Amphibious aircraft fly-by-wire flight control system and working method

Abstract

The application belongs to the technical field of amphibious aircraft flight control, and discloses an amphibious aircraft fly-by-wire flight control system and a working method. A cockpit control device is connected to and controls normal instruction sensors and backup instruction sensors. The normal instruction sensors transmit signals to actuator control electronics. The actuator control electronics and a flight control computer transmit signals to each other. The actuator control electronics and an actuator control module transmit signals to each other. The actuator control electronics also receive signals from external signal sensors, direct angular rate sensors, flap position sensors and a flight control mode switch. The backup instruction sensors transmit signals to a backup flight control computer. The backup flight control computer transmits signals to the actuator control module. The actuator control module is connected to and controls actuators. Angular displacement sensors are arranged on the actuators and transmit collected signals to the actuator control module. The application solves the problem that current fly-by-wire flight control systems do not match the cross-medium application of amphibious aircraft.

Description

Technical Field

[0001] This invention belongs to the field of amphibious aircraft flight control technology, and relates to a fly-by-wire aircraft control architecture system, specifically to an amphibious aircraft fly-by-wire flight control system and its working method. Background Technology

[0002] With the continuous development of aviation technology, aircraft flight control systems have undergone several stages of development: from early simple mechanical control to stability augmentation and control augmentation systems, then to partial authority fly-by-wire systems, and gradually upgrading to full authority fly-by-wire systems, with future development towards fly-by-optical systems. Currently, the application scenarios and industrial scale of full authority fly-by-wire systems are rapidly expanding, but they also face some challenges in architectural design and physical implementation. Limited by the current technological level of the electronics industry, the failure probability of single-channel electronic devices is basically stagnant at 10⁻³ to 10⁻⁴. To meet the airworthiness safety requirements of civil aircraft with a failure probability below 10⁻⁹, the industry typically uses two or more single-channel electronic devices for the same function, designing them as two or more independent control channels, and considering safety and ease of implementation, designing two or more control operating modes. The ultimate backup system, as the last guarantee for safe flight, has always been a hot topic in this field.

[0003] The A320, the world's first large passenger aircraft to employ a fly-by-wire flight control system, consists of nine computers and corresponding actuators. Two elevator and aileron computers (ELACs) control the aileron, elevator, and horizontal stabilizer actuators; three spoiler and elevator computers (SECs) control all spoiler actuators and serve as backups for elevator and horizontal stabilizer control; two flight augmentation computers (FACs) provide yaw damping, trim, and travel limiting functions; and mechanical devices are incorporated into both the horizontal stabilizer and rudder controls as part of a mechanical backup system.

[0004] As Boeing's first aircraft to adopt a fly-by-wire flight control system, the B777's architecture includes three main flight control computers (PFCs), four actuator control electronics (ACEs), and the ARINC 629 bus. The ACEs convert pilot controller position and control surface position signals into digital quantities, which are then uploaded to the PFCs via the ARINC 629 bus. The main function of the PFCs is to process the atmospheric data collected by the system and the input signals from various sensors, combining them with the control laws pre-written in the PFCs. Based on the position feedback signals, it calculates control surface commands and sends them to the ACEs via the ARINC 629 bus. The ACEs then convert these commands into analog signals to directly control the actuators, causing the control surfaces to deflect. The B777 also employs a mechanical backup system, allowing it to control the horizontal stabilizer and two pairs of multi-function spoilers to deflect as needed when all PFCs and ACEs fail.

[0005] The aforementioned fly-by-wire flight control system architectures all retain some mechanical components as a final backup, with only pitch / yaw or pitch / roll axes having control functions. If the flight control computer fails to restart in the air, the aircraft will have to make an emergency landing due to the lack of three-axis control capabilities. With the exposure of the mechanical backup problem and the decline of the mechanical lever industry, the A320 and subsequent B777 models, as well as the C919, have generally adopted a design without mechanical backup. However, their backup control functions are either integrated into an existing secondary computer with a dissimilar design (CN 116767484 A), or the backup computer completely takes over the control channel of a secondary computer (CN112124568 B). Both have disadvantages such as difficulty in implementation, high failure risk, and complex control channel reconfiguration logic, and cannot meet the design requirements of amphibious aircraft for surface mission control functions. Summary of the Invention

[0006] To address this, the present invention provides a fly-by-wire flight control system and its operating method for amphibious aircraft, overcoming the shortcomings of existing fly-by-wire flight control system architectures and adding surface mission control functions, thus filling a gap in the design of fly-by-wire flight control systems for amphibious aircraft in China. It reuses some existing actuator control devices to construct an additional, single-redundant, small fly-by-wire control channel as a final backup, capable of achieving minimum acceptable control. This channel has three-axis backup control authority instead of the existing two-axis backup, ensuring the aircraft can continue to fly and land safely. The calculation and switching of control laws for different operating modes are executed by the relevant flight control computer according to the set logic, avoiding command disputes during control channel reconfiguration.

[0007] The technical solution of the present invention is as follows: A fly-by-wire flight control system for an amphibious aircraft includes cockpit controls, a normal command sensor, a backup command sensor, actuator control electronics, a flight control computer, a flap position sensor, a backup flight control computer, an actuator control module, actuators, external signal sensors, a direct angular rate sensor, an angular displacement sensor, and a flight control mode switch. The cockpit controls connect to and control the normal command sensor and the backup command sensor. The normal command sensor transmits signals to the actuator control electronics. The actuator control electronics and the flight control computer exchange signals. The actuator control electronics and the actuator control module exchange signals. The actuator control electronics also receive signals from the external signal sensors, the direct angular rate sensor, the flap position sensor, and the flight control mode switch. The backup command sensor transmits signals to the backup flight control computer, which then transmits signals to the actuator control module. The actuator control module connects to and controls the actuators. The angular displacement sensor is mounted on the actuators and transmits the collected signals to the actuator control module.

