the second power system into operation, and at the same time receives the actual current value of the motor sent by the first controller. The actual current value of the first controller is obtained, and the current is equalized and filtered using the least squares polynomial parameter identification method to convert it into an actual torque given value. The given torque value is sent to the second controller via the CAN bus as the input, and the second electric power system operates according to the given torque value to provide power for the aircraft.

[0026] The least squares polynomial parameter identification method:

[0027] Given the controller output current, the controller output current can be expressed by a polynomial, namely:

[0028]

[0029] in: is the current output by the first controller 12 at time t, a i are the coefficients of the polynomial, and n represents the order of the polynomial.

[0030] The sum of squares of the minimized residuals between the actual collected value and the polynomial predicted value is minimized, and we can get:

[0031]

[0032] Among them, I(t i ) is the actual observed value, is the polynomial prediction value;

[0033] The minimum residual sum of squares can be obtained from formula (2):

[0034]

[0035] The derivative of the minimum residual sum of squares with respect to the polynomial coefficients is zero:

[0036]

[0037] By calculating formula (4), we can get a0, a1, a2, ..., a n-1 , a n , thereby obtaining the characteristic curve of the output current of the first controller 12;

[0038] The motor output torque is proportional to the controller output current, and it can be obtained from formula (1)

[0039]

[0040] Among them, k M is the torque coefficient; M(t) is the motor output torque.

[0041] The high-voltage box includes a fast fuse and a contactor. The high-voltage box controls the on and off of the contactor according to the received on and off instructions of the contactor, thereby controlling the power DC power supply required by the first controller and the second controller.

[0042] The present invention provides a dual-redundant electric power system suitable for light electric helicopters. The dual-redundant electric power system provides a reliable and safe power solution for electric helicopters, which has a positive impact on helicopter operations from multiple perspectives such as flight safety, reliability and performance optimization.

[0043] It is to be understood that the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the disclosure of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and, together with the description, serve to explain the principles of the invention.

[0045] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, for ordinary technicians in this field, other drawings can be obtained based on these drawings without any creative work.

[0046] Figure 1 A structural diagram of a dual-redundant electric power system suitable for an electric helicopter provided in an embodiment of the present invention;

[0047] Figure 2 The present invention provides a flowchart of the operation of a dual-redundant electric power system suitable for an electric helicopter according to an embodiment of the present invention. DETAILED DESCRIPTION

[0048] Exemplary embodiments will now be described in detail, examples of which are illustrated in the accompanying drawings. In the following description, when referring to the drawings, like numbers in different figures represent the same or similar elements unless otherwise indicated. The embodiments described in the following exemplary embodiments are not intended to represent all possible embodiments consistent with the present invention. Rather, they are merely examples of systems consistent with certain aspects of the present invention, as detailed in the appended claims.

[0049] like Figure 1 As shown, this embodiment provides a dual-redundant electric power system suitable for an electric helicopter, including an intelligent control system, a rotor 3, a battery 4, a high-voltage box 5, a gearbox 7, a rotor speed setting device 8, a power integrated display 9 and a BMS; wherein the intelligent control system includes an electric power system 1, an electric power system 2 and a speed control box 6, and the electric power system 1 and the electric power system 2 exchange information via CAN communication; the battery is preferably a power lithium battery;

[0050] The dual-redundant electric power system for a light electric helicopter provided in this embodiment operates as follows: a power lithium battery acts as an energy storage element to provide electrical energy for the aircraft; a high-voltage box 5 controls the output of DC power from the power lithium battery; a controller converts the DC power from the power lithium battery into three-phase AC power, which in turn provides power to two motors. The two coaxially connected motors drive the rotors through a gearbox to generate the thrust required by the aircraft;

[0051] Specifically, the first electric power system includes a first motor 11 and a first controller 12 connected by signals, and the second electric power system includes a second motor 21 and a second controller 22 connected by signals; the first motor 11 and the second motor 21 are coaxially connected; motor speed sensors are provided at the ends of the first motor 11 and the second motor 21, and are connected to the first controller 12 and the second controller 22 by signals through the motor speed sensors;

[0052] This implementation utilizes a dual-redundant electric power system consisting of two electric power systems to power the aircraft. Each electric power system consists of an electric motor and controller. The two electric power systems exchange information via CAN communication, forming an intelligent control system that implements dual backup, coordination, and automatic switching. During normal flight, the two electric motors are installed in a coaxial series configuration, each equipped with its own controller, and jointly provide power to the aircraft based on its actual power requirements.

