Flight control system architecture of eVTOL aircraft

Abstract

The invention discloses a flight control system architecture of an eVTOL aircraft. The flight control system architecture comprises a navigation system, a sensor system, a servo system and a flight control computer, the navigation system comprises a main navigation subsystem, a second-level navigation subsystem and a third-level navigation subsystem. The sensor system comprises a main data acquisition system, a standby data acquisition system and a height sensor; the servo system comprises a dual-redundancy servo controller, a dual-redundancy steering engine, a dual-redundancy power supply system, a tilting servo driving system and a control switching system; the flight control computer comprises a control law switching system and a control distribution system; and the navigation system, the sensor system and the servo system are all connected with the flight control computer. The interaction logic between the systems is reasonable, the adaptability of the aircraft in a complex environment can be improved on the basis of ensuring the flight safety and reliability, and the design and use cost can be effectively reduced.

Description

Technical Field

[0001] The present invention relates to the field of aircraft technology, and in particular to a flight control system architecture for an eVTOL aircraft. Background Art

[0002] Currently, in the design of eVTOL (electric Vertical Take-off and Landing) aircraft with a tilt-rotating configuration that combines fixed wings and multi-rotors, the flight control system mostly adopts a distributed architecture. However, the design of this distributed architecture is complex, and the design and use costs are too high. Summary of the Invention

[0003] In order to solve the above-mentioned problem of oil-powered multi-rotor aircraft, the present invention provides a flight control system architecture for eVTOL aircraft, which can serve as a key system of tilt-rotating aircraft. The interaction logic between systems is reasonable, which can improve the adaptability of the aircraft in complex environments on the basis of ensuring flight safety and reliability, and can effectively reduce design and use costs.

[0004] The present invention provides the following technical solutions:

[0005] Provided is a flight control system architecture for an eVTOL aircraft, which includes: a navigation system, a sensor system, a servo system, and a flight control computer;

[0006] The navigation system includes a primary navigation subsystem, a secondary navigation subsystem, and a tertiary navigation subsystem, and all three navigation subsystems are capable of obtaining aircraft status and / or ground speed;

[0007] The sensor system includes a primary data acquisition system, a backup data acquisition system, and an altitude sensor; the primary atmospheric data acquisition system and the backup atmospheric data acquisition system are both used to collect atmospheric data of the atmospheric environment in which the aircraft is located and / or flight data of the aircraft, and the altitude sensor is used to obtain the aircraft's altitude above the ground;

[0008] The servo system includes a dual-redundancy servo controller, a dual-redundancy steering gear, a dual-redundancy power supply system, a tilt servo drive system and a control switching system;

[0009] The flight control computer includes a control law switching system and a control distribution system;

[0010] Furthermore, the navigation system, sensor system, and servo system are all connected to the flight control computer.

[0011] Compared with the prior art, the present invention has the following beneficial effects:

[0012] The flight control system architecture logic design of this application is reasonable, and it can serve as a key system of a tilt-rotating configuration aircraft. Through the reasonable interaction logic between the navigation system, sensor system, servo system and flight control computer, as well as the design and switching of control laws at different levels, it can improve the adaptability of the aircraft in complex environments while ensuring flight safety and reliability, and can effectively reduce design and use costs. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] Figure 1 Schematic diagram of the structure of the flight control system architecture of the eVTOL aircraft in the present invention;

[0014] Figure 2 Schematic diagram of the design architecture of the control law in the present invention;

[0015] Figure 3 A schematic diagram of the connection relationship between the flight control computer and other systems in the present invention;

[0016] Figure 4 Schematic diagram of switching between different control laws in the present invention. DETAILED DESCRIPTION

[0017] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0018] Example 1:

[0019] This embodiment provides a flight control system architecture for an eVTOL aircraft, wherein the eVTOL aircraft may adopt a tilt-rotor configuration, which includes a fixed rotor and a tilt-rotor. At the same time, the eVTOL aircraft may be a manned aircraft or an unmanned aircraft. On this basis, Figure 1 As shown, the flight control system architecture includes: a navigation system 1, a sensor system 2, a servo system 3 and a flight control computer 4;

