Electric driving system for electric propulsion system of electric aircraft

Abstract

The invention belongs to the technical field of electric aircraft electric propulsion system design, and particularly relates to an electric drive system for an electric aircraft electric propulsion system, which comprises two inverter controllers and a cascade motor, a motor control instruction sent by a flight control system is received by an inversion controller, high-voltage direct current on the aircraft is converted into alternating current to be output, the output alternating current drives a cascade motor to drive a propeller to rotate, and the propeller rotates to generate flight thrust to provide flight power for the electric aircraft; the rotating speed and torque of the propeller driven by the cascade motor are realized by controlling the frequency and amplitude of the output alternating current through the inverter controller; the cascade motor is composed of two matched motors, 500 kW high-power output is achieved in a coaxial cascade mode, each motor is of a double-winding structure, and winding redundancy is achieved. The working direct-current bus voltage of the inverter controller is 400V-800V, the operation of the cascade motors is controlled through voltage inversion, and two motors are matched to realize coordinated control of multiple motors and multiple windings.

Description

Technical Field

[0001] This application belongs to the field of electric propulsion system design technology for electric aircraft, and specifically relates to an electric drive system for electric propulsion systems of electric aircraft. Background Technology

[0002] Driven by the national "dual-carbon" strategy, the "Medium- and Long-Term Plan for the Development of Low-Altitude Economy," released in 2021, further points out that electric propulsion systems will become a new driving force for the development of the low-altitude economy. As an emerging technology, electric propulsion can not only significantly reduce carbon emissions and pollutant generation, but also effectively reduce aircraft operating costs and improve range.

[0003] In recent years, with the rapid development of emerging industries such as emergency rescue and low-altitude transportation, higher demands have been placed on electric propulsion technology. As a core component of electric propulsion systems, electric drive systems still face technical bottlenecks such as low power density and insufficient output power. In order to meet the future development needs of electric aircraft such as single-propeller, twin-rotor, quadcopter, multi-rotor eVTOL, and tiltrotor aircraft, the application of high-power, high-safety, redundant, and fault-tolerant electric drive systems in the aviation field is particularly important.

[0004] The industry has conducted relatively thorough research on electric drive systems for electric aircraft propulsion systems, with prototypes gradually increasing in power and undergoing on-board verification on different flight platforms. For example, MagniX has developed a series of electric propulsion units (EPUs) for aviation, involving related patents such as US2020381985A1 for a high-torque electric propulsion system and US20200382041_A1 for a high-voltage converter used for power supply. The Magni650 electric propulsion unit system includes four MagniDrive 100 controllers, one Magni650 motor, and a closed-loop integrated thermal management system. Its 4-channel, 3-phase architecture increases reliability through redundancy technology. In its system structure, the motor has two coaxially connected rotors, and each rotor's stator armature winding uses two sets of three-phase windings. The electromagnetic interaction between the two sets of three-phase windings is largely independent, reducing the impact of modular three-phase controllers on the torque output control of each set of three-phase windings in the motor. However, it only describes one electric propulsion system architecture and application examples of dual-channel and quad-channel systems, but does not explain the operating characteristics, redundancy configuration methods and applications of dual-phase / multi-phase electric drive systems.

[0005] This application is made in view of the aforementioned technical deficiencies. Summary of the Invention

[0006] The purpose of this application is to provide an electric drive system for an electric propulsion system of an electric aircraft, which is a novel high-power, high-safety, redundant, and fault-tolerant electric drive system suitable for electric propulsion systems of electric aircraft.

[0007] The technical solution of this application is:

[0008] An electric drive system for an electric propulsion system of an electric aircraft includes two inverter controllers and cascaded electric motors;

[0009] The inverter controller receives motor control commands from the flight control system, converts the high-voltage DC power on the aircraft into AC power output, and drives the cascaded motors to rotate the propeller. The rotation of the propeller generates flight thrust, providing flight power for the electric aircraft.

[0010] The rotational speed and torque of the propeller driven by the cascaded motors are controlled by the inverter controller to control the frequency and amplitude of the output AC power.

[0011] The cascaded motor consists of two matching motors, which achieve a high power output of 500kW through coaxial cascading. Each motor adopts a dual-winding structure to achieve winding redundancy.

[0012] The inverter controller operates at a DC bus voltage of 400V-800V and controls the operation of cascaded motors through voltage inversion. It is equipped with two units to achieve coordinated control of multiple motors and multiple windings.

[0013] Cascaded motors are equipped with two rotary transformers, which are used independently by a single winding of each motor and serve as backups for each other, so that each winding of the motor has an independent working function.