[0008] Furthermore, the normal command sensor collects command signals from the cockpit control devices reflecting the pilot's control intentions and transmits them to the actuator control electronics; the external signal sensor collects dynamic pressure, static pressure, temperature, aircraft attitude, and position information and uploads it to the actuator control electronics; the direct angular rate sensor collects the aircraft's three-axis angular rate information and transmits it to the actuator control electronics; the flap position sensor collects flap position signals and transmits them to the actuator control electronics; the actuator control electronics transmits the received signals to the flight control computer, simultaneously calculates the direct control law, receives normal or degraded control commands from the flight control computer, and selects one type of control command from different operating modes to send to the actuator control module according to the operating mode switching logic.

[0009] Furthermore, after receiving the control command, the actuator control module converts it into a servo control command for the actuator according to a predetermined logic; the actuator receives the servo control command from the actuator control module and performs telescopic movement under the pressure of the hydraulic source, causing the control surfaces to deflect; the angular displacement sensor measures the corresponding control surface deflection and feeds it back to the actuator control module, which then uploads it to the flight control computer via the actuator control electronics.

[0010] Furthermore, the actuator control module adopts a command and monitoring architecture, including a type A module and a type B module. The type A module only accepts control surface commands from the actuator control electronics, while the type B module can accept control surface commands from both the actuator control electronics and the backup flight control computer.

[0011] Furthermore, there are six normal command sensors, all with four redundancies, installed below the left and right control units respectively; three backup command sensors are installed below the left control unit; external signal sensors include a three-redundant atmospheric data reference unit, a one-redundant attitude reference unit, a two-redundant inertial reference unit, a two-redundant radio altimeter, and a four-redundant angle-of-attack sensor; four direct angular rate sensors are configured, installed symmetrically in the middle of the fuselage, outputting four-redundant three-axis angular rate data for degraded control law calculation; four actuator control electronics units, using the same design and interchangeable, each with a command channel and a monitoring channel, without signal interaction between them; three flight control computers, using the same design and interchangeable, each with a command channel and a monitoring channel, with signal interaction between them; and one backup flight control computer, with a command channel and a monitoring channel.

[0012] A method for operating a fly-by-wire flight control system for an amphibious aircraft, using the aforementioned fly-by-wire flight control system for an amphibious aircraft, has four operating modes, including normal operating mode, degraded operating mode, direct operating mode, and backup operating mode. The priority of the four operating modes is as follows: normal operating mode is the first priority, degraded operating mode is the second priority, direct operating mode is the third priority, and backup operating mode is the fourth priority.

[0013] Furthermore, in normal operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the actuator control electronics via the SIB bus, obtains the voted pilot commands, and calculates the control commands for the normal mode control law by combining information on inertial navigation, angle of attack, atmospheric conditions, and flap positions. Normal operating mode is the system's default operating state. At this time, the fly-by-wire flight control system can achieve ideal control functions, including three-axis stabilization control, envelope protection, coordinated turn function, and control surface deflection angle limitation function.

[0014] Furthermore, in degraded operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the actuator control electronics via the SIB bus, obtains the voted pilot commands, and calculates the control commands of the degraded mode control law by combining the signals from the direct angular rate sensor and the flap position sensor, and loses the envelope protection function and control surface limitation function.

[0015] Furthermore, in direct operating mode, the actuator control electronics calculate the control command of the direct mode control law based on the normal command signals it receives, combined with the flap extension and retraction state quantities, and further lose the three-axis stabilization control function and coordinated turning function.

[0016] Furthermore, in backup mode, the backup flight control computer will calculate the control commands of the backup mode control law based on the received pilot backup command signal and send them to the actuator control module, further losing the control function of all spoilers.

[0017] Technical effects: 1. This invention adopts a reasonable configuration of flight control computer, actuator control electronics, backup control computer and actuator control module, and designs a fly-by-wire flight control system architecture and working mode suitable for amphibious aircraft, which makes up for the shortcomings of existing fly-by-wire flight control system architectures, such as difficulty in implementation, high failure risk and complex control channel reconfiguration logic.

[0018] 2. This invention employs a method of using a backup computer instead of a non-similar actuator control electronics design to eliminate the risk of system common mode. This avoids safety issues that may arise from non-similar actuator control electronics designs, such as compatibility, command integrity, and fault monitoring, which could lead to erroneous disconnection of actuator control electronics, frequent restarts, and disruptive alarms. At the same time, it reuses some existing actuator control equipment from the normal flight control system to construct an additional single-channel small fly-by-wire control channel capable of achieving minimum acceptable control as a final backup. This channel has three-axis backup control authority, ensuring that the aircraft can continue to fly and land safely.

[0019] 3. The four different working mode control law calculations and switching designed in this invention are executed by the relevant flight control computer according to the set logic, which solves the problem of command disputes during control channel reconfiguration.

[0020] 4. This invention uses a modular electronic backup system to replace the traditional mechanical backup system, which can reduce the weight of the system to the maximum extent while meeting the safety-related airworthiness requirements of CCAR25.671, 25.1309 and other regulations.

[0021] 5. The fly-by-wire flight control system and working mode design provided by this invention have clear logic, refined structure, reasonable layout, simple control mode switching logic, and are easy to implement in engineering. It creatively solves the engineering problem that the current fly-by-wire flight control system of transport aircraft is not compatible with the cross-media application scenarios of amphibious aircraft. Attached Figure Description

[0022] Figure 1 This is a schematic diagram of a fly-by-wire flight control system according to an embodiment of the present invention.