[0020] Wherein, the navigation system 1 includes a main navigation subsystem 11, a secondary navigation subsystem 12 and a tertiary navigation subsystem 13, and all three navigation subsystems can obtain aircraft state quantities and / or ground speeds, and the aircraft state quantities include one or more of three-axis angular rate, three-axis acceleration, roll angle, pitch angle, heading angle, horizontal speed, vertical speed, longitude and latitude, and altitude (including air altitude and landing altitude). In addition, the main navigation subsystem 11 includes a fiber-optic inertial integrated navigation system, and one or two of the secondary navigation subsystem 12 and the tertiary navigation subsystem 13 are MEMS-based inertial integrated navigation systems, and one or more of the main navigation subsystem 11, the secondary navigation subsystem 12 and the tertiary navigation subsystem 13 include dual-antenna (i.e., main antenna and backup antenna) GNSS modules and support differential positioning;

[0021] Furthermore, the main navigation subsystem 11, the secondary navigation subsystem 12, and the tertiary navigation subsystem 13 are correspondingly level-set according to the first priority, the second priority, and the third priority from high to low (i.e., the navigation subsystem 11 is the first priority, the secondary navigation subsystem 12 is the second priority, and the tertiary navigation subsystem 13 is the third priority). By default, the main navigation subsystem 11 of the first priority is called first to obtain the aircraft status and / or ground speed. When the navigation subsystem of the previous priority fails, the navigation subsystems of the next priority are called in sequence to obtain the aircraft status and / or ground speed. For example, if the main navigation subsystem 11 fails, the secondary navigation subsystem 12 is called. If the secondary navigation subsystem 12 fails, the tertiary navigation subsystem 13 is called. Therefore, the navigation system 1 in this embodiment adopts a triple-redundancy design. Under normal conditions, the inertial integrated navigation system can accurately measure the attitude, altitude, status, and other information of the aircraft to complete the navigation task. When a navigation system fails, the remaining navigation subsystems can also be called to perform the task, thereby meeting the reliability and fault tolerance requirements of the system.

[0022] The sensor system 2 includes a primary data acquisition system 21, a backup data acquisition system 22, and an altitude sensor 23. The primary atmospheric data acquisition system 21 and the backup atmospheric data acquisition system 22 are both used to collect atmospheric data of the atmospheric environment in which the aircraft is located and / or flight data of the aircraft. The atmospheric data include one or more of static pressure and atmospheric temperature, and the flight data include one or more of angle of attack, sideslip angle, airspeed, and vertical speed. Specifically, the primary data acquisition system 21 includes one or more of a data processing unit, a static pressure sensor, an atmospheric temperature sensor, an angle of attack sensor, and a sideslip angle sensor. The backup data acquisition system 22 includes one or more of a data processing unit, a static pressure sensor, and an atmospheric temperature sensor. The altitude sensor 23 is used to obtain the aircraft's altitude above the ground (the altitude above the ground includes the airborne altitude and the landing altitude), and includes one or more of a radio altimeter and a millimeter-wave radar.

[0023] Furthermore, the primary data acquisition system 21, the primary navigation subsystem 11, and the secondary navigation subsystem 12 are level-set according to the first priority, the second priority, and the third priority from high to low (i.e., the primary data acquisition system 21 is the first priority, the navigation subsystem 11 is the second priority, and the secondary navigation subsystem 12 is the third priority). By default, the primary data acquisition system 21 of the first priority is first called to obtain the altitude of the aircraft. When a system of the previous priority fails, the systems of the next priority are called in sequence to obtain the altitude of the aircraft. For example, if the primary data acquisition system 21 fails, the primary navigation subsystem 11 is called. If the primary navigation subsystem 11 fails, the secondary navigation subsystem 12 is called.

[0024] The main navigation subsystem 11, radio altimeter and millimeter-wave radar are level-set according to the first priority, second priority and third priority from high to low (i.e., the main navigation subsystem 11 is the first priority, the radio altimeter is the second priority and the millimeter-wave radar is the third priority). By default, the main navigation subsystem 11 of the first priority is first called to obtain the landing altitude of the aircraft. When the system of the previous priority fails, the system of the next priority is called in sequence to obtain the landing altitude of the aircraft. For example, if the main data acquisition system 21 fails, the radio altimeter is called. If the radio altimeter fails, the millimeter-wave radar is called.

[0025] Furthermore, the primary data acquisition system 21 and the backup data acquisition system 22 are correspondingly level-set according to the first priority and the second priority from high to low (i.e., the primary data acquisition system 21 is the first priority and the backup data acquisition system 22 is the second priority). By default, the primary data acquisition system 21 of the first priority is first called to obtain the airspeed. When the system of the previous priority fails, the system of the next priority is called to obtain the airspeed. For example, if the primary data acquisition system 21 fails, the backup data acquisition system 22 is called;

[0026] Therefore, in this embodiment, in addition to angle of attack and sideslip angle, other flight data and aircraft status can be obtained from two or more sensors. When a system with the highest priority fails, the system with the next highest priority is called in sequence to continue data collection and processing, providing a hardware foundation for sensor redundancy management.