[0014] The inverter controller has at least two CAN communication interfaces. One CAN communication interface is used to receive motor control commands from the flight control system to control the output speed or torque of the cascaded motors. The other CAN communication interface is used for communication between inverter controllers, for coordinated control of multiple motors and multiple windings, and also serves as a backup communication interface for sensor information on the motor shaft position.

[0015] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the inverter controller is a dual three-phase inverter controller.

[0016] The motor is a dual three-phase motor.

[0017] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the stator armature winding of the motor is distributed in a double three-phase Y-type structure, with the neutral lines of the two windings each led out and short-circuited to form a 6-phase motor or each being independent without electrical connection.

[0018] The stator armature windings of the motor are divided into two symmetrical groups that share the same rotor. The windings in the same phase are 180 degrees apart and have the same phase, forming a half-double three-phase armature winding.

[0019] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the inverter controller includes a power supply board, a main power circuit, and a control board.

[0020] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the power supply board receives an input voltage of +28V or high voltage DC, with the inputs serving as backups for each other. The DC / DC module converts the +28V power supply into different voltage levels of +5V and ±15V as the operating voltage of the control board and the main power circuit. The outputs of different voltage levels are configured with two isolated outputs to the channel circuits of different phases.

[0021] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the main power circuit inverts the high-voltage DC power supply on the aircraft into AC power to drive the electric motor.

[0022] The main power circuit includes two three-phase inverter circuits;

[0023] The three-phase inverter circuit includes a support capacitor, an absorption capacitor, a bus voltage sensor, a three-phase current sensor, a bus current sensor, a three-phase inverter power bridge and drive circuit, and a discharge circuit.

[0024] The three-phase inverter power bridge selects high-voltage, high-power IGBT switching transistors, and the DC bus voltage requires a wide input range of 0-800V;

[0025] Each three-phase inverter circuit is equipped with a current sensor at both the positive and negative terminals. The differential current detection of the positive and negative terminals of the DC bus of the three-phase inverter circuit is achieved through two current sensors, which is used for differential current protection.

[0026] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the main power circuit controls the operation of the inverter controller in different working modes through a mode switching switch, including three modes: dual-channel parallel operation, dual-channel series operation, and single-channel independent operation.

[0027] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the control board is used for position signal fault tolerance and dual three-phase cascaded motor control, including the control core circuit.

[0028] The core control circuit adopts a DSP+FPGA architecture, and the position decoding circuit has two paths;

[0029] The core control circuit collects and analyzes signals such as motor position, phase current, phase voltage, bus voltage, and temperature of the main power circuit. It adopts a vector control strategy and generates PWM signals through closed-loop control of speed and current. The main power circuit inverts the high-voltage DC power supply into variable frequency AC power according to the PWM drive signal to control the operation of the motor.

[0030] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the DSP has a dual-core CPU. The first core CPU1 is used to complete the main program calculation and protection, and the second core CPU2 is used to receive instructions from the upper control system and the information transmission between the modular controller. The control instructions are transmitted to the first core CPU1 through the inter-core communication module IPC to realize remote data transmission and monitoring.

[0031] The FPGA completes sampling from two 16-channel A / D chips and sampling of axis angle information from two R / D axis angle conversion chips, as well as providing protection functions for multiple voltage, current, and temperature signals.

[0032] According to at least one embodiment of this application, in the electric drive system of the electric propulsion system of the electric aircraft described above, the control core circuit has two CAN communication channels, wherein the first CAN communication channel is used to receive control commands from the upper layer, and the second CAN communication channel is used for control signal transmission between inverter controllers.

[0033] This application has at least the following beneficial technical effects:

[0034] An electric drive system for electric propulsion systems of electric aircraft is provided, which is equipped with multiple dual-redundancy designs such as parallel redundancy of motor windings, cascade redundancy of motors, redundancy of inverter controller, redundancy of position sensor, and redundancy of communication, resulting in high safety. Attached Figure Description

[0035] Figure 1 This is a schematic diagram of the electric drive system for the electric propulsion system of an electric aircraft provided in the embodiments of this application;

[0036] Figure 2 This is a schematic diagram of the dual three-phase inverter controller provided in an embodiment of this application;

[0037] Figure 3 This is a schematic diagram of the cascaded motors provided in the embodiments of this application;

[0038] Figure 4 This is a schematic diagram of the half-double three-phase armature winding connection provided in the embodiments of this application;

