A tilt-type EVTOL flight control system

Abstract

The utility model discloses a kind of tilting type EVTOL flight control system, it includes redundancy flight control computer device, power device, several wing surface rudder control devices, control device, data link device, controlling means, several main communication buses and several slave communication buses, redundancy flight control computer device includes several flight control computers, power device includes several tilting power devices and several take-off and landing power devices;Every main communication bus or every slave communication bus is connected with all flight control computers, tilting power device, take-off and landing power device and wing surface rudder control device are connected with at least one main communication bus and at least one slave communication bus, redundancy flight control computer device, power device, wing surface rudder control device inside are reserved with matching resistance.The utility model is designed by non-similar double-redundancy bus, precise redundancy management and real-time communication optimization, constructs highly safe, reliable, real-time delivery and easily maintained flight control system.

Description

Technical Field

[0001] This utility model belongs to the field of aircraft technology, and in particular relates to a tilt-type EVTOL flight control system. Background Technology

[0002] Since EVTOL aircraft need to carry cargo or people over cities and are primarily controlled by fly-by-wire, the safety and reliability requirements for the aircraft are extremely stringent. Traditional flight control systems cannot meet the requirements of EVTOL aircraft, and a flight control system with EVTOL characteristics needs to be designed.

[0003] A control architecture for an eVTOL actuator disclosed in Chinese invention patent application publication number CN118466300A still has the following problems:

[0004] (1) The scheme uses a single-redundant bus. If a single bus fails, it may affect the function of some devices, causing the redundancy of the flight control computer to switch, thus affecting data transmission.

[0005] (2) If there are many mechanisms or devices connected on a single bus, the response time will be longer, which is not conducive to the control and monitoring of the system.

[0006] (3) When the rise / fall time of the communication signal is very short and the bus length is long (e.g., more than 1 / 10 of the signal wavelength), the communication bus is no longer a simple wire, but needs to be regarded as a "transmission line". The signal is propagated in the form of electromagnetic waves in the "transmission line". When it is transmitted to the end of the bus, if the end of the bus is open (i.e., the impedance is infinite), the energy of the electromagnetic wave cannot be absorbed and will be reflected back as if it hits a wall. This reflected electromagnetic wave will be superimposed with the subsequently transmitted signal, resulting in severe distortion of the electromagnetic waveform. This will cause the receiver to decode incorrectly and misjudge it as "0" or "1", thereby causing CRC error, communication interruption and other faults, which is especially serious in harsh electromagnetic environments.

[0007] Therefore, it is necessary to provide a tilt-type EVTOL flight control system to solve the above-mentioned technical problems. Utility Model Content

[0008] The main purpose of this utility model is to provide a tilt-type EVTOL flight control system, which, through a non-similar dual-redundant bus design, precise redundancy management, and real-time communication optimization, constructs a highly safe, reliable, real-time transmission, and easy-to-maintain flight control system.

[0009] This utility model achieves the above objectives through the following technical solution: a tilt-type EVTOL flight control system, comprising:

[0010] A redundant flight control computer system, which includes several flight control computers;

[0011] The power unit includes several tilting power units and several take-off and landing power units;

[0012] The wing surface control system has several components.

[0013] The system includes a control device, a data link device, and a control device, all of which are connected to the flight control computer.

[0014] Several main communication buses and several slave communication buses, each of the main communication buses or each of the slave communication buses is connected to all of the flight control computers, and the tilt propulsion device, the take-off and landing propulsion device and the wing surface control device are all connected to at least one of the main communication buses and at least one of the slave communication buses.

[0015] Matching resistors are provided inside the redundant flight control computer, the power unit, and the wing surface control unit.

[0016] Furthermore, the plurality of flight control computers include three independent and redundant first flight control computers and one backup flight control computer, with the three independent first flight control computers forming a master / backup / backup mechanism.

[0017] Furthermore, each of the tilt propulsion devices includes a variable pitch module, a tilt electric drive module, and a tilt rudder control module, and each of the take-off and landing propulsion devices includes a take-off and landing electric drive module and a fixed-pitch propeller.

[0018] Furthermore, the variable pitch module, the tilt electric drive module, the tilt rudder control module, the take-off and landing electric drive module, and the fixed pitch propeller are each provided with at least two bus interfaces, one of which is connected to the main communication bus and the other is connected to the slave communication bus.

[0019] Furthermore, the data link device is equipped with electronic circuitry and a radio frequency module, and both the electronic circuitry and the radio frequency module are provided in duplicate, meaning that the data link device is a dual-redundancy mechanism.