[0027] The servo system 3 includes a dual-redundancy servo controller 31, a dual-redundancy steering gear 32, a dual-redundancy power supply system 33, a tilt servo drive system 34 and a control switching system 35;

[0028] The aircraft in this embodiment has two ailerons, two elevators, and two rudders. Each rudder surface is equipped with a dual-redundant servo 32, and the dual-redundant servo 32 is independently driven by two drive channels.

[0029] The dual-redundant servo controller 31 has two independent control units, each of which includes a control circuit and a control program, and can receive command signals from the flight control computer 4 or other control terminal, and perform signal processing such as resolution and amplification before outputting them to the dual-redundant servo 32, so as to control a corresponding driving channel of the dual-redundant servo 32 to perform a corresponding action, thereby driving the dual-redundant servo 32 to move. At the same time, the dual-redundant servo 32 has a position feedback function for feeding back the actual position of the control surface to the dual-redundant servo controller 31;

[0030] The dual-redundancy power supply system 33 is connected to the dual-redundancy servo controller 31 and / or the dual-redundancy steering gear 32, and is used to provide independent dual-circuit power supplies for the dual-redundancy servo controller 31 and / or the dual-redundancy steering gear 32, ensuring that the system can still operate normally in the event of a single-circuit power failure;

[0031] The control switching system 35 is connected to the dual-redundant servo controller 31 and is used to detect the working status of the dual-redundant servo controller 31 in real time, and when one control unit of the dual-redundant servo controller 31 fails, automatically switch to the other control unit to operate, generate and send fault alarm information, and / or is connected to the dual-redundant servo 32 and is used to detect the working status of the dual-redundant servo 32 (including servo current, control surface position, etc.) in real time, and when one drive channel of the dual-redundant servo 32 fails, automatically switch to the other drive channel to operate, generate and send fault alarm information, and / or is connected to the dual-redundant power supply system 33 and is used to detect the working status of the dual-redundant power supply system 33 in real time (including power supply voltage, etc.), and when one power supply of the dual-redundant servo 32 fails, automatically switch to the other power supply to operate, generate and send fault alarm information;

[0032] The tilt servo drive system 34 is used to drive the tilt rotor to perform a tilting action, so that the aircraft can switch between a multi-rotor mode and a fixed-wing mode. It specifically includes a tilt servo (there can be several, such as four) and a servo controller, wherein the tilt servo is connected to the tilt rotor and includes a brushless DC motor, a reduction mechanism, a position sensor and other structures, and is used to drive the tilt rotor to perform a tilting action; the servo controller is connected to the tilt servo and is used to control the tilt servo to perform a corresponding action. Preferably, the servo controller can be a dual-redundant linear controller, and transmits signals to the host computer via an RS-422 bus;

[0033] Therefore, the servo system 3 in this embodiment adopts dual-channel independent control, has fault isolation and automatic switching functions, and ensures that effective control of the control surface can be maintained in the event of a single-channel failure, which can meet the requirements of high-reliability control of the control surface and improve flight safety.

[0034] The flight control computer 4 includes a CPU board 41, a power board 42, an interface board 43, a control law switching system 44, and a control distribution system 45. The CPU board 41 and / or the power board 42 adopt a triple-redundancy design. The power board 42 is connected to the CPU board 41 to supply power to the CPU board 41. The interface board 43 is used to connect the CPU board 41 and other external devices.

[0035] Furthermore, the flight control computer 4 uses a 4-redundant CAN bus as a bus to achieve full connection of data paths between various functional modules, support module-level reconstruction, facilitate system function expansion, and improve system availability; and at least two of the navigation system 1, sensor system 2, servo system 3 and flight control computer 4 are connected through a cross-channel data link (CCDL) to achieve resource sharing between redundancies; at the same time, one or more of the CPU board 41, power board 42 and interface board 43 are provided with a BIT circuit to improve the system self-detection coverage and enhance the system working reliability; at the same time, one or more of the CPU board 41, power board 42 and interface board 43 are provided with an overcurrent protection circuit to improve the reliability and safety of the system and suppress the spread of system faults;

[0036] At the same time, the control law switching system 44 is connected to one or more of the navigation system 1, sensor system 2, and servo system 3, and is used to complete the switching of the control law according to the working state changes of one or more of the navigation system 1, sensor system 2, and servo system 3, and generate corresponding control law instructions, and control one or more of the navigation system 1, sensor system 2, and servo system 3 to perform corresponding actions according to the control law instructions;