[0039] Figure 5 This is a block diagram of the overall architecture of the dual three-phase inverter controller provided in the embodiments of this application;

[0040] Figure 6 This is a schematic diagram of the control power supply architecture provided in the embodiments of this application;

[0041] Figure 7 This is a schematic diagram of the main power circuit provided in an embodiment of this application;

[0042] Figure 8 This is a schematic diagram of the working principle of the control board provided in the embodiments of this application;

[0043] Figure 9 This is a basic layout diagram of the control board functional modules provided in the embodiments of this application;

[0044] Figure 10 This is a diagram of the multi-core parallel control strategy provided in the embodiments of this application;

[0045] Figure 11 This is a flowchart of the main program of the high-power electric drive system provided in the embodiments of this application;

[0046] Figure 12 This is a flowchart of the A / D and R / D sampling process provided in the embodiments of this application;

[0047] Figure 13 This is a diagram of the electric propulsion system architecture for a single-propeller electric aircraft provided in an embodiment of this application;

[0048] Figure 14 This is an independent architecture diagram of the electric propulsion system for a dual-propeller electric aircraft provided in an embodiment of this application;

[0049] Figure 15 This is a redundant architecture diagram of the electric propulsion system for a dual-propeller electric aircraft provided in an embodiment of this application;

[0050] Figure 16 This is a diagram of the dual-power redundant architecture of the electric propulsion system for a dual-propeller electric aircraft provided in this application embodiment;

[0051] Figure 17 This is a closed-loop redundant architecture diagram of the electric propulsion system for a quad propeller electric aircraft provided in an embodiment of this application.

[0052] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation

[0053] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings, and other related parts can be referred to the general design.

[0054] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.

[0055] Furthermore, the terms indicating location used in the description of this application are only used to indicate relative directions or positional relationships. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. It should also be noted that, unless otherwise explicitly specified and limited, terms such as "installation" and "connection" used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0056] An electric drive system for an electric propulsion system of an electric aircraft, such as Figure 1 As shown, its basic function is to convert electrical energy on the aircraft into mechanical energy. The inverter controller receives the motor control commands issued by the flight control system, converts the high-voltage DC power on the aircraft into AC power output, and the output AC power drives the cascaded motor to drive the propeller to rotate. The rotation of the propeller generates flight thrust, providing flight power for the electric aircraft. At this time, the rotation speed and torque of the propeller driven by the cascaded motor are controlled by the inverter controller to control the frequency and amplitude of the output AC power.

[0057] The interconnection between the electric drive system and the aircraft is as follows:

[0058] It is installed on the aircraft using bolts.

[0059] High-voltage DC voltage and controller operating voltage are input from the aircraft power system via the input connector;

[0060] Receive motor control commands from the flight control system via the input connector;

[0061] The rotor of the cascaded motor is coaxially connected to the propeller by bolts.

[0062] The electric drive system mainly consists of two inverter controllers and cascaded motors. The inverter controllers are dual three-phase inverter controllers, such as... Figure 2 As shown.

[0063] Cascaded motors, such as Figure 3 As shown, the high-torque output is directly connected to the power load to drive the propeller, providing flight power for the aircraft. The cascaded motor consists of two matching motors, achieving a 500kW high-power output through coaxial cascading. Each motor adopts a dual-winding structure to achieve winding redundancy (the motor windings are independent of each other, representing parallel redundancy; two motors connected coaxially represent series redundancy). The inverter controller operates on a DC bus voltage of 400V-800V and controls the operation of the cascaded motors through voltage inversion. Due to the dual-winding motor cascaded structure, the inverter controller needs to achieve coordinated control of multiple motors and windings; therefore, two units are required.

[0064] In a cascaded motor system, two motors are dual three-phase motors. The cascaded motor system is equipped with two rotary transformers, which allow each motor's single winding to be used independently and serve as a backup for the other. This enables each winding of the motor to have an independent working function. When a fault occurs, the faulty winding can be directly disconnected, while the other windings continue to operate normally, giving the electric drive system a certain degree of fault tolerance.

[0065] The inverter controller has at least two CAN communication interfaces. One CAN communication interface is used to receive motor control commands from the flight control system to control the speed or torque output of the cascaded motors. The other CAN communication interface is used for communication between inverter controllers, for coordinated control of multiple motors and windings, and also serves as a backup communication interface for sensor information on the motor shaft position. When one resolver fails while another sensor is operating normally, the controller of the faulty winding receives sensor position information from the controller of the normal winding via the CAN communication interface. The faulty winding directly uses the position information of the normal sensor transmitted via CAN communication for control output, realizing the redundancy and fault tolerance function of the sensor.