[0020] Furthermore, the control device includes an atmospheric data module, an inertial navigation module, a global navigation satellite module, a radar altimeter, a display and control module, an energy module, a power distribution module, and an environmental control module.

[0021] Furthermore, there are six tilting propulsion devices, six takeoff and landing propulsion devices, and five wing surface control devices.

[0022] Furthermore, the main communication bus has six lines, and the slave communication bus has six lines.

[0023] Furthermore, a first tilting power unit and a second tilting power unit are connected to the first main communication bus;

[0024] The second main communication bus is connected to the third tilting power unit and the fourth tilting power unit.

[0025] The third main communication bus connects the fifth tilting power unit and the sixth tilting power unit;

[0026] The fourth main communication bus is connected to the first takeoff and landing power unit, the second takeoff and landing power unit, the first wing surface control unit, and the second wing surface control unit.

[0027] The fifth main communication bus is connected to the third takeoff and landing power unit, the fourth takeoff and landing power unit, and the third wing surface control unit;

[0028] The sixth main communication bus is connected to the fifth takeoff and landing power unit, the sixth takeoff and landing power unit, the fourth wing surface control unit, and the fifth wing surface control unit.

[0029] The first is connected to the fourth tilt propulsion unit, the fourth takeoff and landing propulsion unit and the fifth wing surface control unit via the communication bus.

[0030] The second is connected to the communication bus to the second tilt propulsion unit, the second takeoff and landing propulsion unit and the fourth wing surface control unit.

[0031] The third is connected to the fifth tilt propulsion unit, the fifth takeoff and landing propulsion unit and the first wing surface control unit via the communication bus;

[0032] The fourth is connected to the third tilt propulsion unit, the third takeoff and landing propulsion unit, and the second wing surface control unit via the communication bus;

[0033] The fifth device is connected to the first tilt propulsion unit, the first takeoff and landing propulsion unit, and the third wing surface control unit via the communication bus.

[0034] The sixth tilting power unit and the sixth take-off and landing power unit are connected to the communication bus.

[0035] Furthermore, the flight control computer is equipped with at least twelve bus interfaces.

[0036] Compared with the prior art, the beneficial effects of the tilt-type EVTOL flight control system of this utility model are as follows:

[0037] 1. Extremely high system safety and reliability: The system has a multi-level redundancy architecture and a comprehensive redundancy design. Its core includes: (1) Flight control computer redundancy: The system adopts a hybrid redundancy mode of "three main backups (main / backup / backup) + one independent backup". The three main flight control computers achieve hot backup through redundancy management to ensure that the control function is not lost under single or even double point failure. The independent backup flight control computer effectively prevents possible common mode failures of the main system and meets the stringent safety requirements of the EVTOL aircraft; (2) Redundancy of key actuators: The core components such as motors and controllers inside the tilt propulsion device and wing surface control device adopt redundant designs such as dual windings and dual electronic control, which significantly improves the reliability of the execution level; (3) Dual redundancy data link: Ensures the data link between the ground station and the aircraft. The smooth operation of the remote control and telemetry channel ensures communication even in the event of interference in a single frequency band or equipment failure; (4) Non-similar dual-redundant bus network: The non-similar dual-redundant communication network is constructed using two buses with different characteristics: the master communication bus (CAN bus) and the slave communication bus (RS485 bus); (5) Anti-common-mode failure: Due to the different physical layers and protocol layers of the two buses, the risk of common-mode failure due to design defects, electromagnetic interference, etc., is extremely high for one bus, while the other bus can still work normally, fundamentally avoiding the risk of system paralysis caused by single-type bus failure; (6) Fault isolation: The careful bus connection layout ensures that when any one or two buses fail, at most only one power unit or control surface unit fails, minimizing the impact of local faults and ensuring the controllability of the aircraft;

[0038] 2. Excellent real-time and deterministic communication: (1) Optimized bus load distribution, which evenly distributes the seventeen key execution devices (six tilt power units + six take-off and landing power units + five wing surface control units) to six main communication buses and six slave communication buses, ensuring that the number of devices connected to each communication bus is reasonable (maximum six); (2) Guaranteeing RS485 bus response time: Through the "call-by-name" communication mechanism, the flight control computer can complete a complete interaction with all slave devices on the RS485 bus within a maximum of 60ms. This deterministic delay design meets the strict real-time requirements of the flight control system and ensures that control commands can be responded to in a timely manner;