[0023] Figure 2 This is a schematic diagram of the actuator control electronics and peripheral equipment configuration according to an embodiment of the present invention.

[0024] Figure 3 This is a schematic diagram of actuator control and energy configuration according to an embodiment of the present invention.

[0025] Figure 4 This is a schematic diagram of the working mode switching of a fly-by-wire flight control system according to an embodiment of the present invention.

[0026] Among them, 1—cockpit control devices, 2—backup command sensor, 3—backup flight control computer, 4—actuator, 5—angular displacement sensor, 6—actuator control module, 7—flap position sensor, 8—automatic flight control computer, 9—flight control computer, 10—indication and recording system, 11—direct angular rate sensor, 12—flight control mode switch, 13—actuator control electronics, 14—external signal sensor, and 15—normal command sensor. Detailed Implementation

[0027] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. Numerous specific design details are set forth in the following detailed description to provide a more complete understanding of the invention. However, it will be apparent to those skilled in the art that the invention can be practiced without some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setup and method set forth below, but covers any improvements, substitutions, and modifications to the structures, methods, and devices without departing from the spirit of the invention. In the drawings and the following description, any parts not exhaustively described are considered to be common knowledge or conventional practices in the art.

[0028] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0029] Example 1:

[0030] A fly-by-wire flight control system for an amphibious aircraft includes a cockpit control unit 1, a normal command sensor 2, a backup command sensor 15, actuator control electronics 13, a flight control computer 9, a flap position sensor 7, a backup flight control computer 3, an actuator control module 6, actuators 4, an external signal sensor 14, a direct angular rate sensor 11, an angular displacement sensor 5, and a flight control mode switch 12. The cockpit control unit 1 connects to and controls the normal command sensor 2 and the backup command sensor 15. The normal command sensor 2 transmits signals to the actuator control electronics 13, which controls the actuators. Electronics 13 and flight control computer 9 exchange signals. Actuator control electronics 13 and actuator control module 6 exchange signals. Actuator control electronics 13 also receives signals from external signal sensor 14, direct angular rate sensor 11, flap position sensor 7 and flight control mode switch 12. Backup command sensor 15 transmits signals to backup flight control computer 3. Backup flight control computer 3 transmits signals to actuator control module 6. Actuator control module 6 connects to and controls actuator 4. Angular displacement sensor 5 is installed on actuator 4 and transmits the collected signals to actuator control module 6.

[0031] Normal command sensor 2 collects command signals from cockpit control devices 1 reflecting the pilot's control intentions and transmits them to actuator control electronics 13; external signal sensor 14 collects dynamic pressure, static pressure, temperature, aircraft attitude and position information and uploads it to actuator control electronics 13; direct angular rate sensor 11 collects aircraft three-axis angular rate information and transmits it to actuator control electronics 13; flap position sensor collects flap position signals and transmits them to actuator control electronics 13; actuator control electronics 13 transmits the received signals to flight control computer 9, simultaneously calculates the direct control law, receives normal or degraded control commands issued by flight control computer 9, and selects one type of control command from different working modes to issue to actuator control module 6 according to the working mode switching logic.

[0032] After receiving the control command, the actuator control module 6 converts it into a servo control command for the actuator 4 according to a predetermined logic; the actuator 4 receives the servo control command from the actuator control module 6 and performs telescopic movement under the pressure of the hydraulic source, causing the control surface to deflect; the angular displacement sensor measures the corresponding control surface deflection and feeds it back to the actuator control module 6, which then uploads it to the flight control computer 9 via the actuator control electronics 13.

[0033] The actuator control module 6 adopts a command and monitoring architecture, including a type A module and a type B module. The type A module only accepts control surface commands from the actuator control electronics 13, while the type B module can accept control surface commands from both the actuator control electronics 13 and the backup flight control computer 3.

[0034] There are six normal command sensors 2, all with 4 redundancy, installed below the left and right control devices respectively; there are three backup command sensors 15, installed below the left control device; external signal sensors 14 include a 3-redundant atmospheric data reference unit, a 1-redundant attitude reference unit, a 2-redundant inertial reference unit, a 2-redundant radio altimeter, and a 4-redundant angle of attack sensor; four direct angular rate sensors 11 are configured, installed symmetrically in the middle of the fuselage, outputting 4-redundant three-axis angular rate data for degraded control law calculation; there are four actuator control electronics 13, all with the same design and interchangeability, each with a command channel and a monitoring channel, without signal interaction between them; there are three flight control computers 9, all with the same design and interchangeability, each with a command channel and a monitoring channel, with signal interaction between them; there is one backup flight control computer 3, with a command channel and a monitoring channel.

[0035] A method for operating a fly-by-wire flight control system for an amphibious aircraft, using the aforementioned fly-by-wire flight control system for an amphibious aircraft, has four operating modes, including normal operating mode, degraded operating mode, direct operating mode, and backup operating mode. The priority of the four operating modes is as follows: normal operating mode is the first priority, degraded operating mode is the second priority, direct operating mode is the third priority, and backup operating mode is the fourth priority.

[0036] In normal operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the actuator control electronics via the SIB bus, obtains the voted pilot commands, and calculates the control commands of the normal mode control law by combining information on inertial navigation, angle of attack, atmosphere, and flap position. Normal operating mode is the system's default operating state. At this time, the fly-by-wire flight control system can achieve ideal control functions, including three-axis stabilization control, envelope protection, coordinated turn function, and control surface deflection angle limitation function.

[0037] In degraded operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the actuator control electronics via the SIB bus, obtains the voted pilot commands, and calculates the control commands of the degraded mode control law by combining the signals from the direct angular rate sensor and the flap position sensor, and loses the envelope protection function and control surface limitation function.