[0037] In this embodiment, the control law includes a primary mode control law, a secondary mode control law, and a backup mode control law;

[0038] Wherein, the main mode control law includes:

[0039] When the aircraft is in multi-rotor mode, in the longitudinal channel, the forward speed of the aircraft is controlled, and when the forward speed command is 0, the aircraft is controlled to maintain its current position; in the transverse channel, the lateral speed of the aircraft is controlled, and when the lateral speed command is 0, the aircraft is controlled to maintain its current position; in the vertical channel, the vertical speed of the aircraft is controlled, and when the vertical speed command is 0, the aircraft is controlled to maintain its current altitude; and in the heading channel, the yaw rate of the aircraft is controlled, and when the yaw angle command is 0, the aircraft is controlled to maintain its current heading angle;

[0040] When the aircraft is in tilt transition mode (i.e., the process of switching from multi-rotor mode to fixed-wing mode), the forward speed of the aircraft is controlled on the longitudinal channel; the roll angle of the aircraft is controlled on the lateral channel; the vertical speed of the aircraft is controlled on the vertical channel, and when the vertical speed command is 0, the aircraft is controlled to maintain the current altitude unchanged; and the yaw rate of the aircraft is controlled on the heading channel, and when the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle unchanged;

[0041] When the aircraft is in fixed-wing mode, the longitudinal channel controls the aircraft's airspeed; the transverse channel controls the aircraft's lateral speed, and when the lateral speed command is 0, controls the aircraft to maintain its current position; the vertical channel controls the aircraft's vertical speed, and when the vertical speed command is 0, controls the aircraft to maintain its current altitude; and the heading channel controls the aircraft to perform coordinated turns.

[0042] The sub-mode control law includes:

[0043] When the aircraft is in multi-rotor mode, the longitudinal channel controls the pitch angle of the aircraft; the lateral channel controls the roll angle of the aircraft; the vertical channel controls the vertical speed of the aircraft, and when the vertical speed command is 0, the aircraft is controlled to maintain the current altitude; and the heading channel controls the yaw rate of the aircraft, and when the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle.

[0044] When the aircraft is in the bank transition mode, the forward speed of the aircraft is controlled on the longitudinal channel; the roll angle of the aircraft is controlled on the lateral channel; the vertical speed of the aircraft is controlled on the vertical channel, and when the vertical speed command is 0, the aircraft is controlled to maintain the current altitude; and the yaw rate of the aircraft is controlled on the heading channel, and when the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle.

[0045] When the aircraft is in fixed-wing mode, the longitudinal channel controls the aircraft's airspeed; the lateral channel controls the aircraft's roll angle; the vertical channel controls the aircraft's vertical speed, and when the vertical speed command is 0, controls the aircraft to maintain its current altitude; and the heading channel controls the aircraft to perform coordinated turns.

[0046] The standby mode control law includes:

[0047] When the aircraft is in multi-rotor mode, the longitudinal channel controls the aircraft's pitch angle; the lateral channel controls the aircraft's roll angle; the vertical channel controls the throttle of the aircraft's rotors according to the stick value; and the heading channel controls the aircraft's yaw rate. When the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle.

[0048] When the aircraft is in bank transition mode, the longitudinal channel controls the forward speed of the aircraft; the lateral channel controls the roll angle of the aircraft; the vertical channel controls the throttle of the aircraft's rotors according to the stick value; and the heading channel controls the aircraft's yaw rate. When the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle.

[0049] When the aircraft is in fixed-wing mode, the longitudinal channel controls the throttle of the aircraft's fixed wings according to the stick value; the lateral channel controls the aircraft's ailerons; the vertical channel controls the aircraft's elevator; and the heading channel controls the aircraft to perform coordinated turns.

[0050] In the standby mode control law, when the aircraft is a manned aircraft, the lever amount is the lever amount of the throttle control stick in the cockpit; when the aircraft is an unmanned aircraft, the lever amount is the lever amount of the remote control control stick of the control terminal (such as a remote controller, ground station, etc.);

[0051] Further, such as Figure 2 As shown, the design architecture of the above control law includes an inner loop position loop, an inner loop attitude loop, an outer loop speed loop, an outer loop position loop, and an outer loop attitude loop;

[0052] wherein, based on the inner position loop, vertical velocity control of the aircraft is performed and a vertical velocity instruction is outputted; and the outer position loop performs altitude control of the aircraft according to the vertical velocity instruction; wherein the vertical velocity control may be performed according to proportional-integral control;

[0053] Based on the inner attitude loop, the aircraft's attitude angle and angular rate control are completed, and an angular rate instruction is output. The outer attitude loop completes the aircraft's attitude angle control according to the angular rate instruction. Among them, the differential term of the angular rate mainly affects the short period, which can improve the short-period damping of the attitude and suppress the oscillation amplitude of the angular rate. The proportional term of the attitude angle acts on the long period, which can increase the long-period damping ratio of the attitude, thereby improving the stability of the pitch channel. Unlike the conventional damping loop control method, the outer attitude loop in this embodiment adds an integral term of the pitch angle, thereby improving the control accuracy based on the pitch angle integral term, reducing the steady-state error, and enhancing the anti-interference performance.