[0039] 3. Precise fault management capabilities: (1) A sound redundancy management strategy: the three main flight control computers dynamically elect the host through synchronous calculation and voting mechanism, realizing seamless hot backup switching and ensuring a smooth transition of control; (2) Comprehensive status monitoring and feedback: all key actuators (such as tilt control module and wing surface control device) can provide real-time feedback of their status information and sensor data (such as control surface angle and tilt position) to the flight control computer, providing a data foundation for the flight control computer to perform system health assessment, fault diagnosis and reconfiguration control;

[0040] 4. Flexible configuration and maintainability: (1) Modular and standardized design: The entire system is connected by clearly defined modules (flight control computer, power unit, rudder control unit, etc.) through the main communication bus and the slave communication bus, which facilitates the design, testing, integration and maintenance of the system; (2) Flexible ID identification mechanism: The device ID is defined by the discrete signal of "ground / on" on the plug. No complicated software configuration is required. The device can be flexibly assigned a unique identity through physical jumpers, which simplifies the equipment replacement process in production assembly and later maintenance.

[0041] 5. Performance optimization: (1) High-precision servo control: The wing surface servo adopts the scheme of "DC brushless motor + reduction gear + harmonic reducer". The final output has the characteristics of low speed, high torque and high precision (low backlash), which is crucial for achieving precise attitude control of the aircraft; (2) Signal integrity guarantee: The redundant flight control computer device, power device and wing surface servo control device are all equipped with matching resistors. When it is necessary to eliminate the reflection of the signal at the end of the communication bus, the matching resistor in the device located at the end of the communication bus is connected to the main communication bus and the slave communication bus by shorting the plug pin. When the signal is transmitted to the end of the communication bus, its energy is completely absorbed by the matching resistor and converted into heat energy, which can eliminate the reflection of the signal at the end of the communication bus, effectively absorb the signal energy, eliminate the signal reflection, and ensure the quality and stability of long-distance, multi-node bus communication;

[0042] Therefore, this tilt-type EVTOL flight control system, through its multi-level redundant architecture, dissimilar bus design, precise redundancy management, and real-time communication optimization, constructs a highly secure, reliable, real-time-transmitting, and easy-to-maintain flight control system. Attached Figure Description

[0043] Figure 1 This is a schematic diagram of the tilt-type EVTOL flight control system according to an embodiment of the present invention;

[0044] Figure 2 This is a schematic diagram showing the installation positions of the matching resistor on the main communication bus and the slave communication bus according to an embodiment of the present invention.

[0045] The numbers in the image represent:

[0046] 100-tilt-type EVTOL flight control system;

[0047] 1-Redundant flight control computer unit; 11-Flight control computer;

[0048] 21-Tilting power unit, 21a-First tilting power unit, 21b-Second tilting power unit, 21c-Third tilting power unit, 21d-Fourth tilting power unit, 21e-Fifth tilting power unit, 21f-Sixth tilting power unit, 211-Variable pitch module, 212-Tilting electric drive module, 213-Tilting rudder control module, 22-Landing and takeoff power unit, 22a-First landing and takeoff power unit, 22b-Second landing and takeoff power unit, 22c-Third landing and takeoff power unit, 22d-Fourth landing and takeoff power unit, 22e-Fifth landing and takeoff power unit, 22f-Sixth landing and takeoff power unit, 221-Landing and takeoff electric drive module;

[0049] 3-wing surface control device, 3a-first wing surface control device, 3b-second wing surface control device, 3c-third wing surface control device, 3d-fourth wing surface control device, 3e-fifth wing surface control device;

[0050] 4-Operating device; 5-Data link device; 6-Control device;

[0051] 7-Main communication bus, 7a-First main communication bus, 7b-Second main communication bus, 7c-Third main communication bus, 7d-Fourth main communication bus, 7e-Fifth main communication bus, 7f-Sixth main communication bus;

[0052] 8 - Slave communication bus, 8a - First slave communication bus, 8b - Second slave communication bus, 8c - Third slave communication bus, 8d - Fourth slave communication bus, 8e - Fifth slave communication bus, 8f - Sixth slave communication bus;

[0053] 9-Matching resistor. Detailed Implementation

[0054] Please refer to Figures 1-2 This embodiment is a tilt-type EVTOL flight control system 100, which includes:

[0055] A redundant flight control computer device 1, which includes several flight control computers 11;

[0056] The power unit includes several tilting power units 21 and several take-off and landing power units 22;

[0057] The wing surface control device 3 has several components;

[0058] The system includes a control device 4, a data link device 5, and a control device 6, all of which are connected to the flight control computer 11.