[0053] The first controller 12 and the second controller 22 are connected to the battery 4 and are used to convert the direct current output by the battery 4 into three-phase alternating current. The battery 4 is used to provide power to the first motor 11 and the second motor 21. The battery 4 provides the three-phase alternating current to the first motor 11 and the second motor 21.

[0054] The rotor speed setting device 8 is installed in the pilot's cabin and is used to adjust the rotor speed. The rotor speed setting device 8 is connected to the speed control box 6 by signal;

[0055] The speed control box 6 is signal-connected to the first controller 12 and the second controller 22 respectively;

[0056] The dual redundant electric power system also includes an auxiliary power supply battery 101, which serves as an auxiliary power source to power the power integrated display 9, BMS, first controller 12, second controller 22, high voltage box 5, rotor speed setting device 8 and speed control box 6.

[0057] Both auxiliary power supply batteries and power lithium batteries are placed in the aircraft's battery compartment. When the battery power is low, the battery needs to be removed from the aircraft's battery compartment and placed on a charging station near the hangar for charging. When the number of charging times reaches the limit or the capacity drops to 80%, the battery is eliminated.

[0058] The power integrated display 9 can monitor the operating status of the dual-redundant electric power system in real time, collect and display the operating status information of the motor, power lithium battery and controller, and provide the pilot with necessary flight information. This flight information is not limited to the motor speed, motor temperature, controller temperature, battery power, power lithium battery temperature, etc. When the electric power system components such as the motor, controller and battery fail, the power integrated display 9 will issue an alarm message for the pilot to use. The pilot can adjust the rotor speed by operating the rotor setting device 8 in the cockpit, thereby controlling the operating status of the aircraft.

[0059] The high-voltage box 5 is mainly composed of a fast fuse and a contactor. The high-voltage box 5 controls the on and off of the contactor according to the received contactor on and off instructions, thereby realizing the control of the DC power supply required by the first controller 12 and the second controller 22; when a short-term overcurrent fault such as a short circuit occurs, the fast fuse can be quickly disconnected, cutting off the DC power supply, protecting the dual redundant electric power system from fire due to short circuit or overcurrent, which endangers the operational safety of the aircraft.

[0060] BMS is used to collect real-time information about the single cells of the power lithium battery. This information is not limited to voltage, current, temperature and power, and monitors the operating status of the single cells in real time. When a single cell or power lithium battery fault is detected, it can be reported in time through the CAN bus, displayed on the power integrated display, and transmitted to the pilot. The pilot will handle it in time according to the fault level to ensure the safe operation of the aircraft.

[0061] The lightweight electric helicopter's dual-redundant electric powertrain utilizes a comprehensive control strategy to ensure precise rotor speed control under various flight conditions. This control approach integrates both speed and torque control, achieving stable, efficient, and safe flight by coordinating the outputs of the dual-redundant electric powertrain.

[0062] First, the controller uses rotor speed setting device 8 to set the desired rotor speed as the control target. A motor speed sensor mounted on the motor end monitors the actual rotor speed in real time and feeds it back to the controller. This feedback is compared with the set speed value, and a PI speed closed-loop control strategy is employed. This allows the controller to precisely track the rotor speed, ensuring stable flight.

[0063] At the same time, the current transformer installed at the three-phase output end of the controller monitors the input current of the corresponding motor in real time. The intelligent control system collects the three-phase current output by the first controller 12 in real time through CAN. Since the output three-phase current contains high-order harmonics and has large fluctuations, the least squares polynomial parameter identification method is used for current balancing and filtering to obtain the actual controller output three-phase current value, and the three-phase current value is converted into the torque set value required by the second controller 22, and input to the second controller 22 through the CAN bus. The second controller 22 receives the torque set value issued by the speed controller box, and the second controller 22 performs torque closed-loop control according to the torque set value; and performs real-time adjustment according to the real-time torque set value, and finally achieves the same output torque of the first controller 12 and the second controller 22, ensuring that the system will not affect the system output efficiency or directly cause a parking failure due to overcurrent of a single electric power system.