[0038] In direct operating mode, the actuator control electronics calculate the control command of the direct mode control law based on the normal command signals it receives and the flap extension / retraction state quantities, and further lose the three-axis stabilization control function and coordinated turning function.

[0039] In backup mode, the backup flight control computer will calculate the control commands of the backup mode control law based on the received pilot backup command signal and send them to the actuator control module, further losing the control function of all spoilers.

[0040] Example 2:

[0041] A fly-by-wire flight control system and its operating modes for an amphibious aircraft are disclosed. The system architecture comprises a cockpit control unit 1, a first command sensor 2, a second command sensor 15, actuator control electronics 13, a flight control computer 9, an automatic flight control computer 8, a flap position sensor 7, a backup flight control computer 3, an actuator control module 6, actuators 4, an external signal sensor 14, a direct angular rate sensor 11, an angular displacement sensor 5, a flight control mode switch 12, and an instruction recording system 10. To reduce common-mode risk and control channel reconfiguration complexity, four operating modes are designed: normal operating mode, degraded operating mode, direct operating mode, and backup operating mode. Control law calculation and switching between different operating modes are executed by the relevant flight control computer according to set logic, avoiding command disputes during control channel reconfiguration. Among them: The normal command sensor collects command signals from the cockpit controls reflecting the pilot's control intentions and transmits them to the actuator control electronics; the external signal sensor collects information such as dynamic pressure, static pressure, temperature, aircraft attitude, and position and transmits it to the actuator control electronics; the direct angular rate sensor collects the aircraft's three-axis angular rate information and transmits it to the actuator control electronics; the flap position sensor collects flap position signals and transmits them to the actuator control electronics; the actuator control electronics transmits the above signals to the flight control computer, simultaneously calculates the direct control law, receives normal / degraded control commands from the flight control computer, and selects the appropriate control mode based on the operating mode switching logic. A control command of one of the following types—normal, degraded, or direct—is issued to the actuator control module; the backup flight control computer calculates the backup control command according to the backup command sensor signal and a predetermined control logic, and issues it to the actuator control module; upon receiving the control command, the actuator control module converts it into a servo control command for the actuator according to a predetermined logic; the actuator receives the servo control command from the actuator control module and performs an extension / retraction movement under the pressure of a hydraulic source, causing the control surface hinged to it to deflect; the angular displacement sensor measures the deflection of the corresponding control surface and feeds it back to the actuator control module, which then uploads it to the flight control computer via the actuator control electronics.

[0042] The flight control computer receives signals uploaded by the actuator control electronics, performs redundancy voting, simultaneously calculates normal / degraded control laws according to predetermined control logic, and sends them to the actuator control electronics; determines whether to respond to the automatic flight control computer command; and sends relevant status and fault information of the flight control system electronic equipment to the instruction recording system for instruction and alarm.

[0043] The flight control mode switch is used by the pilot to manually switch between direct working mode and elevation mode.

[0044] The actuator control module is divided into type A and type B, and adopts a command / monitoring architecture. The type A actuator control module only accepts control surface commands from the actuator control electronics, while the type B actuator control module can accept control surface commands from both the actuator control electronics and the backup flight control computer.

[0045] There are 6 normal command sensors, all with 4 redundancy, installed under the left and right control devices respectively; there are 3 backup command sensors, all with 1 redundancy, installed only under the left control device.

[0046] The external signal sensors include a triplet atmospheric data reference unit, a single-redundant attitude reference unit, a dual-redundant inertial reference unit, a dual-redundant radio altimeter, and a quadruple-redundant angle-of-attack sensor.

[0047] Four direct angular rate sensors are configured and installed symmetrically in the middle of the fuselage, outputting a total of four redundancies of the aircraft's three-axis angular rate data for degraded control law calculation.

[0048] The actuator control electronics are configured with four units, which adopt the same design and are interchangeable. Each unit is equipped with a command channel and a monitoring channel to prevent the output of erroneous control commands. There is no signal interaction between them, and they supply power to the normal command sensors, flight control computer and actuator control module.

[0049] The flight control computer is configured with three units, which adopt the same design and are interchangeable. Each unit is equipped with a command channel and a monitoring channel to prevent the output of erroneous control commands. The two channels are designed based on different processor models to eliminate the risk of common mode. They have signal interaction with each other and with all actuator control electronics.

[0050] The backup flight control computer is configured with one unit, which supplies power to the backup command sensor and has a command channel and a monitoring channel to prevent the output of erroneous control commands. Based on the backup command sensor signal, it calculates the backup control command according to the predetermined control logic and sends it to the B-type actuator control module to eliminate the common mode risk caused by the similar design of the actuator control electronics.

[0051] The actuator control module is configured with 19 units, 13 of which are type A and 6 of type B. The two types of actuator control modules use different processors for dissimilar design. Each actuator control module has an instruction channel and a monitoring channel to prevent the output of erroneous control commands. The same model is interchangeable.

[0052] The system is equipped with 19 actuators: 2 for each elevator and aileron, 3 for each rudder, 1 for each spoiler, and 2 for each surface control surface. Each actuator on a control surface is capable of driving that control surface to work normally, and each actuator is fixedly controlled by an actuator control module.

[0053] Four angular displacement sensors are configured and installed on the left / right elevators and left / right ailerons, respectively. They are powered by the B-type actuator control module to measure the deflection of the corresponding control surfaces.

[0054] One flight control mode switch is provided, which is mainly used by the pilot to manually switch the flight control system to direct working mode and ascend mode.

[0055] The automatic flight control computer is configured with one unit and adopts a 3-channel design, with each channel communicating with one flight control computer. After receiving the autopilot activation command, it generates automatic flight commands according to the predetermined control logic and transmits them to the flight control computer. When the aircraft status allows, the flight control computer controls each control surface to complete the autopilot.