[0059] Several main communication buses 7 and several slave communication buses 8, each main communication bus 7 or each slave communication bus 8 is connected to all flight control computers 11, and the tilt propulsion unit 21, the takeoff and landing propulsion unit 22 and the wing surface control unit 3 are all connected to at least one main communication bus 7 and at least one slave communication bus 8.

[0060] Matching resistor 9 is pre-installed inside the redundant flight control computer unit 1, the power unit 2, and the wing surface control unit 3. When it is necessary to eliminate signal reflection at the end of the communication bus, the matching resistor 9 inside the device located at the end of the communication bus is connected to the main communication bus 7 and the slave communication bus 8 by shorting the connector pins. Figure 2 As shown, when a signal is transmitted to the end of the communication bus, its energy is completely absorbed by the matching resistor and converted into heat energy, which can eliminate signal reflection at the end of the communication bus and ensure communication quality. The resistance value of the matching resistor 9 can be set according to the actual situation and is not limited here. Preferably, its resistance value is 120Ω or 100Ω.

[0061] In this embodiment, the flight control computers 11 include three independent and redundant primary flight control computers and one backup flight control computer. The three independent primary flight control computers operate on a master / backup / backup mechanism, achieving redundant control through connections to the main communication bus 7 and the slave communication bus 8, ensuring the availability of critical controls. The three independent primary flight control computers employ redundancy management, synchronously calculating and determining one primary flight control computer as the master during each run, with the other two serving as slave flight control computers for hot backup. The master flight control computer is responsible for sending commands to external devices or receiving data from external devices, while the slave flight control computers are responsible for receiving data from external devices. The backup flight control computer prevents common-mode issues among the three primary flight control computers. When all three primary flight control computers fail, the backup flight control computer provides an independent backup control system, ensuring the safety of the aircraft control meets the requirements of the EVTOL aircraft.

[0062] In other embodiments, a plurality of flight control computers 11 can satisfy redundant control. The number of flight control computers 11 is not limited here and can be set according to actual conditions.

[0063] Each tilt propulsion unit 21 includes a pitch control module 211, a tilt electric drive module 212, and a tilt rudder control module 213. Each take-off and landing propulsion unit 22 includes a take-off and landing electric drive module 221 and a fixed-pitch propeller. Each of the pitch control module 211, tilt electric drive module 212, tilt rudder control module 213, take-off and landing electric drive module, and fixed-pitch propeller is provided with at least two bus interfaces, one of which is connected to the main communication bus 7 and the other is connected to the slave communication bus 8.

[0064] The pitch module 211 includes a pitch actuator and a pitch propeller. The pitch actuator executes the control commands of the flight control computer 11 to change the pitch of the pitch propeller and provide different thrust or lift to the aircraft.

[0065] Both the tilt electric drive module 212 and the takeoff and landing electric drive module include a motor, a motor controller, and an external wind deflector. The motor and motor controller are integrated into a single structure, used to execute control commands from the flight control computer 11, rotating and driving the propeller to perform work, providing power to the aircraft. The motor has a dual-winding structure, and the motor controller has a dual-control mechanism. The motor controller is simultaneously connected to at least one main communication bus 7 and at least one slave communication bus 8 to achieve redundancy.

[0066] The tilt control module 213 includes a tilt servo unit and a tilt angle sensor. The tilt control module 213 receives and executes tilt control commands from the flight control computer 11 and sends its own status information to the flight control computer 11. The tilt angle sensor monitors the rotational position of the tilt rotor in real time and sends the tilt rotor's position information to the flight control computer 11 in real time through the tilt servo unit. In multi-rotor mode, the fixed-pitch propeller provides the lift required by the aircraft.

[0067] Each wing surface control device 3 includes one wing surface servo controller and two wing surface servos. The wing surface servo controller is equipped with at least two bus interfaces, one of which is connected to the main communication bus 7 and the other is connected to the slave communication bus 8, controlling the two wing surface servos respectively. The wing surface control device 3 is used to execute the wing surface control commands of the flight control computer 11, drive the corresponding wing surface servos to rotate, and send back the status of the wing surface control device 3 and the angle information of the wing surface servos to the flight control computer 11 in real time, so that the aircraft flies in a predetermined attitude.