[0054] The outer loop velocity loop adopts proportional-integral control to output acceleration instructions, and the acceleration instructions can be converted into attitude angle instructions (such as roll angle, pitch angle, etc.) of the aircraft system and input into the outer loop attitude loop, so that the integral term of the pitch angle can not only improve the speed control accuracy, but also reduce the influence of feedforward compensation uncertainty.

[0055] The control allocation system 45 is used to complete the control allocation according to the control allocation plan, wherein the control allocation plan includes:

[0056] Determine whether there is a failed rotor drive motor. If so, no throttle control amount distribution is required for the failed motor. The throttle control amount distribution is performed on the remaining valid motors as follows:

[0057] The throttles for the roll and pitch channels are pre-allocated according to the maximum throttle control amount, and the thrust is designed to meet the maximum throttle control amount of the roll channel and the maximum throttle control amount of the pitch channel;

[0058] When all rotors are operating, the rotational speed of the rotational drive motor of each rotor is obtained. When the rotational drive motor of a particular rotor is saturated, the unallocated throttle control amount is allocated to the remaining rotor rotational drive motors whose speeds are not saturated. The above-mentioned motor speed acquisition and throttle control amount allocation process to the motors whose speeds are not saturated is repeated until the speeds of a predetermined number (e.g., greater than or equal to 6) of the motors are saturated. The saturation includes high saturation, where the speed is greater than the maximum speed of the motor, and low saturation, where the speed is less than the minimum speed of the motor.

[0059] At the same time, if the rotation drive motor of a certain rotor is saturated, the throttle control amount of the roll channel and the pitch channel will be scaled proportionally, and the allocation of the throttle control amount of the heading channel will be abandoned. If the rotation drive motors of all rotors are not saturated, the throttle control amount of the heading channel will be maximized.

[0060] Therefore, the flight control system architecture in this embodiment can serve as a key system of a tilt-rotating configuration aircraft. Through the reasonable interaction logic between the navigation system, sensor system, servo system and flight control computer, as well as the design of control laws at different levels, it can improve the adaptability of the aircraft in complex environments while ensuring flight safety and reliability, and can effectively reduce design and use costs.

[0061] Example 2:

[0062] The only difference between this embodiment and embodiment 1 is that Figure 4 As shown, the switching of the control law includes:

[0063] (1) If the horizontal position signal is lost, the control law automatically switches from the primary mode to the secondary mode;

[0064] (2) If the horizontal position signal and the altitude signal are lost, the main mode control law is automatically switched to the backup mode control law, and the mode switch is manually switched to the backup mode corresponding to the backup mode control law to indicate that the backup mode control law has been switched to.

[0065] (3) If the altitude signal is lost, the secondary mode control law automatically switches to the backup mode control law, and the mode switch is manually switched to the backup mode corresponding to the backup mode control law to indicate that the backup mode control law has been switched to.

[0066] (4) Under the standby mode control law, the control law switching system 45 is also used to determine whether the aircraft meets the requirements of the main mode control law or the sub-mode control law. If the requirements of the main mode control law are met, the aircraft automatically switches from the standby mode control law to the main mode control law. If the requirements of the sub-mode control law are met, the aircraft automatically switches from the standby mode control law to the sub-mode control law.

[0067] Example 3:

[0068] The only difference between this embodiment and embodiment 1 is that the flight control system architecture further includes a communication system 5, and the communication system 5 includes one or more of a UHF communication system, a Tiantong communication system, a 4G / 5G communication system, a 5G communication system, and an L-link communication system connected to the flight control computer 4 via an RS422 serial port; at the same time, the communication system 5 is also connected to the flight control ground inspection port to complete ground debugging of the flight control system;

[0069] And, avionics system equipment 6, which includes one or more of a voice radio, an ADS-B link communication system, a joystick, a battery management system, a cockpit display system, and a power controller.