[0056] Two flap position sensors are configured, one on each of the left and right flaps, to collect flap position signals and send them to the actuator control electronics.

[0057] The actuator control electronics, flight control mode switch, direct angular rate sensor, backup flight control computer, flap position sensor, automatic flight control computer, external signal sensor, and indication and recording system all require power from the onboard power system to function properly, and the actuators all require pressure from the onboard hydraulic power source system to function properly.

[0058] The priority of the four operating modes of the fly-by-wire flight control system is as follows: normal operating mode > degraded operating mode > direct operating mode > backup operating mode. The control law calculation and switching between different operating modes are executed by the relevant flight control computer according to the set logic to avoid command disputes during control channel reconfiguration. The same control surface cannot execute two or more operating mode control commands of the fly-by-wire flight control system at the same time, but different control surfaces can execute different control commands at the same time. When a higher-level operating mode is available, the system cannot automatically upgrade to the higher-level mode, but will issue a high-level flight control operating mode availability alarm to the crew, and the pilot must manually rotate the flight control mode switch to manually upgrade to the higher-level mode.

[0059] The instruction recording system sends mode alarm information to the crew when the flight control system switches to different operating modes.

[0060] Example 3:

[0061] This invention discloses a fly-by-wire flight control system and its operating modes for amphibious aircraft. By rationally configuring command sensors, flight control computers, actuator control electronics, external signal sensors, actuator control modules, backup flight control computers, and actuators, it achieves cross-platform control of amphibious aircraft while minimizing system weight, while meeting the airworthiness requirements of safety-related regulations such as CCAR25671 and 251309. This invention reuses some existing actuator control devices to construct an additional, single-redundant, small fly-by-wire control channel as a final backup, providing three-axis backup control authority instead of the existing two-axis backup, ensuring the aircraft can continue to fly and land safely. Considering factors such as common-mode risk, functional implementation, design cost, and implementation difficulty, this invention designs four operating modes: normal, degraded, direct, and backup. The control law calculation and switching between different operating modes are executed by the relevant flight control computer according to set logic, avoiding command disputes during control channel reconfiguration.

[0062] The amphibious aircraft has an aileron and three spoilers on each of its left and right wings, an elevator on each of its left and right horizontal stabilizers, a rudder behind the vertical stabilizer, and a water rudder on the lower part of the tail fuselage.

[0063] like Figure 1 The diagram shown is a schematic block diagram of the fly-by-wire flight control system of the present invention. The fly-by-wire flight control system architecture includes a cockpit control device 1, a command sensor, an actuator control electronics 13, a flight control computer 9, an automatic flight control computer 8, a flap position sensor 7, a backup flight control computer 3, an actuator control module 6, an actuator 4, an external signal sensor 14, a direct angular rate sensor 11, an angular displacement sensor 5, a flight control mode switch 12, and an indication and recording system 10.

[0064] like Figure 1 As shown, the actuator control electronics will Figure 2 The collected signals are uploaded to the flight control computer via the SIB bus. The system also calculates the direct control law according to the predetermined control logic, and simultaneously receives normal / degraded control commands from the flight control computer via the SIB bus. Figure 4 The aforementioned working mode switching method sends one of the control commands—normal, degraded, or direct—to the actuator control module via the ADB bus.

[0065] like Figure 1 As shown, the flight control computer performs redundancy voting on the signals uploaded by the actuator control electronics, calculates normal / degraded control commands according to the predetermined control logic and sends them to the actuator control electronics, monitors the status of the flight control electronics equipment, and sends the status information to the instruction recording system.

[0066] like Figure 1As shown, the flight control computer will determine whether to respond to the instructions of the automatic flight control computer.

[0067] like Figure 1 As shown, the backup flight control computer supplies power to the backup command sensor, calculates backup control commands based on the backup command sensor signals, and sends them to the actuator control module via hardwired connection.

[0068] like Figure 1 As shown, one flight control mode switch is configured to receive the pilot's command to switch the main flight control system to "automatic" or "direct", allowing the pilot to manually switch to direct working mode and perform mode-up operations.

[0069] like Figure 2 The diagram shown is a schematic of the actuator control electronics and peripheral equipment configuration of the present invention. The six quadruple redundancy normal command sensors are defined according to their installation positions as follows: left longitudinal normal command sensor, left lateral normal command sensor, left heading normal command sensor, longitudinal backup command sensor, lateral backup command sensor, heading backup command sensor, right longitudinal normal command sensor, right lateral normal command sensor, and right heading normal command sensor. The left longitudinal normal command sensor has 2% redundancy connected to actuator control electronics 1 and the remaining 2% redundancy connected to actuator control electronics 2. The right longitudinal normal command sensor has 2% redundancy connected to actuator control electronics 3 and the remaining 2% redundancy connected to actuator control electronics 4. The left lateral normal command sensor has 2% redundancy connected to actuator control electronics 1 and the remaining 2% redundancy connected to actuator control electronics 3. The right lateral normal command sensor has 2% redundancy connected to actuator control electronics 2 and the remaining 2% redundancy connected to actuator control electronics 4. The left and right heading normal command sensors have 3% redundancy connected to actuator control electronics 1, actuator control electronics 2, and actuator control electronics 4, respectively. The remaining 1% redundancy is used by the front wheel steering system. The four actuator control electronics use the same design, are interchangeable, and do not interact with each other.