[0064] The intelligent control system composed of the speed control box of the present invention can also flexibly adjust the coordination mode of the two electric power systems according to the different flight states and requirements of the aircraft to optimize energy utilization and improve system efficiency. The specific implementation method is as follows:

[0065] The speed control box 6 serves as an intelligent control system. The pilot operates the rotor speed setting device 8 in the cockpit and sends the speed control signal of the output speed setting device to the speed control box 6 by changing the speed control signal of the rotor speed setting device. The speed control box 6 receives the speed instruction of the rotor speed setting device 8 and sends the speed instruction to the first controller 12 through the CAN bus. It determines whether two electric power systems need to run simultaneously according to the actual speed demand. When it is determined that the actual speed demand can be provided by the electric power system composed of the first controller 12 and the first motor 11 alone, the speed control box will issue a shutdown instruction to the second controller 22, so that the electric power system composed of the second controller 22 and the second motor 21 will stop running. Similarly, according to the actual speed demand of the system, the speed control box determines that two electric power systems need to run simultaneously, and the speed control box sends a controller start instruction to the second controller 22, so that the electric power system composed of the second controller 22 and the second motor 21 will stop running. The electric power system composed of the first controller 12 and the second motor 21 is put into operation, and at the same time receives the actual current value of the motor sent by the first controller 12. The actual current value of the first controller 12 is obtained, and after current averaging and filtering, it is converted into an actual torque given value, which is sent to the second controller 22 via the CAN bus as the input given torque value. The electric power system composed of the second controller 22 and the second motor 21 operates according to the given torque value to provide power for the aircraft; the operating status of the electric power system 1 and the electric power system 2 is monitored. When a fault is detected in one of the electric power systems, the faulty electric power system exits operation, and the other electric power system automatically switches to the speed mode and sends a fault instruction to the power integrated display for the pilot to use. After receiving the fault instruction, the pilot quickly selects the nearest location for forced landing to ensure the flight safety of the aircraft. The control flow chart is as follows Figure 2 shown.

[0066] The dual-redundant electric power system of the light electric helicopter precisely coordinates and synchronizes the two electric power systems, with their motors rotating coaxially in series to achieve stable flight and precise maneuverability. This design not only ensures the required thrust for the aircraft but also maintains lateral stability during flight, enhancing the helicopter's safety. An intelligent control system adjusts the output speed and torque of electric power systems 1 and 2 in real time, enabling the pilot to precisely control the flight state and ensuring the aircraft's stability and reliability during all flight phases, including takeoff, cruising, and landing. Furthermore, the redundant design of the dual-redundant electric power system ensures that if one electric power system fails, the other can still provide the power required for an emergency landing, ensuring a safe emergency landing. The dual-redundant electric power system of the light electric helicopter utilizes an integrated control strategy to ensure precise control of the rotor speed under different flight conditions. This control method integrates both speed and torque control, achieving stable, efficient, and safe helicopter flight by coordinating the outputs of the dual-redundant electric power systems.

[0067] For a dual-redundant electric power control system, the three-phase current output by the first controller 12 at different times is collected, and the least squares polynomial parameter identification method is used to obtain the actual output current characteristic curve of the first controller 12. The least squares polynomial parameter identification method is used to establish a mathematical model based on the relationship between the input and output current of the first controller 12, and to find the optimal model parameters from the actual collected data so that the output of the model is closest to the actual observed value. The least squares polynomial parameter identification method:

[0068] Given the controller output current, the controller output current can be expressed by a polynomial, namely:

[0069]

[0070] in: is the current output by the first controller 12 at time t, a i are the coefficients of the polynomial, and n represents the order of the polynomial.