[0066] Both the inverter controller and the motor adopt a modular design. Each inverter controller can be directly matched with one motor to form the smallest electric drive system unit, or multiple inverter controllers and motors can form electric drive systems with different architecture types.

[0067] The electric propulsion system of the aforementioned electric aircraft is equipped with multiple dual-redundancy designs, including parallel redundancy of motor windings, cascade redundancy of motors, redundancy of inverter controllers, redundancy of position sensors, and redundancy of communication, which ensures high safety.

[0068] In a cascaded motor, the stator armature windings are arranged in a double three-phase Y-type configuration. The neutral lines of the two windings are each led out separately. They can be short-circuited to form a six-phase motor or remain independent without electrical connection, thus becoming a double three-phase motor. To minimize electromagnetic coupling between the two windings, the stator armature windings are divided into two symmetrical groups sharing the same rotor. The windings of the same phase are 180 degrees apart but in phase, forming a symmetrical double three-phase armature winding. The connection method is as follows: Figure 4 As shown.

[0069] The aforementioned motor structure is similar to a dual-stator motor structure, achieving redundancy in the motor windings. Each of the two stators forms a three-phase Y-type structure, allowing for independent control using an inverter controller. Under the same speed and torque output conditions, the current output value of the inverter controller's power devices can be reduced, theoretically by half. When one winding is not in operation, the motor still retains half of its output torque capacity.

[0070] Inverter controller principle as follows Figure 5 As shown, it mainly consists of structural components such as a power supply board, main power circuit, control board, and drive board assembly. The chassis assembly consists of a chassis and a heat dissipation section. In addition to supporting the functions of the various components of the inverter controller, it also needs to provide heat dissipation for the power components to meet the heat dissipation requirements of the electric drive system during operation.

[0071] The power board uses a DC / DC module to convert the +28V power supply to different voltage levels such as +5V and ±15V, which are used as the operating voltage for the control board and the main power circuit.

[0072] The power board outputs +5V and ±15V. +5V powers the control core circuitry of the DSP and FPGA, while ±15V primarily powers the voltage and current sensors, some sampling and conditioning circuits, and the main power circuit driver board. The power board can accept input voltages of +28V or high-voltage DC, with inputs serving as backups to achieve redundancy in the control power input. The power board requires different DC / DC modules to achieve different voltage levels. Each voltage level is configured with two isolated outputs to different phase circuits, achieving power output isolation and redundancy. The control power supply architecture is as follows: Figure 6 As shown.

[0073] Based on the functional requirements of the electric drive system, the main function of the main power circuit is to invert the aircraft's high-voltage DC power supply (typically 540VDC) into AC power to drive the motor. The main power circuit consists of two three-phase inverter circuits, a mode switching switch, a soft-start circuit, and an input voltage sensor. Each three-phase inverter circuit comprises a support capacitor, an absorption capacitor, a bus voltage sensor, a three-phase current sensor, a bus current sensor, a three-phase inverter power bridge, a drive circuit, and a discharge circuit. Figure 7As shown. The three-phase inverter power bridge selects high-voltage, high-power IGBT switching transistors. The DC bus voltage requires a wide input range of 0-800V, typically 540VDC. The range of the DC voltage sensor meets the voltage range requirements, and the range of the current sensor should meet the short-time overload current measurement requirements as much as possible.

[0074] In the main power circuit, each three-phase inverter circuit is equipped with a current sensor at both the positive and negative terminals. The differential current detection of the positive and negative terminals of the DC bus of the three-phase inverter circuit can be achieved through two current sensors, which is used for differential current protection to ensure fault isolation of each three-phase inverter circuit and guarantee safety.

[0075] By controlling the mode switching switches (S3, S4, S5) in the main power circuit, the inverter controller can operate in different modes, including dual-channel parallel operation, dual-channel series operation, and single-channel independent operation. When switches S3 and S5 are closed, it operates in dual-channel parallel mode; when only S4 is closed, it operates in dual-channel series mode; when S4 and S3 are open and S5 is closed, channel 2 operates independently; when S4 and S5 are open and S3 is closed, channel 1 operates independently.

[0076] The above three modes can also be combined with motor and operating modes to achieve multiple system operating modes. For example, when S6 is open, the motor center is disconnected, which is a dual three-phase electric drive system; when S6 is closed, the motor center is connected, which is a multi-phase electric drive system. Other possible coordination modes and multi-motor three-phase channel coordinated control can also be achieved. The above S6 switch can be controlled through the controller IO interface.