[0070] like Figure 2 As shown, among the external signal sensors, atmospheric data reference unit 1 is connected to actuator control electronics 1, atmospheric data reference unit 2 is connected to actuator control electronics 3, and atmospheric data reference unit 3 is connected to actuator control electronics 4; inertial reference unit 1 is connected to actuator control electronics 1, inertial reference unit 2 is connected to actuator control electronics 2, and attitude reference unit is connected to actuator control electronics 4; radio altimeter 1 is connected to actuator control electronics 2, and radio altimeter 2 is connected to actuator control electronics 4; angle of attack sensors 1 to 4 are connected one-to-one with actuator control electronics 1 to 4 (Note: the device numbers are only for ease of description and have no meaning in terms of order).

[0071] like Figure 2As shown, the flap position sensor 1 is connected to actuator control electronics 1 and actuator control electronics 3, and the flap position sensor 2 is connected to actuator control electronics 1 to 4; the direct angular rate sensors 1 to 4 are connected one-to-one with actuator control electronics 1 to 4.

[0072] like Figure 3 The diagram shown (the a-type and b-type actuator control modules are referred to as aREU and bREU, respectively) illustrates the actuator control and energy configuration of this invention. Three flight control computers are configured, employing the same design and ensuring interchangeability. Each computer has a command channel and a monitoring channel to prevent the output of erroneous control commands. The two channels are designed based on different processor models to eliminate common-mode risk. The three flight control computers provide redundancy for each other. Each flight control computer communicates bidirectionally with all actuator control electronics via the SIB bus, and the flight control computers also communicate bidirectionally with each other via the SIB bus, providing redundant cross-data communication functionality.

[0073] like Figure 3 As shown, there are 19 actuators: two actuators for each elevator and aileron, three actuators for each rudder, one actuator for each spoiler, and two actuators for each water rudder. Each actuator is fixedly controlled by an actuator control module, and performs extension / retraction movements under the servo control commands of the actuator control module and the assistance of a hydraulic power source, thereby driving the deflection of the corresponding control surface. Different actuators on the same control surface are redundantly distributed to two or three independent hydraulic power sources to prevent catastrophic events caused by the failure of a single hydraulic power source.

[0074] like Figure 3 As shown, the actuator control module (REU) is configured with 19 units, divided into 13 type A and 6 type B. The two configurations use different processors for dissimilar design. Each actuator control module has an instruction channel and a monitoring channel to prevent the output of erroneous actuator servo control commands. The same model is interchangeable. The type A actuator control module receives control surface commands from the actuator control electronics via the ADB bus, while the type B actuator control module receives control surface commands from the actuator control electronics via the ADB bus and simultaneously receives control surface commands from the backup flight control computer via hardwired connections. The actuator control module converts the received control commands from the actuator control electronics and the backup flight control computer into servo control commands for the actuator according to predetermined logic, thus achieving servo control of the actuator.

[0075] like Figure 3As shown, four angular position sensors are configured and installed on the left / right elevators and left / right ailerons respectively to measure the deflection of the corresponding control surfaces. The data is then transmitted to the corresponding actuator control electronics via the B-type actuator control module. The actuator control electronics then transmit the data to the flight control computer via the SIB bus, and the flight control computer reports it to the indication and recording system for display and recording.

[0076] like Figure 3 As shown, the actuator control electronics, backup flight control computer, actuator control module, and actuator combined with control surface functions are defined as follows: Actuator control electronics 1 controls the left elevator actuator 1, left spoiler actuator 1, right spoiler actuator 1, left aileron actuator 1, rudder actuator 1, and water rudder actuator 1 respectively through 6 aREUs; Actuator control electronics 2 controls the left elevator actuator 2 and right aileron actuator 1 respectively through 2 bREUs, receives the left elevator deflection and right aileron deflection collected by two angular displacement sensors, and controls the rudder actuator 2 through 1 aREU; Actuator control electronics 3 controls the right elevator actuator 2 respectively through 2 bREUs. The rudder actuator 2 and the left aileron actuator 2 receive the right elevator deflection and left aileron deflection collected by two angular displacement sensors, and control the left 2nd spoiler actuator and the right 2nd spoiler actuator through two aREUs; the actuator control electronics 4 controls the right elevator actuator 1, the left 3rd spoiler actuator, the right 3rd spoiler actuator and the right aileron actuator 2 through four aREUs, and controls the rudder actuator 3 and the water rudder actuator 2 through two bREUs; when necessary, the backup flight control computer can reuse six b-type actuator control modules to reconstruct the minimum acceptable control loop of the ultimate backup system, realize three-axis flight control, and ensure the aircraft continues to fly and land safely.

[0077] like Figure 3 As shown, the actuator control electronics 1 to 4 are all powered by the onboard power system with 2 redundancy. Actuator control electronics 1 powers flight control computer 1, actuator control electronics 2 powers flight control computer 2, actuator control electronics 3 powers flight control computer 3, and the backup flight control computer is powered by the onboard power system with 2 redundancy.

[0078] like Figure 3 As shown, the power system includes at least a dual-redundant normal busbar and at least a dual-redundant emergency busbar. Actuator control electronics 1 to 4 are connected to one normal busbar and one emergency busbar respectively according to certain rules. Flight control battery 1 provides backup power to actuator control electronics 1, actuator control electronics 2 and backup flight control computer. Flight control battery 2 provides backup power to actuator control electronics 3 and actuator control electronics 4, ensuring that the flight control system can work normally for at least 20 minutes after the power system fails, preventing catastrophic events caused by single-redundant busbar failure.

[0079] like Figure 3 As shown, the hydraulic power system includes three independent hydraulic power sources. To prevent the failure of a single hydraulic power source from causing the loss of control function of a certain main control surface, multiple actuators of each main control surface are redundantly distributed to different hydraulic power sources to provide hydraulic pressure. The same pair of spoiler actuators are selectively connected to a hydraulic power source.