[0071] The sum of squares of the minimized residuals between the actual collected value and the polynomial predicted value is minimized, and we can get:

[0072]

[0073] Among them, I(t i ) is the actual observed value, is the polynomial prediction value;

[0074] The minimum residual sum of squares can be obtained from formula (2):

[0075]

[0076] The derivative of the minimum residual sum of squares with respect to the polynomial coefficients is zero:

[0077]

[0078] By calculating formula (4), we can get a0, a1, a2, ..., a n-1 , a n , thereby obtaining the characteristic curve of the output current of the first controller 12;

[0079] The motor output torque is proportional to the controller output current, and it can be obtained from formula (1)

[0080]

[0081] Among them, k M is the torque coefficient; M(t) is the motor output torque.

[0082] The dual-redundant electric power system of the light electric helicopter proposed in this implementation plan uses two relatively independent electric power systems to provide power for the aircraft. When one electric power system fails and shuts down, the other electric power system can quickly switch to the speed control mode, ensuring that the power of the helicopter power system is not lost, providing power guarantee for the emergency landing of the helicopter, and ensuring the flight safety of the aircraft from the perspective of hardware redundancy.

[0083] As the helicopter's intelligent control system, the speed controller box monitors the helicopter's power requirements in real time under different operating conditions, and balances the power output of the two electric power systems according to the power requirements, ensuring the efficient operation of the helicopter's electric power system, increasing the aircraft's flight time, and reducing the heat generated by the helicopter's electric power system.

[0084] The least squares polynomial parameter identification method is applied to the current balancing control of the electric power system, which ensures the consistency of the output torque of the two electric power systems, improves the overall efficiency of the dual-redundant electric power system of the light electric helicopter, and ensures the flight safety of the helicopter, eliminating the occurrence of shutdown failures due to current imbalance between the two electric power systems.

[0085] The dual-redundant electric power system of a light electric helicopter involved in the present invention is a core component of the aircraft and is composed of multiple key components that work together to achieve propulsion, control, communication, display and fault diagnosis of the aircraft.

[0086] Other embodiments of the present invention will readily occur to those skilled in the art after considering the specification and practicing the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the present invention that follow the general principles of the invention and include common knowledge or customary techniques in the art not disclosed herein. The description and examples are to be considered as exemplary only, with the true scope and spirit of the invention being indicated by the claims.

Claims

1. A dual-redundant electric power system suitable for a light electric helicopter, characterized in that: The invention comprises an intelligent control system, a rotor (3), a battery (4), a high-voltage box (5), a gearbox (7), a rotor speed setting device (8), a power integrated display (9) and a BMS (10); the intelligent control system comprises an electric power system 1, an electric power system 2 and a speed control box (6); the electric power system 1 and the electric power system 2 exchange information via CAN communication; The first electric power system comprises a first motor (11) and a first controller (12) connected by signals, and the second electric power system comprises a second motor (21) and a second controller (22) connected by signals; the first motor (11) and the second motor (21) are coaxially connected; the first controller (12) and the second controller (22) are provided with two control modes: speed control and torque control; The high-voltage box (5) is connected to the battery (4) and is used to control the battery (4) to output direct current power; The first controller (12) and the second controller (22) are connected to the battery (4) and are used to convert the direct current output by the battery (4) into three-phase alternating current. The battery (4) is used to provide electric energy for the first motor (11) and the second motor (21). The battery (4) provides the three-phase alternating current to the first motor (11) and the second motor (21). A rotor speed setting device (8) is installed in the pilot's cabin and is used to adjust the rotor speed. The rotor speed setting device (8) is connected to the speed control box (6) via a signal. The speed control box (6) is respectively connected to the first controller (12) and the second controller (22) via signals; The power integrated display (9) is used to monitor the operating status of the dual-redundant electric power system in real time, and collect and display the operating status information of the motor, battery and controller; The gearbox (7) is used to drive the rotors of the first motor (11) and the second motor (21) to rotate to generate the thrust required by the aircraft; The BMS (10) is used to collect and detect the single cell information and operating status of the battery (4) in real time; The BMS (10) is connected to the first controller (12) and the second controller (22) via a CAN bus for signal communication, and the BMS (10) is also connected to a power integrated display (9); The three-phase output ends of the first controller (12) and the second controller (22) are provided with current transformers, and the input current of the motor is monitored through the current transformers; The current transformer of the first controller monitors the input current of the first motor, and the speed control box (6) collects the three-phase current output by the first controller in real time through the CAN bus, uses the least squares polynomial parameter identification method to perform current balancing and filtering, obtains the actual controller output three-phase current value, and converts the three-phase current value into the torque set value required by the second controller, and inputs it to the second controller through the CAN bus; The second controller receives a torque given value sent by the speed control box (6), performs torque closed-loop control according to the torque given value, and performs real-time adjustment according to the real-time torque given value, ultimately achieving the same output torque of the first controller and the second controller; The rotor speed setting device (8) outputs a speed control signal to the speed control box (6); The speed control box (6) sends a speed instruction to the first controller (12) via the CAN bus, and the speed control box (6) determines whether two electric power systems need to operate simultaneously based on actual speed requirements; When the speed control box (6) determines that the actual speed requirement can be provided by the electric power system 1 alone, the speed control box (6) will issue a shutdown instruction to the second controller to stop the electric power system 2 from running; According to the actual speed requirement of the system, the speed control box (6) determines that two electric power systems need to be operated simultaneously, and sends a controller start instruction to the second controller to put the electric power system 2 into operation, and at the same time receives the actual current value of the motor sent by the first controller, and converts the obtained actual current value of the first controller into an actual torque given value after current averaging and filtering using the least squares polynomial parameter identification method, and sends the given torque value as an input to the second controller through the CAN bus, so that the electric power system 2 operates according to the given torque value to provide power for the aircraft.