[0077] By controlling the mode switching switch, fault isolation of two three-phase inverter circuits can be achieved. In parallel operation mode, if a fault occurs in a certain inverter channel, the faulty channel can be electrically isolated by disconnecting switch S3 or S5. After the faulty channel is cleared, the remaining channels and motor windings can still work independently, which has high fault tolerance and improves system safety.

[0078] The soft power-on process in the main power circuit described above is as follows:

[0079] When a high-voltage DC power supply is input, and the controller input voltage sensor detects that the voltage has reached the operable threshold, the inverter controller first controls the operation of switches S3, S4, and S5 according to the given instructions to switch the mode in real time. Then, it controls switch S1 to perform soft power-on for the bus support capacitor of the three-phase inverter circuit. When the inverter controller detects that the voltage of the support capacitor has reached the specified threshold through the bus voltage sensor, it controls switch S2 to close, and the soft power-on process ends.

[0080] The three-phase inverter circuit in the main power circuit described above can be expanded to multiple channels to achieve inverter output of more phases or more three-phase channels.

[0081] The control board mainly consists of a core control circuit, two 16-channel A / D sampling circuits, and two shaft angle conversion circuits, used for position signal fault tolerance and dual three-phase cascaded motor control. The basic working principle of the control board is as follows: Figure 8 As shown.

[0082] The control board uses a DSP+FPGA architecture as its control core. It collects and analyzes signals such as motor position, phase current, phase voltage, bus voltage, and temperature of the main power circuit. Employing a vector control strategy, it generates PWM signals to drive the three-phase inverter power bridge through closed-loop control of speed and current. The main power circuit inverts the high-voltage DC power into variable-frequency AC power based on the PWM drive information, controlling the motor operation. It features two 16-channel sampling circuits, two resolver excitation circuits, and a position sampling circuit. Each sampling circuit is independent and provides redundancy for the others.

[0083] The core control circuit adopts a DSP (TMS320F28379) + FPGA (XC7A200T) architecture, with a dual-channel position decoding circuit. The DSP has a dual-core CPU, which can improve control calculation efficiency through proper configuration. It features 24 independent PWM output pins, allowing for simultaneous control of up to four sets of three-phase windings by adding a main power circuit. There are at least two CAN communication channels. The first CAN communication channel is mainly used to receive control commands from the upper layer, while the second CAN communication channel is used for control signal transmission between modular inverter controllers and can also serve as redundancy between the inverter controller and the upper-level control. The RS485 and RS232 ports are backup communication ports for the modular inverter controller. The control board functional module layout is as follows: Figure 9 As shown.

[0084] The control board outputs 24 PWM signals to control the upper and lower bridge arms of the IGBT power transistors. Each group of six bridge arms can control the operation of three-phase, five-phase, and six-phase permanent magnet synchronous motors, dual three-phase permanent magnet synchronous motors, and single three-phase cascaded motor systems. Furthermore, this circuit can expand the number of three-phase inverter bridges in the main power circuit to control two dual three-phase permanent magnet motors, two six-phase permanent magnet synchronous cascaded motors, or one twelve-phase permanent magnet synchronous motor.

[0085] The dual-core DSP architecture in the control core circuit utilizes its first core, CPU1, to perform main program calculations and protection. The second core, CPU2, receives instructions from the upper-level control system and facilitates information exchange between the modular controller and the CPU. Control instructions are transmitted to the first core, CPU1, via the inter-core communication module (IPC), enabling remote data transmission and monitoring. The FPGA primarily performs sampling from two AD7616 16-channel A / D chips and two AD2S1210 R / D axis-angle conversion chips, while also providing multi-channel voltage, current, and temperature protection functions.

[0086] The dual-core DSP employs a multi-core parallel control strategy to improve the operating efficiency of the multi-channel controller. Its operation and control strategy is as follows: Figure 10 As shown.

[0087] When using a single control board to control multiple motor windings, high CPU computational efficiency is required, especially when communicating with the upper-level system. Computational resources need to be allocated reasonably to avoid communication functions consuming excessive main program computation time. The key program of the entire system is the control section of the switching transistors. The timing requirements for its signal input, control algorithm, and output control are extremely high. If the system switching and control cycles are 10kHz and the PWM update cycle is 0.1ms, the short control time requires the designer to rationally and effectively divide system tasks to ensure the system completes all tasks efficiently, stably, and completely. The design utilizes the first CPU core (CPU1) to handle main program calculations and protection, while the second CPU core (CPU2) receives instructions from the upper-level control system and facilitates information transmission between the modular controller. Control instructions are transmitted to the first CPU core (CPU1) via the IPC module, enabling remote data transmission and monitoring.