[0080] like Figure 4 The diagram illustrates the switching of operating modes in the fly-by-wire flight control system of this invention. The control laws for the four operating modes—normal, degraded, direct, and backup—are calculated by flight control computers at different levels, with control priority decreasing progressively and flight quality decreasing accordingly. The normal mode control law and the degraded mode control law are calculated synchronously by the flight control computer, the direct mode control law is calculated by the actuator control electronics, and the backup mode control law is calculated by the backup flight control computer. The normal-degraded mode command switching logic is executed by the flight control computer, the normal-direct mode command switching logic is executed by the actuator control electronics, and the backup command access logic is executed by the bREU.

[0081] like Figure 1 , 2 As shown in Figure 4, when the fly-by-wire flight control system of this invention is in normal operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the control electronics of each actuator via the SIB bus to obtain the voted pilot commands. Combined with information such as inertial navigation, angle of attack, atmospheric conditions, and flap position, the control commands for the normal mode control law are calculated. Normal operating mode is the system's default operating state. In this state, the fly-by-wire flight control system can achieve ideal control functions, including three-axis stability augmentation control, envelope protection, coordinated turn control, and control surface deflection angle limitation.

[0082] like Figure 1 , 2 As shown in Figure 4, when the fly-by-wire flight control system of the present invention is in degraded operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the control electronics of each actuator through the SIB bus, obtains the voted pilot commands, and calculates the control commands of the degraded mode control law by combining the signals of the direct angular rate sensor and the flap position sensor. In this operating mode, the envelope protection function and the control surface limitation function will be lost.

[0083] like Figure 1 , 2 As shown in Figure 4, when the fly-by-wire flight control system of the present invention is in direct operating mode, the actuator control electronics calculate the control command of the direct mode control law based on the normal command signal it receives and the flap extension / retraction state quantity. In this operating mode, the three-axis stabilization control function and coordinated turning function will be further lost.

[0084] like Figures 1 to 4 As shown, when the fly-by-wire flight control system is in backup mode, the backup flight control computer will calculate the control command of the backup mode control law based on the received pilot backup command signal and send it to bREU. In this mode, the control function of all spoilers will be further lost.

[0085] As a preferred option, the three flight control computers rotate in a "master-backup-backup" manner based on the number of power-on cycles, the validity of their own votes, and the integrity of their interfaces. Each computer independently calculates control commands for normal / degraded operating modes and sends the control commands to all actuator control electronics. The actuator control electronics then forward the control commands from the "master" flight control computer to the actuator control module based on the master / backup information provided by the flight control computer.

[0086] Preferably, each actuator on a control surface is capable of driving the control surface to operate normally. The two actuators of the single-piece elevator, single-piece aileron, and water rudder operate in a "primary-backup" mode, while the rudder operates in a "primary-backup-backup" mode. The actuator control modules of the same control surface can communicate with each other, with priority given to the actuator control module corresponding to the "primary" actuator. There is no communication between actuator control modules of different control surfaces.

[0087] Preferably, the same control surface cannot simultaneously execute control commands for two or more operating modes of the fly-by-wire flight control system to avoid control command conflicts and abrupt changes. When a higher-level operating mode is available, the system cannot automatically upgrade to the new mode, but will issue a high-level flight control operating mode availability alarm to the crew, requiring the pilot to manually switch the flight control mode switch to manually upgrade. The instruction recording system issues mode alarm information to the crew when the flight control system switches to different operating modes.

[0088] Preferably, when the flight control computer has sufficient control law calculation capability, but one of the following situations occurs, it will automatically output a degraded control law (normal-degraded mode command switching logic): 1) Atmospheric data is less than two redundancies; or 2) Inertial data is less than two redundancies (including inertial navigation and attitude); or 3) Angle of attack data is less than two redundancies; or 4) ≥2 actuator control electronics switch to direct mode. When one of the following situations occurs, a direct control law (normal-direct mode command switching logic) will be output: 1) All 3 flight control computers fail; or 2) The manual control mode selection switch is set to the "direct" position; or 3) 2 or more actuator control electronics cannot execute the commands of the flight control computer. When the following situation occurs, a backup control law (backup command access logic) will be activated: 1) A type B actuator control module on a certain control surface cannot obtain valid control commands from the actuator control electronics, and another type A actuator control module on that control surface also cannot execute the commands of the actuator control electronics.

[0089] The fly-by-wire flight control system and its operating modes proposed in this invention are the first complete public disclosure in the field, both domestically and internationally, of a fly-by-wire flight control system and its operating modes specifically designed for the special purposes of amphibious aircraft. This fills a gap in the design of fly-by-wire flight control systems for amphibious aircraft operating across different media, and is of pioneering significance. The fly-by-wire flight control system proposed in this invention employs a method of using a backup computer instead of a dissimilar actuator control electronics design to eliminate the risk of system common-mode issues. This avoids safety problems such as compatibility, command integrity, and fault monitoring that may be introduced by dissimilar actuator control electronics designs, leading to incorrect disconnection of actuator control electronics, frequent restarts, and disruptive alarms. Simultaneously, it reuses some existing actuator control equipment from the normal flight control system to construct an additional, single-channel, small fly-by-wire control channel capable of achieving minimum acceptable control as a final backup. This channel has three-axis backup control authority, ensuring the aircraft can continue to fly and land safely, and is lighter than mechanical backup systems. The control law calculation and switching of the four different operating modes designed in this invention are executed by the relevant flight control computer according to the set logic, solving the problem of command disputes during control channel reconfiguration. The fly-by-wire flight control system and operating mode proposed in this invention have many advantages, including high safety, complete functions, refined structure, reasonable layout, simple control mode switching logic, and ease of engineering implementation.