[0068] The wing-surface servo controller includes a control circuit and a drive circuit. The control circuit uses a digital control scheme of "MCU (Microcontroller Unit) + logic chip," while the drive circuit uses a scheme of "optical isolation + MOSFET-based three-phase bridge drive." The control circuit and the drive circuit are completely electrically isolated. The control signals from the flight control computer 11 are transmitted to the MCU. The MCU collects sensor position information and outputs motor PWM control signals after PID calculation. The PWM control signals are sent to the logic chip for logic calculation. The logic chip performs processing functions such as three-phase bridge control signal output, Hall sensor acquisition, and Hall fault signal feedback. The drive circuit drives the corresponding motor windings after passing through optical isolation and drive / inverter circuits.

[0069] The wing surface servo consists of a brushless DC motor, a reduction gear, and a harmonic reducer. The brushless DC motor is the power source, the reduction gear performs primary speed reduction, responsible for initially reducing speed and increasing torque. The harmonic reducer performs the final speed reduction and output, converting the power source into the low-speed, high-torque, and high-precision motion required by the control surface.

[0070] The control device 4 is used by the pilot to control the aircraft when it is manned. It includes, but is not limited to, the throttle lever, control stick, and pedals, and may also include other structures. The control device 4 can adopt the design of the existing technology, and there are no restrictions here. The control device 4 is connected to the flight control computer 11 via a serial port (e.g., RS-422). After the pilot operates the control device 4, the physical operation is converted into an electrical signal. The flight control computer 11 processes the signal and sends instructions to various devices to control the aircraft to perform corresponding actions.

[0071] Data link device 5 is used for wireless data transmission. It employs a dual-redundancy mechanism, meaning it contains two complete, independent, and functionally identical sets of electronic circuitry and radio frequency modules (including processors, modems, transmitters, and receivers). Each redundancy has two sets of electronic circuitry and radio frequency modules, with each redundancy using a different wireless transmission frequency (dual-band). Data link device 5 connects to flight control computer 11 via a serial port (e.g., RS-422), allowing it to receive video signals and also supporting the transmission of image data and remote control / telemetry data. In unmanned flight operations, data link device 5 allows the pilot to control and monitor the aircraft via a ground station. The specific methods for control and monitoring are existing technologies and will not be elaborated upon here.

[0072] The control device 6 is connected to the flight control computer 11 via a serial port (e.g., RS-422). The control device 6 includes an air data module, an inertial navigation module, a global navigation satellite module, a radar altimeter, a display and control module, an energy module, a power distribution module, and an environmental control module. The air data module measures atmospheric parameters such as airspeed and altitude for flight control calculations; the inertial navigation module autonomously and continuously calculates position and attitude without relying on external signals; the global navigation satellite module provides high-precision global position, velocity, and time information; the radar altimeter measures the actual altitude to the ground for low-altitude flight and takeoff and landing; the display and control module is a human-machine interface that displays information and receives commands; the energy module generates and stores electrical energy (power generation system); the power distribution module distributes, protects, and controls electrical energy transmission; and the environmental control module regulates the cockpit environment and cools the equipment (life support). The aforementioned air data module, inertial navigation module, global navigation satellite module, radar altimeter, display and control module, energy module, power distribution module, and environmental control module are all existing technologies and will not be described in detail here.

[0073] In this embodiment, six tilting power units 21, six take-off and landing power units 22, and five wing surface control units 3 are provided. Correspondingly, six main communication buses 7 and six slave communication buses 8 are provided. The main communication bus 7 is a CAN bus, and the slave communication bus 8 is an RS485 bus.

[0074] The CAN bus is used for communication and transmission of critical signals, while the RS485 bus can transmit critical signals between devices in the event of a CAN bus failure. This forms a non-similar dual-redundant bus communication logic to prevent aircraft loss of control caused by CAN bus failure.

[0075] Because the CAN bus has an arbitration mechanism, the communication logic of the CAN bus is as follows: The master flight control computer sends relevant control protocol packets at preset intervals. When a slave device (including the tilt propulsion unit 21, the takeoff and landing propulsion unit 22, and the wing surface control unit 3) receives a control protocol packet with its own target ID, it receives, parses, and executes the instructions in the control protocol packet. Furthermore, the slave device (including the tilt propulsion unit 21, the takeoff and landing propulsion unit 22, and the wing surface control unit 3) will send relevant operation data packets at fixed intervals according to its own data frame characteristics. When both the master and slave flight control computers receive an operation data packet with the target ID of either flight control computer, they both receive and parse the operation data packet. The preset interval can be 10ms, or other interval times, which can be set according to actual conditions and are not limited here; the fixed interval can also be set according to actual conditions and is not limited here.

