An aircraft fault reconfiguration and fault-tolerant control method, device, equipment, medium and product

Abstract

The application discloses an aircraft fault reconstruction and fault-tolerant control method, device, equipment, medium and product, and relates to the technical field of aircraft control and safety. The method comprises the following steps: acquiring the running state information of each rotor; determining complete failure faults based on the running state information of each rotor to obtain a determination result; if the determination result is that a complete failure fault occurs, locating the physical position of the fault rotor to obtain fault location information; based on the fault location information, performing reconstruction and execution processing of a target configuration according to preset decision logic to obtain a reconstructed configuration; and performing full-attitude control on the reconstructed configuration by using a nonlinear control law to stabilize the attitude of the aircraft and realize safe landing. After detecting a single-rotor fault, the aircraft configuration can be actively reconstructed and switched to the corresponding nonlinear control law, so that rapid attitude stabilization and safe landing after the fault are realized, and the survival ability of the aircraft is greatly improved.

Description

Technical Field

[0001] This application relates to the field of aircraft control and safety technology, and in particular to an aircraft fault reconfiguration and fault-tolerant control method, apparatus, equipment, medium and product. Background Technology

[0002] The flight safety of multi-rotor aircraft is highly dependent on the redundancy of their power systems. For configurations such as hexa-rotor or octo-rotor, if a single rotor fails, thrust and torque can be redistributed to a certain extent by adjusting the speed or pitch of the remaining normal rotors, thus maintaining the basic stability of the aircraft. This fault-tolerant control strategy based on hardware redundancy is relatively mature.

[0003] However, for a quadcopter, the rotors are symmetrically distributed in a cruciform shape, resulting in limited redundancy. If one rotor experiences a complete failure, the system will be forced to degenerate into an asymmetric tri-rotor mode. This mode presents a series of serious challenges: (1) Breaking of dynamic symmetry: The original symmetrical thrust pattern is broken, generating an inherent unbalanced torque that cannot be offset by simple speed adjustment, especially a strong yaw torque, which leads to a sharp loss of attitude of the aircraft.

[0004] (2) Increased nonlinearity and strong coupling effects: The dynamic model of the system after the fault exhibits significant nonlinear characteristics, and the coupling between pitch, roll and yaw channels is sharply enhanced. Traditional linear controllers (such as PID controllers) rely on linearized models and fixed parameters, making it difficult to adapt to such dynamic changes. The control performance drops sharply, and instability may even be exacerbated by control command mismatch.

[0005] (3) Insufficient control efficiency: In the tri-rotor mode, the control authority of the remaining actuators is greatly reduced, making it difficult to effectively counteract the continuous interference introduced by the fault.

[0006] Therefore, current strategies based on hardware redundancy and linear control are difficult to directly apply to single-rotor fatal failure scenarios in quad tiltrotor aircraft. There is an urgent need for a novel fault-tolerant control scheme that can fundamentally reconstruct the force and torque generation mechanism from a configurational perspective. Summary of the Invention

[0007] The purpose of this application is to provide a method, device, equipment, medium, and product for aircraft fault reconfiguration and fault-tolerant control, which can actively reconfigure the aircraft configuration and switch to the corresponding nonlinear control law after detecting a single rotor fault, thereby achieving rapid attitude stabilization and safe landing after the fault, and greatly improving the survivability of the aircraft.

[0008] To achieve the above objectives, this application provides the following solution: Firstly, this application provides a method for aircraft fault reconstruction and fault-tolerant control, which is applied to a scenario in a four-tilt rotor helicopter mode where a single rotor completely fails; the method includes: The system acquires the operating status information of each rotor; the operating status information is based on multi-source sensors integrated on each rotor of the four tiltrotor aircraft; the operating status information includes: rotational speed, motor current, drive torque, and fuselage vibration acceleration; Based on the operating status information of each rotor, a complete failure fault is determined, and the determination result is obtained; If the determination result is a complete failure, then locate the physical location of the faulty rotor to obtain the fault location information; Based on the fault location information, the target configuration is reconstructed and processed according to the preset decision logic to obtain the reconstructed configuration. A nonlinear control law is used to perform full attitude control on the reconstructed configuration in order to stabilize the aircraft's attitude and achieve a safe landing.

[0009] In one embodiment, a complete failure fault is determined based on the operating status information of each rotor, and a determination result is obtained, specifically including: The speed, motor current, and fuselage vibration acceleration in the operating status information of each rotor are compared using a threshold comparison method to obtain the judgment result. The driving torque in the operating status information of each rotor is compared using a state estimation method to obtain a judgment result; if the deviation obtained from the comparison continues to exceed the preset range, the judgment result is that a complete failure has occurred.