[0088] Furthermore, the losses in high-power controllers are mainly due to the frequent switching of power devices. Therefore, when implementing high-power drive control, it is necessary to appropriately reduce the switching frequency of the power devices, generally around 6kHz. Due to the relatively lower switching frequency, the sinusoidal nature of the motor phase current deteriorates, and the main program waits a considerable amount of time after calculating the PWM switching time and updating the corresponding register, resulting in a lag between the updated PWM information and the motor state. This phenomenon is exacerbated in high-speed motors, affecting the motor system control. For high-power drive motors, when the motor speed is relatively low, the current and shaft angle information can be sampled and calculated multiple times within a single PWM interrupt cycle, updating the speed and current loop PI outputs multiple times to approximate the current state of the motor as closely as possible. The main program workflow of a high-power electric drive system is as follows: Figure 11 As shown.

[0089] When controlling a high-speed motor, it is not possible to support multiple calculations of speed loop and current loop PI data. However, it is possible to perform multiple samplings of voltage, current, and shaft angle transformation information and calculate the PWM conduction time using data close to the current motor state.

[0090] The main functions of the FPGA, such as A / D sampling of DC bus voltage and current, phase current, and R / D axis angle signal sampling, are all implemented based on FPGA programming. A / D and R / D axis angle transformations are performed in parallel, and their working principles are basically the same. The basic program flowchart is as follows: Figure 12 As shown.

[0091] The electric drive system for electric propulsion systems of electric aircraft disclosed in the above embodiments is equipped with multiple dual-redundancy designs, such as parallel redundancy of motor windings, cascade redundancy of motors, redundancy of inverter controllers, redundancy of position sensors, and redundancy of communication, which ensures high safety.

[0092] The electric drive system for electric propulsion systems of electric aircraft disclosed in the above embodiments has multiple communication functions, overvoltage, overcurrent, and temperature protection functions, and fault-tolerant functions such as multi-motor coordinated control and the ability of the remaining windings to work independently after the faulty winding is removed.

[0093] The electric drive system for electric propulsion systems of electric aircraft disclosed in the above embodiments adopts multiple dual-redundancy designs, such as redundant backup of position sensors, resulting in high safety. Both the inverter controller and the electric motor in the system employ a modular design. By flexibly expanding and combining the inverter controller and electric motor in the electric drive system, and coordinating with communication networking, an electric drive system architecture adaptable to different power levels, voltage levels, and fault-tolerant operating modes can be formed to meet the application requirements of high-power electric propulsion systems for different types of electric aircraft.

[0094] First application example:

[0095] The electric drive system for electric propulsion systems of electric aircraft disclosed in the above embodiments is applied to electric propulsion systems of single-propeller electric aircraft, specifically as follows: Figure 13 As shown.

[0096] Two dual three-phase motors are cascaded in series and connected to the propeller of the electric aircraft. They receive commands from the flight control system via the CAN1 communication bus. The two dual three-phase inverter controllers control the cascaded motors to achieve power output. The two inverter controllers exchange data via the CAN2 communication bus to perform power balancing control and power distribution between different windings of the two motors. At the same time, they monitor the status of each channel of the system to achieve fault isolation and switching of operating status. This enables multi-channel redundancy of the power circuit, ensuring continuous power output and achieving high flight safety.

[0097] In practical applications, the inverter controller can be configured to set the operating states of multiple controller inverter output channels and motor winding channels, allowing them to operate in either active or standby mode. The channel status can be configured according to actual power requirements. When a channel in active mode fails, the channel operating in standby mode can serve as a backup redundancy for the electric drive system's power output, achieving system fault channel isolation and improving the redundancy and fault tolerance performance of the system's power output.

[0098] Depending on the power requirements, one or more minimum electric drive system units can be cascaded to achieve power output and control of a single propeller electric propulsion system. When multiple minimum electric drive system units drive the propeller simultaneously, an imbalance in output torque occurs among them, affecting system efficiency. Therefore, coordinated control of the multiple electric drive systems is necessary. Since each minimum electric drive system unit can operate independently, the coordinated control function of the multiple electric drive systems is implemented in a single dual three-phase controller. One of these controllers acts as the master controller, receiving commands from the flight control system via the CAN1 communication bus to achieve dual closed-loop control of speed and torque. The output value of the speed loop PI regulator is then distributed to other modular controllers via the CAN2 communication bus as the torque control setpoint. Other modular controllers track the speed in real time. When the master controller fails, such as a position sensor failure or a CAN1 communication failure, the other controllers act as redundant backups to ensure normal system operation.