[0076] In RS5485 bus communication, if a response is required from a slave device (including tilt propulsion unit 21, takeoff and landing propulsion unit 22, and wing surface control unit 3), it must be called by name by the flight control computer. The RS485 bus control protocol packet includes a source device ID and a destination device ID. The source device ID refers to the ID of the flight control computer, and the destination device ID refers to the ID of the slave device (including tilt propulsion unit 21, takeoff and landing propulsion unit 22, and wing surface control unit 3). The RS485 bus communication logic is as follows: The master flight control computer sends relevant control protocol packets periodically according to a preset cycle. If a slave device (including tilt propulsion unit 21, takeoff and landing propulsion unit 22, and wing surface control unit 3) identifies the source device ID in the control protocol packet as any flight control computer, it immediately executes the control commands in the control protocol packet. When the destination device ID is any slave device (including tilt propulsion unit 21, takeoff and landing propulsion unit 22, and wing surface control unit 3), the master flight control computer sends relevant control protocol packets periodically. 1. When the takeoff and landing power unit 22 and the wing surface control device 3 are received, the slave device (including the tilt power unit 21, the takeoff and landing power unit 22 and the wing surface control device 3) immediately sends a feedback data packet (response time < 5ms). Whether the slave device (including the tilt power unit 21, the takeoff and landing power unit 22 and the wing surface control device 3) sends the feedback data packet depends only on the destination device ID in the flight control computer control protocol packet; otherwise, it will not be sent. When the master flight control computer and the slave flight control computer receive the data packet sent by the slave device (including the tilt power unit 21, the takeoff and landing power unit 22 and the wing surface control device 3) with the target ID of any flight control computer, they will both receive and parse the feedback data packet.

[0077] RS485 bus communication process: The master flight control computer requests all slave devices to execute instructions and specifically requests slave device 1 to send a feedback data packet → All slave devices receive and execute the master flight control computer's instructions → Slave device 1 immediately sends a feedback data packet according to the master flight control computer's request → After a preset period (e.g., 10ms), the master flight control computer requests all slave devices to execute instructions and specifically requests slave device 2 to send a feedback data packet → All slave devices receive and execute the master flight control computer's instructions → Slave device 2 sends a feedback data packet according to the master flight control computer's request → This process repeats continuously, and the master flight control computer completes the interaction with up to 6 slave devices on the RS485 bus within a maximum of 60ms.

[0078] The aforementioned ID refers to the unique identification information of each device, unit, module, etc. This ID is defined by the state of several discrete "ground / on" input signals on the socket. By shorting these discrete "ground / on" input signals to GND (ground) or leaving them floating on the connector, the "0" and "1" states can be distinguished. For example, three of these discrete "ground / on" input signals can form a maximum of eight ID information, capable of identifying eight devices. Multiple devices, units, modules, etc., can have different discrete "ground / on" input signals set according to the actual number, giving each device, unit, module, etc., its own ID, facilitating identification or naming by other devices.

[0079] The six tilting power devices 21 are: first tilting power device 21a, second tilting power device 21b, third tilting power device 21c, fourth tilting power device 21d, fifth tilting power device 21e, and sixth tilting power device 21f.

[0080] The six take-off and landing power units 22 are: the first take-off and landing power unit 22a, the second take-off and landing power unit 22b, the third take-off and landing power unit 22c, the fourth take-off and landing power unit 22d, the fifth take-off and landing power unit 22e, and the sixth take-off and landing power unit 22f.

[0081] The five wing surface control devices 3 are: first wing surface control device 3a, second wing surface control device 3b, third wing surface control device 3c, fourth wing surface control device 3d, and fifth wing surface control device 3e.

[0082] The six main communication buses 7 are: the first main communication bus 7a, the second main communication bus 7b, the third main communication bus 7c, the fourth main communication bus 7d, the fifth main communication bus 7e, and the sixth main communication bus 7f.

[0083] The six slave communication buses 8 are: first slave communication bus 8a, second slave communication bus 8b, third slave communication bus 8c, fourth slave communication bus 8d, fifth slave communication bus 8e, and sixth slave communication bus 8f.

[0084] Correspondingly, each main communication bus 7 or each slave communication bus 8 is connected to all flight control computers 11. That is, each flight control computer 11 is connected to six main communication buses 7 and six slave communication buses 8, so each flight control computer 11 has at least twelve bus interfaces. The tilt propulsion unit 21, the takeoff and landing propulsion unit 22, and the wing surface control unit 3 are all connected to at least one main communication bus 7 and at least one slave communication bus 8, so each of the tilt propulsion unit 21, the takeoff and landing propulsion unit 22, and the wing surface control unit 3 has at least two bus interfaces.