[0010] In one embodiment, the preset decision logic specifically includes: The four rotors are divided into two diagonal combinations. When a fault is detected in one of the rotors, the other rotor in the same combination is actively controlled to stop, in order to eliminate the inherent unbalanced torque caused by the asymmetrical structure. The remaining two rotors in another combination were identified as the reconfigured power unit.

[0011] In one embodiment, the target configuration adopts a tandem twin-rotor configuration.

[0012] In one embodiment, based on the fault location information, the target configuration is reconstructed and processed according to a preset decision logic to obtain the reconstructed configuration, specifically including: By using electrical signals or mechanical actuators, the power supply to the faulty rotor and the stopped rotor is cut off, and all power resources are concentrated on the reconfigured power unit. The aircraft's body coordinate system is reconstructed, and the control relationship mapping is reconstructed. The reconstruction of the body coordinate system redefines the heading reference as the direction from the front rotor to the rear rotor. The reconstruction of the control relationship mapping includes: mapping the collective pitch of the control input to the same direction of the collective pitch of the two rotors to control altitude / lift; mapping the longitudinal periodic pitch of the control input to the same direction of the longitudinal periodic pitch of the two rotors to generate pitch moment about the Y-axis; and mapping the lateral periodic pitch of the control input to the same direction of the lateral periodic pitch of the two rotors to generate roll moment about the X-axis.

[0013] In one embodiment, the nonlinear control law includes: altitude and collective pitch control, pitch control, roll control, and yaw control; The altitude and collective pitch control is achieved by synchronously adjusting the collective pitch of the two working rotors, combined with the aircraft's descent speed, so that the working rotors maintain their rotational speed within a preset acceptable range, while the lift generated by the working rotors overcomes gravity. The pitch control is achieved by differentially changing the collective pitch or speed of the front and rear rotors, or by direct longitudinal cyclic pitch control, to generate a pitch control torque around the lateral axis. The roll control is achieved by synchronously and periodically varying the lateral pitch of the front and rear rotors to generate a roll torque around the longitudinal axis. The yaw control is achieved by differentially adjusting the counter-torque generated by the front and rear rotors to control the yaw torque.

[0014] Secondly, this application provides an aircraft fault reconfiguration and fault-tolerant control device, comprising: The information acquisition module is used to acquire the operating status information of each rotor; the operating status information is based on the multi-source sensors integrated on each rotor of the four tiltrotor aircraft; the operating status information includes: rotational speed, motor current, drive torque and fuselage vibration acceleration; The determination module is used to determine the complete failure fault based on the operating status information of each rotor and obtain the determination result; The fault location module is used to locate the physical location of the faulty rotor and obtain fault location information when the determination result is a complete failure fault. The reconstructed execution module is used to reconstruct and execute the target configuration based on the fault location information and according to the preset decision logic to obtain the reconstructed configuration. The attitude control module is used to perform full attitude control on the reconstructed configuration using nonlinear control laws to stabilize the aircraft's attitude and achieve a safe landing.

[0015] Thirdly, this application provides a computer device, including: a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the above-described aircraft fault reconfiguration and fault-tolerant control method.

[0016] Fourthly, this application provides a computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the aforementioned aircraft fault reconfiguration and fault-tolerant control method.

[0017] Fifthly, this application provides a computer program product, including a computer program that, when executed by a processor, implements the aforementioned aircraft fault reconfiguration and fault-tolerant control method.

[0018] According to the specific embodiments provided in this application, the following technical effects are disclosed: This application provides a method, apparatus, device, medium, and product for aircraft fault reconstruction and fault-tolerant control. It acquires the operating status information of each rotor; determines a complete failure based on this information; if the determination is a complete failure, it locates the physical position of the faulty rotor, obtaining fault location information; based on this location information, it reconstructs and executes the target configuration according to a preset decision logic, obtaining the reconstructed configuration; and uses a nonlinear control law to perform full-attitude control on the reconstructed configuration to stabilize the aircraft's attitude and achieve a safe landing. This application can proactively reconstruct the aircraft configuration and switch to the corresponding nonlinear control law after detecting a single rotor fault, thereby achieving rapid attitude stabilization and safe landing after a fault, greatly improving the aircraft's survivability. Attached Figure Description

[0019] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0020] Figure 1 This is a flowchart of the aircraft fault reconstruction and fault-tolerant control method.

[0021] Figure 2 This is the overall control logic flowchart.

[0022] Figure 3 This is a schematic diagram of the reconstruction of the aircraft's coordinate system after a malfunction.

[0023] Figure 4This is a schematic diagram of the manipulation relationship mapping under the tandem configuration.