[0099] Second application example:

[0100] The electric drive system described above, which is used in single-propeller electric propulsion systems, can be extended to multi-propeller electric aircraft or multi-rotor eVTOL electric aircraft.

[0101] Independent architecture of electric propulsion system for twin-propeller electric aircraft, such as Figure 14 As shown, two propellers are driven by two minimum electric drive system units. The two minimum electric drive system units can receive flight control system commands through the communication bus CAN1 to achieve independent operation, or they can interact with each other through the communication bus CAN2 to achieve coordinated control of power output between the two propellers.

[0102] Third application example:

[0103] In the electric drive system architecture described above for a twin-propeller electric propulsion system, if both inverter output channels of the inverter controller in one of the smallest electric drive system units fail, one propeller will have no power output, affecting the normal flight of the electric aircraft.

[0104] The two three-phase output channels of the two inverter controllers are cross-connected to the two sets of three-phase windings of the two dual-winding motors, forming a configuration as follows: Figure 15The diagram shows a redundant architecture for the electric propulsion system of a twin-propeller electric aircraft.

[0105] Inverter output channel 1 of inverter controller 1 is connected to one three-phase winding of dual-winding motor 1, inverter output channel 2 is connected to one three-phase winding of dual-winding motor 1; inverter output channel 3 of inverter controller 2 is connected to the other three-phase winding of dual-winding motor 1, and inverter output channel 4 is also connected to the other three-phase winding of dual-winding motor 2.

[0106] When the power requirements are met, the direct-connected inverter output channels in the two minimum electric drive system units can be configured as active operation, while the cross-connected inverter output channels can be configured as standby operation. Data exchange is performed through the CAN2 communication bus to monitor the inverter output channels in active operation between the systems. This ensures that if a channel fails, the standby operation channel can be promptly connected as a backup, achieving redundant backup of the power output of the electric drive system and improving the redundancy of the electric propulsion system of the twin-propeller electric aircraft.

[0107] If each minimum electric drive system unit also has its own independent high-voltage DC power supply, power supply redundancy of the electric drive system can be achieved, forming a dual-power redundancy architecture for the electric propulsion system of a twin-propeller electric aircraft. Figure 16 As shown, this allows the electric propulsion system to have higher power output redundancy.

[0108] Fourth application example:

[0109] The dual-power redundant architecture of the electric propulsion system for twin-propeller electric aircraft can be further extended to more propeller-driven electric aircraft or multi-rotor eVTOLs. The closed-loop redundant architecture of the electric propulsion system for quad-propeller electric aircraft is as follows: Figure 17 As shown, other rotorcraft electric aircraft have similar electric propulsion architectures.

[0110] The four propellers are driven by four minimum electric drive system units. The inverter output channels and motor windings of the four minimum electric drive system units are also cross-connected, forming a closed-loop redundancy of the inverter output channels of each minimum electric drive system unit, that is, the closed-loop redundancy of the power output system of the electric propulsion system of the electric aircraft.

[0111] Each propeller can be expanded to be driven by multiple cascaded minimum electric drive systems, improving the redundancy and fault tolerance of the electric drive system, and further ensuring the flight safety of electric aircraft.

[0112] In the above application examples, the high-voltage DC power supply mentioned can be one or more backups or independent power supplies, such as batteries, high-voltage DC generators or fuel cells, etc. The CAN communication bus architecture can be of various types, such as star topology architecture, linear topology architecture, etc., and the communication form is not limited to CAN communication. The motor and inverter controller are not limited to the motor and inverter controller included in the electric drive system of this invention, but can be single-phase or three-phase inverters or single-phase or three-phase motor cascades, etc.

[0113] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the protection scope of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the protection scope of this application.