[0070] Further, such as Figure 4 As shown, the navigation system 1 is connected to the flight control computer 4 via the RS422 serial port, and the RTK data and airspeed data are transmitted to the navigation system 1 via the RS23 and RS422 serial ports. The main navigation subsystem in the navigation system 1 uses two RTK antennas (i.e., the main antenna and the backup antenna) separately, and the secondary navigation subsystem 12 and the tertiary navigation subsystem 13 share two RTK antennas through a power splitter;

[0071] The sensor system 2 is connected to the flight control computer 4 via an RS422 serial port;

[0072] The servo system 3, the power distribution box, and the domain controller are all connected to the flight control computer via the RS422 serial port; one or more of the engine, the generator, the motor electronic regulator, and the variable pitch propeller are connected to the flight control computer 4 via the CAN bus;

[0073] One or both of the navigation lights and the cockpit lighting system are connected to the flight control computer 4 via a DO interface;

[0074] The fuel gauge is connected to the flight control computer 4 via an RS422 serial port, etc., to obtain fuel quantity data from the AI ​​system of the flight control computer 4;

[0075] The flight control ground inspection port is connected to the flight control computer 4 through the RS232 serial port and the USB interface to complete program burning through the RS232 serial port and complete data downloading and uploading through the USB interface.

[0076] To sum up, the flight control system architecture logic design in the present invention is reasonable, and it can serve as a key system of a tilt-rotating configuration aircraft. Through the reasonable interaction logic between the navigation system, sensor system, servo system and flight control computer, as well as the design and switching of control laws at different levels, it can improve the adaptability of the aircraft in complex environments while ensuring flight safety and reliability, and can effectively reduce design and use costs.

[0077] It should be noted that the technical features in the above two embodiments can be combined in any way, and the technical solutions formed by the combination all belong to the scope of protection of this application. In this article, terms such as "including", "comprising" or any other variants thereof are intended to cover non-exclusive inclusion, so that the process, method, article or equipment including a series of elements includes not only those elements, but also includes other elements that are not explicitly listed, or also includes elements inherent to such process, method, article or equipment. In the absence of further restrictions, the elements defined by the sentence "including a..." do not exclude the presence of other identical elements in the process, method, article or equipment including the elements.

[0078] While embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that various changes, modifications, substitutions, and variations may be made to these embodiments without departing from the principles and spirit of the invention, and that the scope of the invention is defined by the appended claims and their equivalents.

Claims

1. A flight control system architecture for an eVTOL aircraft, characterized in that: include: Navigation systems, sensor systems, servo systems, and flight control computers; The navigation system includes a primary navigation subsystem, a secondary navigation subsystem, and a tertiary navigation subsystem, and all three navigation subsystems are capable of obtaining aircraft status and / or ground speed; The sensor system includes a primary data acquisition system, a backup data acquisition system, and an altitude sensor; the primary atmospheric data acquisition system and the backup atmospheric data acquisition system are both used to collect atmospheric data of the atmospheric environment in which the aircraft is located and / or flight data of the aircraft, and the altitude sensor is used to obtain the aircraft's altitude above the ground; The servo system includes a dual-redundancy servo controller, a dual-redundancy steering gear, a dual-redundancy power supply system, a tilt servo drive system and a control switching system; The flight control computer includes a control law switching system and a control distribution system; Furthermore, the navigation system, sensor system, and servo system are all connected to the flight control computer.

2. The flight control system architecture according to claim 1, wherein: The main navigation subsystem, secondary navigation subsystem and tertiary navigation subsystem are level-set according to the first priority, second priority and third priority from high to low, and the first-priority main navigation subsystem is called by default to obtain the aircraft status and / or ground speed. When the navigation subsystem of the previous priority fails, the navigation subsystem of the next priority is called in turn to obtain the aircraft status and / or ground speed.

3. The flight control system architecture according to claim 1, wherein: The main data acquisition system, main navigation subsystem and secondary navigation subsystem are level-set according to the first priority, second priority and third priority from high to low, and the first-priority main data acquisition system is called by default to obtain the altitude of the aircraft. When the system of the current priority fails, the system of the next priority is called in turn to obtain the altitude of the aircraft.

4. The flight control system architecture according to claim 1, wherein: The altitude sensor includes a radio altimeter and a millimeter-wave radar. The primary navigation subsystem, radio altimeter, and millimeter-wave radar are prioritized according to a first priority, a second priority, and a third priority, from high to low. By default, the first-priority primary navigation subsystem is first called to obtain the aircraft's landing altitude. If a system with the highest priority fails, the next-highest priority systems are called in sequence to obtain the aircraft's landing altitude. In addition, the main data acquisition system and the backup data acquisition system are level-set according to the first priority and the second priority from high to low, and the main data acquisition system with the first priority is called first by default to obtain the airspeed. When the system with the current priority fails, the system with the next priority is called to obtain the airspeed.