[0024] Figure 5 This is a structural diagram of the aircraft fault reconstruction and fault-tolerant control device.

[0025] Figure 6 This is a schematic diagram of the structure of a computer device provided in an embodiment of this application. Detailed Implementation

[0026] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0027] The proposed solution is to actively reconfigure the aircraft configuration and control strategy when a single rotor of a quad tiltrotor aircraft experiences a complete failure in helicopter mode (vertical takeoff, hovering, low-speed forward flight), thereby stabilizing the aircraft and enabling a safe landing by converting it into a tandem dual-rotor mode.

[0028] To make the above-mentioned objectives, features and advantages of this application more apparent and understandable, the application will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0029] In one exemplary embodiment, an aircraft fault reconfiguration and fault-tolerant control method is provided, which is applied to a scenario in which a single rotor completely fails in a four-tilt rotor aircraft helicopter mode.

[0030] like Figure 1 As shown, the aircraft fault reconstruction and fault-tolerant control method includes: Step 100: Obtain the operating status information of each rotor. The operating status information is obtained based on multi-source sensors integrated on each rotor of the four tiltrotor aircraft; the operating status information includes: rotational speed, motor current, drive torque, and fuselage vibration acceleration.

[0031] Step 200: Determine the complete failure fault based on the operating status information of each rotor, and obtain the determination result.

[0032] Among them, a complete failure fault is determined based on the operating status information of each rotor, and the determination result is obtained, specifically including: The speed, motor current, and fuselage vibration acceleration of each rotor are compared using a threshold comparison method to obtain the judgment result.

[0033] The driving torque in the operating status information of each rotor is compared using a state estimation method to obtain a judgment result; if the deviation obtained from the comparison continues to exceed the preset range, the judgment result is that a complete failure has occurred.

[0034] Step 300: If the determination result is a complete failure, locate the physical location of the faulty rotor and obtain the fault location information.

[0035] Step 400: Based on the fault location information, the target configuration is reconstructed and processed according to the preset decision logic to obtain the reconstructed configuration. The target configuration adopts a tandem twin-rotor configuration.

[0036] The pre-defined decision-making logic specifically includes: The four rotors are divided into two diagonal combinations. When a fault is detected in one rotor, the other rotor in the same combination is actively controlled to stop, in order to eliminate the inherent unbalanced torque caused by the asymmetrical structure.

[0037] The remaining two rotors in another combination were identified as the reconfigured power unit.

[0038] In one embodiment, based on fault location information, the target configuration is reconstructed and processed according to preset decision logic to obtain the reconstructed configuration, specifically including: By using electrical signals or mechanical actuators, the power supply to the faulty rotor and the stopped rotor is cut off, and all power resources are concentrated on the reconfigured power unit.

[0039] The aircraft's body coordinate system is reconstructed, and the control relationship mapping is reconstructed. The reconstruction of the body coordinate system redefines the heading reference as the direction from the front rotor to the rear rotor. The reconstruction of the control relationship mapping includes: mapping the collective pitch of the control input to the same direction of the collective pitch of the two rotors to control altitude / lift; mapping the longitudinal periodic pitch of the control input to the same direction of the longitudinal periodic pitch of the two rotors to generate pitch moment about the Y-axis; and mapping the lateral periodic pitch of the control input to the same direction of the lateral periodic pitch of the two rotors to generate roll moment about the X-axis.

[0040] Step 500: Use a nonlinear control law to perform full attitude control on the reconstructed configuration to stabilize the aircraft's attitude and achieve a safe landing.

[0041] Nonlinear control laws include: altitude and collective pitch control, pitch control, roll control, and yaw control.

[0042] Among them, altitude and collective pitch control is achieved by synchronously adjusting the collective pitch of the two working rotors, combined with the aircraft's descent speed, so that the working rotors maintain their rotational speed within a preset acceptable range, while the lift generated by the working rotors overcomes gravity.

[0043] Pitch control generates a pitch control torque around the lateral axis by differentially changing the collective pitch or speed of the front and rear rotors, or by direct longitudinal cyclic pitch control.

[0044] Roll control generates a roll moment about the longitudinal axis by synchronously and periodically varying the lateral pitch of the front and rear rotors. Yaw control controls the yaw moment by differentially adjusting the counter-torque generated by the front and rear rotors.

[0045] The purpose of this application is to overcome the shortcomings of the current technology. The method mentioned in this application does not rely on additional hardware redundancy. Instead, after detecting a single rotor failure, it actively reconstructs the aircraft configuration into an aerodynamically stable tandem helicopter mode and switches to a nonlinear control law designed specifically for this configuration, thereby achieving rapid attitude stabilization and safe landing after the failure, greatly improving the survivability of the aircraft.