[0085] In other embodiments, the number of bus interfaces on each flight control computer 11, including the tilt propulsion unit 21, takeoff and landing propulsion unit 22, wing surface control unit 3, main communication bus 7, slave communication bus 8, and tilt propulsion unit 21, is not limited and can be adjusted according to actual conditions. Moreover, the types of main communication bus 7 and slave communication bus 8 can also be adjusted according to actual conditions and are not limited to the CAN bus and RS485 bus mentioned above.

[0086] In this embodiment, the specific connection methods between the main communication bus 7, the slave communication bus 8, the tilt propulsion device 21, the takeoff and landing propulsion device 22, and the wing surface control device 3 are as follows:

[0087] The first main communication bus 7a is connected to the first tilting power device 21a and the second tilting power device 21b.

[0088] The third tilting power unit 21c and the fourth tilting power unit 21d are connected to the second main communication bus 7b;

[0089] The fifth tilting power unit 21e and the sixth tilting power unit 21f are connected to the third main communication bus 7c;

[0090] The first takeoff and landing power unit 22a, the second takeoff and landing power unit 22b, the first wing surface control device 3a, and the second wing surface control device 3b are connected to the fourth main communication bus 7d.

[0091] The third takeoff and landing power unit 22c, the fourth takeoff and landing power unit 22d, and the third wing surface control unit 3c are connected to the fifth main communication bus 7e.

[0092] The sixth main communication bus 7f is connected to the fifth takeoff and landing power unit 22e, the sixth takeoff and landing power unit 22f, the fourth wing surface control unit 3d, and the fifth wing surface control unit 3e;

[0093] The first communication bus 8a is connected to the fourth tilt propulsion unit 21d, the fourth takeoff and landing propulsion unit 22d and the fifth wing surface control unit 3e.

[0094] The second tilting power unit 21b, the second takeoff and landing power unit 22b and the fourth wing surface control unit 3d are connected to the second communication bus 8b.

[0095] The fifth tilt propulsion unit 21e, the fifth takeoff and landing propulsion unit 22e, and the first wing surface control unit 3a are connected to the communication bus 8c.

[0096] The fourth is connected to the third tilt propulsion unit 21c, the third takeoff and landing propulsion unit 22c and the second wing surface control unit 3b via the communication bus 8d;

[0097] The fifth device is connected to the first tilting power unit 21a, the first takeoff and landing power unit 22a and the third wing surface control unit 3c via the communication bus 8e.

[0098] The sixth tilting power unit 21f and the sixth take-off and landing power unit 22f are connected to the sixth communication bus 8f.

[0099] Specifically, when the main communication bus 7 and the slave communication bus 8 are connected to the tilt propulsion unit 21, they are all connected to the internal pitch control module 211, tilt electric drive module 212, and tilt rudder control module 213. That is, each of the pitch control module 211, tilt electric drive module 212, and tilt rudder control module 213 is provided with at least two bus interfaces, one of which is connected to the main communication bus 7 and the other is connected to the slave communication bus 8. When the main communication bus 7 and the slave communication bus 8 are connected to the take-off and landing propulsion unit 22a, they are connected to the internal take-off and landing electric drive module. The take-off and landing electric drive module is provided with at least two bus interfaces, one of which is connected to the main communication bus 7 and the other is connected to the slave communication bus 8.

[0100] Each flight control computer 11 is interconnected with six tilt propulsion units 21, six takeoff and landing propulsion units 22, and five wing surface control units 3 via six CAN buses and six RS485 buses, thus forming a dissimilar dual-redundant bus communication architecture, effectively reducing the risk of common-mode failure caused by bus failure. The advantages of the dissimilar dual-redundant bus communication interconnection method are mainly twofold: (1) If any bus fails, it will not affect the function of any device; if any two buses fail, at most one propulsion unit or one wing surface control unit 3 will fail; (2) The load is basically evenly distributed, so that the load on each bus is not too large, and the number of slave devices (propulsion units and wing surface control units) on each bus does not exceed 6, so that the flight control computer 11 can complete the interaction with the slave devices on the RS485 bus within 6 cycles, ensuring the real-time performance of the system.