5. The flight control system architecture according to claim 1, wherein: The dual-redundant servos are independently driven by two drive channels; The dual-redundant servo controller has two independent control units, each of which includes a control circuit and a control program, and controls a corresponding drive channel of the dual-redundant servo to perform a corresponding action; The dual-redundancy power supply system is connected to the dual-redundancy servo controller and / or the dual-redundancy steering gear, and is used to provide independent dual-path power supplies for the dual-redundancy servo controller and / or the dual-redundancy steering gear; The control switching system is connected to the dual-redundant servo controller and is used to detect the working status of the dual-redundant servo controller, and when one control unit of the dual-redundant servo controller fails, automatically switch to the other control unit to operate, and / or is connected to the dual-redundant steering gear and is used to detect the working status of the dual-redundant steering gear, and when one drive channel of the dual-redundant steering gear fails, automatically switch to the other drive channel to operate, and / or is connected to the dual-redundant power supply system and is used to detect the working status of the dual-redundant power supply system, and when one power supply of the dual-redundant steering gear fails, automatically switch to the other power supply to operate; The tilt servo drive system is used to drive the tilt rotor to perform a tilting action, so that the aircraft can switch between a multi-rotor mode and a fixed-wing mode.

6. The flight control system architecture according to claim 1, wherein: The control law switching system is connected to one or more of the navigation system, sensor system, and servo system, and is used to complete the switching of the control law according to the working state changes of one or more of the navigation system, sensor system, and servo system, and generate corresponding control law instructions, and control one or more of the navigation system, sensor system, and servo system to perform corresponding actions according to the control law instructions.

7. The flight control system architecture according to claim 6, wherein: The control law includes a primary mode control law, a secondary mode control law and a backup mode control law; Wherein, the main mode control law includes: When the aircraft is in multi-rotor mode, in the longitudinal channel, the forward speed of the aircraft is controlled, and when the forward speed command is 0, the aircraft is controlled to maintain its current position; in the transverse channel, the lateral speed of the aircraft is controlled, and when the lateral speed command is 0, the aircraft is controlled to maintain its current position; in the vertical channel, the vertical speed of the aircraft is controlled, and when the vertical speed command is 0, the aircraft is controlled to maintain its current altitude; and in the heading channel, the yaw rate of the aircraft is controlled, and when the yaw angle command is 0, the aircraft is controlled to maintain its current heading angle; When the aircraft is in the bank transition mode, the forward speed of the aircraft is controlled on the longitudinal channel; the roll angle of the aircraft is controlled on the lateral channel; the vertical speed of the aircraft is controlled on the vertical channel, and when the vertical speed command is 0, the aircraft is controlled to maintain the current altitude; and the yaw rate of the aircraft is controlled on the heading channel, and when the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle. When the aircraft is in fixed-wing mode, the longitudinal channel controls the aircraft's airspeed; the transverse channel controls the aircraft's lateral speed, and when the lateral speed command is 0, controls the aircraft to maintain its current position; the vertical channel controls the aircraft's vertical speed, and when the vertical speed command is 0, controls the aircraft to maintain its current altitude; and the heading channel controls the aircraft to perform coordinated turns. The sub-mode control law includes: When the aircraft is in multi-rotor mode, the longitudinal channel controls the pitch angle of the aircraft; the lateral channel controls the roll angle of the aircraft; the vertical channel controls the vertical speed of the aircraft, and when the vertical speed command is 0, the aircraft is controlled to maintain the current altitude; and the heading channel controls the yaw rate of the aircraft, and when the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle. When the aircraft is in the bank transition mode, the forward speed of the aircraft is controlled on the longitudinal channel; the roll angle of the aircraft is controlled on the lateral channel; the vertical speed of the aircraft is controlled on the vertical channel, and when the vertical speed command is 0, the aircraft is controlled to maintain the current altitude; and the yaw rate of the aircraft is controlled on the heading channel, and when the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle. When the aircraft is in fixed-wing mode, the longitudinal channel controls the aircraft's airspeed; the lateral channel controls the aircraft's roll angle; the vertical channel controls the aircraft's vertical speed, and when the vertical speed command is 0, controls the aircraft to maintain its current altitude; and the heading channel controls the aircraft to perform coordinated turns. The standby mode control law includes: When the aircraft is in multi-rotor mode, the longitudinal channel controls the aircraft's pitch angle; the lateral channel controls the aircraft's roll angle; the vertical channel controls the throttle of the aircraft's rotors according to the stick value; and the heading channel controls the aircraft's yaw rate. When the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle. When the aircraft is in bank transition mode, the longitudinal channel controls the forward speed of the aircraft; the lateral channel controls the roll angle of the aircraft; the vertical channel controls the throttle of the aircraft's rotors according to the stick value; and the heading channel controls the aircraft's yaw rate. When the yaw angle command is 0, the aircraft is controlled to maintain the current heading angle. When the aircraft is in fixed-wing mode, the longitudinal channel controls the throttle of the aircraft's fixed wings according to the stick value; the lateral channel controls the aircraft's ailerons; the vertical channel controls the aircraft's elevator; and the heading channel controls the aircraft to perform coordinated turns. In the standby mode control law, when the aircraft is a manned aircraft, the lever amount is the lever amount of the cockpit throttle joystick; when the aircraft is an unmanned aircraft, the lever amount is the lever amount of the remote control joystick of the control terminal (such as a remote controller, ground station, etc.).