2. A dual-redundant electric power system suitable for a light electric helicopter according to claim 1, characterized in that: The invention also includes an auxiliary power supply battery (101), wherein the auxiliary power supply battery (101) is used to supply power to a power integrated display (9), a BMS (10), a first controller (12), a second controller (22), a high-voltage box (5), a rotor speed setting device (8), and a speed control box (6).

3. A dual-redundant electric power system suitable for a light electric helicopter according to claim 1, characterized in that: Motor speed sensors are provided at the ends of the first motor (11) and the second motor (21), and are connected to the first controller (12) and the second controller (22) via the motor speed sensors. The frequency of the three-phase AC power output by the controller is used to adjust the rotor speed in real time. The motor speed sensor feeds back the actual motor speed and position information to the corresponding controller in real time, forming a closed-loop control of the motor speed, thereby controlling the operating conditions of the aircraft.

4. A dual-redundant electric power system suitable for a light electric helicopter according to claim 3, characterized in that: The controller uses a rotor speed setting device (8) to set a desired rotor speed as a control target; The motor speed sensor monitors the actual rotor speed in real time and feeds it back to the controller. The feedback speed value is compared with the given speed value, and a PI speed closed-loop control strategy is adopted to achieve precise follow-up control of the rotor speed by the controller.

5. The dual-redundant electric power system for a light electric helicopter according to claim 1, characterized in that: The least squares polynomial parameter identification method: Given the controller output current, the controller output current can be expressed by a polynomial, namely: in: is the current output by the first controller (12) at time t, a i are the coefficients of the polynomial, and n represents the order of the polynomial; The sum of squares of the minimized residuals between the actual collected value and the polynomial predicted value is minimized, and we can get: Among them, I(t i ) is the actual observed value, is the polynomial prediction value; The minimum residual sum of squares can be obtained from formula (2): The derivative of the minimum residual sum of squares with respect to the polynomial coefficients is zero; By calculating formula (4), we can get a0, a1, a2, ..., a n-1 , a n , thereby obtaining a characteristic curve of the output current of the first controller (12); The motor output torque is proportional to the controller output current, and it can be obtained from formula (1) Among them, k M is the torque coefficient; M(t) is the motor output torque.

6. A dual-redundant electric power system suitable for a light electric helicopter according to claim 1, characterized in that: The high-voltage box (5) includes a fast fuse and a contactor. The high-voltage box controls the on-off of the contactor according to the received on-off instruction of the contactor, thereby controlling the power DC power supply required by the first controller (12) and the second controller (22).

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