[0101] In other embodiments, the connection methods between the main communication bus 7, the slave communication bus 8 and the tilt propulsion device 21, the take-off and landing propulsion device 22, and the wing surface control device 3 can be adjusted according to the actual situation, as long as the tilt propulsion device 21, the take-off and landing propulsion device 22 and the wing surface control device 3 are all connected to at least one main communication bus 7 and at least one slave communication bus 8.

[0102] The above descriptions are merely some embodiments of this utility model. For those skilled in the art, various modifications and improvements can be made without departing from the inventive concept of this utility model, and all such modifications and improvements fall within the protection scope of this utility model.

Claims

1. A tilt-type EVTOL flight control system, characterized in that, It includes: A redundant flight control computer system, which includes several flight control computers; The power unit includes several tilting power units and several take-off and landing power units; The wing surface control system has several components. The system includes a control device, a data link device, and a control device, all of which are connected to the flight control computer. Several main communication buses and several slave communication buses, each of the main communication buses or each of the slave communication buses is connected to all of the flight control computers, and the tilt propulsion device, the take-off and landing propulsion device and the wing surface control device are all connected to at least one of the main communication buses and at least one of the slave communication buses. Matching resistors are provided inside the redundant flight control computer, the power unit, and the wing surface control unit.

2. The tilt-type EVTOL flight control system as described in claim 1, characterized in that: The flight control computers include three independent and redundant first flight control computers and one backup flight control computer, with the three independent first flight control computers forming a master / backup / backup mechanism.

3. The tilt-type EVTOL flight control system as described in claim 1, characterized in that: Each of the tilt propulsion units includes a pitch control module, a tilt electric drive module, and a tilt rudder control module; each of the takeoff and landing propulsion units includes a takeoff and landing electric drive module and a fixed-pitch propeller.

4. The tilt-type EVTOL flight control system as described in claim 3, characterized in that: The variable pitch module, the tilt electric drive module, the tilt rudder control module, and the take-off and landing electric drive module, as well as the fixed pitch propeller, are each provided with at least two bus interfaces, one of which is connected to the main communication bus and the other is connected to the slave communication bus.

5. The tilt-type EVTOL flight control system as described in claim 1, characterized in that: The data link device is equipped with electronic circuits and radio frequency modules, and there are two sets of both the electronic circuits and the radio frequency modules, that is, the data link device is a dual-redundancy mechanism.

6. The tilt-type EVTOL flight control system as described in claim 1, characterized in that: The control device includes an atmospheric data module, an inertial navigation module, a global navigation satellite module, a radar altimeter, a display and control module, an energy module, a power distribution module, and an environmental control module.

7. The tilt-type EVTOL flight control system as described in claim 1, characterized in that: The tilting propulsion system is provided in six units, the takeoff and landing propulsion system is provided in six units, and the wing surface control system is provided in five units.

8. The tilt-type EVTOL flight control system as described in claim 7, characterized in that: The main communication bus has six lines, and the slave communication bus has six lines.

9. The tilt-type EVTOL flight control system as described in claim 8, characterized in that: The first main communication bus is connected to the first tilting power unit and the second tilting power unit. The second main communication bus is connected to the third tilting power unit and the fourth tilting power unit. The fifth tilting power unit and the sixth tilting power unit are connected to the third main communication bus; The fourth main communication bus is connected to the first takeoff and landing power unit, the second takeoff and landing power unit, the first wing surface control unit, and the second wing surface control unit. The fifth main communication bus is connected to the third takeoff and landing power unit, the fourth takeoff and landing power unit, and the third wing surface control unit; The sixth main communication bus is connected to the fifth takeoff and landing power unit, the sixth takeoff and landing power unit, the fourth wing surface control unit, and the fifth wing surface control unit. The first is connected to the fourth tilt propulsion unit, the fourth takeoff and landing propulsion unit and the fifth wing surface control unit via the communication bus. The second is connected to the communication bus to the second tilt propulsion unit, the second takeoff and landing propulsion unit and the fourth wing surface control unit. The third is connected to the fifth tilt propulsion unit, the fifth takeoff and landing propulsion unit and the first wing surface control unit via the communication bus; The fourth is connected to the third tilt propulsion unit, the third takeoff and landing propulsion unit, and the second wing surface control unit via the communication bus; The fifth device is connected to the first tilt propulsion unit, the first takeoff and landing propulsion unit, and the third wing surface control unit via the communication bus. The sixth tilting power unit and the sixth take-off and landing power unit are connected to the communication bus.

10. The tilt-type EVTOL flight control system as described in claim 8, characterized in that: The flight control computer is equipped with at least twelve bus interfaces.

Images