8. The flight control system architecture according to claim 1, wherein: The design framework of the control law includes an inner loop position loop, an inner loop attitude loop, an outer loop speed loop, an outer loop position loop, and an outer loop attitude loop; The inner position loop controls the vertical velocity of the aircraft and outputs a vertical velocity instruction, and the outer position loop controls the altitude of the aircraft according to the vertical velocity instruction. The inner attitude loop completes the attitude angle and angular rate control of the aircraft and outputs an angular rate instruction. The outer attitude loop completes the attitude angle and angle control of the aircraft according to the angular rate instruction. The outer loop speed loop is used to output an acceleration instruction, and the acceleration instruction can be converted into an attitude angle instruction of the machine system and input into the outer loop attitude loop.

9. The flight control system architecture according to claim 1, wherein: The control allocation system is used to complete control allocation according to a control allocation plan, wherein the control allocation plan includes: Determine whether there is a failed rotor drive motor. If so, no throttle control amount distribution is required for the failed motor. The throttle control amount distribution is performed on the remaining valid motors as follows: The throttles for the roll and pitch channels are pre-allocated according to the maximum throttle control amount, and the thrust is designed to meet the maximum throttle control amount of the roll channel and the maximum throttle control amount of the pitch channel; When all rotors are working, the speed of the rotation drive motor of each rotor is obtained. When the rotation drive motor of a certain rotor is saturated, the unallocated throttle control amount is allocated to the remaining rotor rotation drive motors whose speeds are not saturated. The above-mentioned motor speed acquisition and throttle control amount allocation process to the motors whose speeds are not saturated are repeated until the speeds of a predetermined number of motors are saturated.

10. The flight control system architecture according to claim 1, wherein: The control allocation scheme also includes: if the rotation drive motor of a certain rotor is saturated, the throttle control amount of the roll channel and the pitch channel will be scaled proportionally, and the allocation of the throttle control amount of the heading channel will be abandoned. If the rotation drive motors of all rotors are not saturated, the throttle control amount of the heading channel will be maximized.

11. The flight control system architecture according to claim 1, wherein: The switching of the control law includes: (1) If the horizontal position signal is lost, the control law automatically switches from the primary mode to the secondary mode; (2) If the horizontal position signal and the altitude signal are lost, the control law of the primary mode is automatically switched to the backup mode control law; (3) If the altitude signal is lost, the secondary mode control law automatically switches to the backup mode control law.

12. The flight control system architecture according to claim 11, wherein: Under the backup mode control law, it is determined whether the aircraft meets the requirements of the main mode control law or the secondary mode control law. If the requirements of the main mode control law are met, the backup mode control law is automatically switched to the main mode control law. If the requirements of the secondary mode control law are met, the backup mode control law is switched to the secondary mode control law.

13. The flight control system architecture according to claim 1, wherein: The flight control computer further includes a CPU board, a power board, and an interface board, wherein the CPU board and / or the power board adopt a triple-redundancy design, the power board is connected to the CPU board to supply power to the CPU board, and the interface board is used to connect the CPU board and other external devices; At least two of the navigation system, sensor system, servo system and flight control computer are connected through a cross-channel data link to achieve redundancy resource sharing; at the same time, one or more of the CPU board, power board and interface board are provided with a BIT circuit.

14. The flight control system architecture according to claim 1, wherein: The eVTO aircraft may adopt a tilt-rotor configuration.

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