Claims

1. An electric drive system for an electric propulsion system of an electric aircraft, characterized in that, Includes two inverter controllers and cascaded motors; The inverter controller receives motor control commands from the flight control system, converts the high-voltage DC power on the aircraft into AC power output, and drives the cascaded motors to rotate the propeller. The rotation of the propeller generates flight thrust, providing flight power for the electric aircraft. The rotational speed and torque of the propeller driven by the cascaded motors are controlled by the inverter controller to control the frequency and amplitude of the output AC power. The cascaded motor consists of two matching motors, which achieve a high power output of 500kW through coaxial cascading. Each motor adopts a dual-winding structure to achieve winding redundancy. The inverter controller operates at a DC bus voltage of 400V-800V and controls the operation of cascaded motors through voltage inversion. It is equipped with two units to achieve coordinated control of multiple motors and multiple windings. Cascaded motors are equipped with two rotary transformers, which are used independently by a single winding of each motor and serve as backups for each other, so that each winding of the motor has an independent working function. The inverter controller has at least two CAN communication interfaces. One CAN communication interface is used to receive motor control commands from the flight control system to control the output speed or torque of the cascaded motors. The other CAN communication interface is used for communication between inverter controllers, for coordinated control of multiple motors and multiple windings, and also serves as a backup communication interface for sensor information on the motor shaft position.

2. The electric drive system for an electric propulsion system of an electric aircraft according to claim 1, characterized in that, The inverter controller is a dual three-phase inverter controller; The motor is a dual three-phase motor.

3. The electric drive system for an electric propulsion system of an electric aircraft according to claim 2, characterized in that, The stator armature windings in the motor are distributed in a double three-phase Y-type structure, with the neutral lines of the two windings each led out and short-circuited to form a 6-phase motor or each winding being independent without electrical connection. The stator armature windings of the motor are divided into two symmetrical groups that share the same rotor. The windings in the same phase are 180 degrees apart and have the same phase, forming a half-double three-phase armature winding.

4. The electric drive system for an electric propulsion system of an electric aircraft according to claim 3, characterized in that, The inverter controller includes a power supply board, a main power circuit, and a control board.

5. The electric drive system for an electric propulsion system of an electric aircraft according to claim 4, characterized in that, The power board receives an input voltage of +28V or high voltage DC, with the inputs serving as backups for each other. It uses a DC / DC module to convert the +28V power supply to different voltage levels of +5V and ±15V, which are used as the operating voltage for the control board and the main power circuit. The outputs of different voltage levels are configured with two isolated outputs to the channel circuits of different phases.

6. The electric drive system for an electric propulsion system of an electric aircraft according to claim 5, characterized in that, The main power circuit inverts the high-voltage DC power supply on the aircraft into AC power to drive the electric motor. The main power circuit includes two three-phase inverter circuits; The three-phase inverter circuit includes a support capacitor, an absorption capacitor, a bus voltage sensor, a three-phase current sensor, a bus current sensor, a three-phase inverter power bridge and drive circuit, and a discharge circuit. The three-phase inverter power bridge selects high-voltage, high-power IGBT switching transistors, and the DC bus voltage requires a wide input range of 0-800V; Each three-phase inverter circuit is equipped with a current sensor at both the positive and negative terminals. The differential current detection of the positive and negative terminals of the DC bus of the three-phase inverter circuit is achieved through two current sensors, which is used for differential current protection.

7. The electric drive system for an electric propulsion system of an electric aircraft according to claim 6, characterized in that, The main power circuit controls the operation of the inverter controller in different working modes through a mode switching switch, including three modes: dual-channel parallel operation, dual-channel series operation, and single-channel independent operation.

8. The electric drive system for an electric propulsion system of an electric aircraft according to claim 7, characterized in that, The control board is used for position signal fault tolerance and dual three-phase cascaded motor control, including the core control circuit. The core control circuit adopts a DSP+FPGA architecture, and the position decoding circuit has two paths; The core control circuit collects and analyzes signals such as motor position, phase current, phase voltage, bus voltage, and temperature of the main power circuit. It adopts a vector control strategy and generates PWM signals through closed-loop control of speed and current. The main power circuit inverts the high-voltage DC power supply into variable frequency AC power according to the PWM drive signal to control the operation of the motor.

9. The electric drive system for an electric propulsion system of an electric aircraft according to claim 8, characterized in that, The DSP has a dual-core CPU. The first core CPU1 is used to complete the main program calculation and protection, while the second core CPU2 is used to receive instructions from the upper control system and to transmit information between the modular controller. The control instructions are transmitted to the first core CPU1 through the inter-core communication module IPC to realize remote data transmission and monitoring. The FPGA completes sampling from two 16-channel A / D chips and sampling of axis angle information from two R / D axis angle conversion chips, as well as providing protection functions for multiple voltage, current, and temperature signals.

10. The electric drive system for an electric propulsion system of an electric aircraft according to claim 9, characterized in that, The control core circuit has two CAN communication channels. The first CAN communication channel is used to receive control commands from the upper layer, and the second CAN communication channel is used for the transmission of control signals between inverter controllers.

Images