[0090] The above description is only a preferred embodiment of the present invention. It should be noted that those skilled in the art can make several improvements without departing from the principle of the present invention, and these improvements should also be considered within the scope of protection of the present invention.

Claims

1. A fly-by-wire flight control system for an amphibious aircraft, characterized in that, The system includes a cockpit control unit (1), a normal command sensor (2), a backup command sensor (15), an actuator control electronics (13), a flight control computer (9), a flap position sensor (7), a backup flight control computer (3), an actuator control module (6), an actuator (4), an external signal sensor (14), a direct angular rate sensor (11), an angular displacement sensor (5), and a flight control mode switch (12). The cockpit control unit (1) connects to and controls the normal command sensor (2) and the backup command sensor (15). The normal command sensor (2) transmits signals to the actuator control electronics (13). The actuator control electronics (13) and the flight control mode switch (9) control the flight control mode. The flight control computer (9) and the actuator control electronics (13) and the actuator control module (6) exchange signals. The actuator control electronics (13) also receives signals from the external signal sensor (14), the direct angular rate sensor (11), the flap position sensor (7), and the flight control mode switch (12). The backup command sensor (15) transmits signals to the backup flight control computer (3), and the backup flight control computer (3) transmits signals to the actuator control module (6). The actuator control module (6) connects to and controls the actuator (4). The angular displacement sensor (5) is installed on the actuator (4) and transmits the collected signals to the actuator control module (6). Normal command sensor (2) collects command signals from cockpit control devices (1) reflecting the pilot's control intentions and transmits them to actuator control electronics (13); external signal sensor (14) collects dynamic pressure, static pressure, temperature, aircraft attitude and position information and transmits it to actuator control electronics (13); direct angular rate sensor (11) collects aircraft three-axis angular rate information and transmits it to actuator control electronics (13); flap position sensor collects flap position signals and transmits them to actuator control electronics (13); actuator control electronics (13) transmits the received signals to flight control computer (9), calculates the direct control law, receives normal or degraded control commands issued by flight control computer (9), and selects one control command from different working modes to issue to actuator control module (6) according to working mode switching logic; After receiving the control command, the actuator control module (6) converts it into a servo control command for the actuator (4) according to a predetermined logic; the actuator (4) receives the servo control command from the actuator control module (6) and performs telescopic movement under the pressure of the hydraulic source, causing the control surface to deflect; the angular displacement sensor (5) measures the deflection of the corresponding control surface and feeds it back to the actuator control module (6), which then uploads it to the flight control computer (9) via the actuator control electronics (13); The actuator control module (6) adopts a command and monitoring architecture, including a type A module and a type B module. The type A module only accepts control surface commands from the actuator control electronics (13), while the type B module can accept control surface commands from the actuator control electronics (13) and the backup flight control computer (3).

2. The fly-by-wire flight control system for an amphibious aircraft according to claim 1, characterized in that, There are 6 normal command sensors (2), all with 4 redundancy, installed under the left and right control devices respectively; there are 3 backup command sensors (15), installed under the left control device; the external signal sensors (14) include a 3-redundant atmospheric data reference unit, a 1-redundant attitude reference unit, a 2-redundant inertial reference unit, a 2-redundant radio altimeter and a 4-redundant angle of attack sensor; there are 4 direct angular rate sensors (11), installed symmetrically in the middle of the fuselage, which output 4-redundant aircraft three-axis angular rate data for degraded control law calculation; there are 4 actuator control electronics (13), which adopt the same design and are interchangeable. Each actuator control electronics (13) is equipped with a command channel and a monitoring channel, and there is no signal interaction between them; there are 3 flight control computers (9), which adopt the same design and are interchangeable. Each flight control computer (9) is equipped with a command channel and a monitoring channel, and there is signal interaction between them; there is 1 backup flight control computer (3), which is equipped with a command channel and a monitoring channel.

3. A method for operating a fly-by-wire flight control system for an amphibious aircraft, using the fly-by-wire flight control system for an amphibious aircraft as described in any one of claims 1-2, characterized in that, There are four working modes: normal working mode, degraded working mode, direct working mode, and backup working mode. The priority of the four working modes is as follows: normal working mode is the first priority, degraded working mode is the second priority, direct working mode is the third priority, and backup working mode is the fourth priority.

4. The method for operating a fly-by-wire flight control system for an amphibious aircraft according to claim 3, characterized in that, In normal operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the actuator control electronics via the SIB bus, obtains the voted pilot commands, and calculates the control commands of the normal mode control law by combining information on inertial navigation, angle of attack, atmosphere, and flap position. Normal operating mode is the system's default operating state. At this time, the fly-by-wire flight control system can achieve ideal control functions, including three-axis stabilization control, envelope protection, coordinated turn function, and control surface deflection angle limitation function.

5. The method for operating a fly-by-wire flight control system for an amphibious aircraft according to claim 3, characterized in that, In degraded operating mode, the flight control computer performs redundancy voting on the pilot control commands uploaded by the actuator control electronics via the SIB bus, obtains the voted pilot commands, and calculates the control commands of the degraded mode control law by combining the signals from the direct angular rate sensor and the flap position sensor, and loses the envelope protection function and control surface limitation function.

6. The method for operating a fly-by-wire flight control system for an amphibious aircraft according to claim 3, characterized in that, In direct operating mode, the actuator control electronics calculate the control command of the direct mode control law based on the normal command signals it receives and the flap extension / retraction state, and lose the three-axis stabilization control function and coordinated turning function.

7. The method for operating a fly-by-wire flight control system for an amphibious aircraft according to claim 3, characterized in that, In backup mode, the backup flight control computer will calculate the control commands of the backup mode control law based on the received pilot backup command signal and send them to the actuator control module, thus losing the control function of all spoilers.

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