[0046] like Figure 2 As shown, the operation process of the method mentioned in this application is as follows: Fault detection and diagnosis: The operating status of each rotor is monitored in real time by integrating multiple source sensors (including but not limited to rotor speed sensors, motor current / torque sensors, and fuselage vibration acceleration sensors) on each rotor subsystem. Using threshold-based or model-based state estimation algorithms, when a rotor's speed abnormally drops below a safe threshold, or the drive torque changes significantly or disappears completely, the rotor is immediately determined to have suffered a complete failure, and the physical location of the faulty rotor (e.g., front left, front right, rear left, rear right) is precisely located.

[0047] Threshold-based methods directly compare real-time sensor data (such as rotational speed and current) with preset safety thresholds. For example, a fault can be identified when a rotor's rotational speed falls below the minimum threshold of its normal operating range, or when the motor current approaches zero (indicating no torque output). This is a direct data comparison.

[0048] Model-based methods: A mathematical model of the rotor or motor system is established to predict its normal output (e.g., predicted torque) under the current command. The actual sensor measurements are then compared with the model predictions. If the deviation consistently exceeds a certain range, it indicates abnormal system behavior and a potential malfunction. This is a more advanced method based on comparing "expected" and "actual" data. The mathematical model of the rotor or motor system is a mechanistic model based on actual rotor or motor parameters. In one embodiment, a currently known mechanistic model can be used.

[0049] Reconfigurable configuration decision: Based on fault location information, this system automatically selects the optimal reconfiguration target configuration—a tandem twin-rotor configuration—according to preset decision logic. The decision logic is as follows: The four rotors are divided into two diagonal combinations: Group 1 (left front rotor and right rear rotor) and Group 2 (right front rotor and left rear rotor).

[0050] When a rotor malfunction is detected, the control system immediately issues a command to actively stop the other healthy rotor in the same group. This is intended to completely eliminate the inherent unbalanced torque caused by the asymmetrical structure.

[0051] The remaining two healthy rotors in another group were identified as the reconfigured power unit. The rotor located at the front of the aircraft was defined as the front rotor, and the rotor located at the rear was defined as the rear rotor, together forming a tandem configuration.

[0052] Example: If the left front rotor fails, the decision system commands the right rear rotor to stop. At this time, the right front rotor (front rotor) and the left rear rotor (rear rotor) form a new power system.

[0053] Configuration refactoring execution: This step is responsible for physically implementing the instructions of the decision-making module, including: Power system reconfiguration: By using electrical signals or mechanical actuators, the power supply to the faulty rotor and the rotor that has been ordered to stop is cut off (such as motor power failure or clutch disengagement), and all power resources are concentrated on the remaining two working rotors.

[0054] Flight control system reconfiguration: Coordinate system redefinition: such as Figure 3 As shown, the aircraft's coordinate system is reconstructed, and in particular, the heading reference (longitudinal axis) is redefined as the direction from the front rotor to the rear rotor.

[0055] Manipulating relational mapping reconstruction: such as Figure 4As shown, the mapping relationship between the control inputs (such as collective pitch, longitudinal cyclic pitch, and lateral cyclic pitch) and the two working rotor actuators is redefined to conform to the control characteristics of a tandem helicopter. Assume that the longitudinal / lateral cyclic pitch before the failure is defined as... Figure 3 As shown on the left, rotating the propeller disk plane around the X-axis results in longitudinal pitch variation, and rotating it around the Y-axis results in lateral pitch variation. The reconfigured result is shown on the right, where rotating the propeller disk plane around the X-axis results in longitudinal pitch variation, and rotating it around the Y-axis results in lateral pitch variation.

[0056] Coordinate system transformation and target redefinition: such as Figure 3 As shown, the longitudinal axis (X-axis) of the reconstructed airframe is redefined as "the direction from the front rotor to the rear rotor". Therefore, torque control around the new lateral axis (Y-axis) is defined as pitch control, and torque control around the new longitudinal axis (X-axis) is defined as roll control. This is the physical basis of the mapping relationship.

[0057] Matching the handling characteristics of tandem helicopters: The specific mathematical form of the mapping relationship (i.e., how the control input is calculated as actuator commands for the two rotors) needs to be designed based on the aerodynamic characteristics of the tandem helicopter. This typically refers to simplified formulas or coefficients derived from classical helicopter flight dynamics theories (such as blade element theory and momentum theory), describing the relationship between rotor collective pitch, cyclic pitch, and the generated forces / torques; these will not be elaborated upon here. During control law design, these formulas or models obtained through system identification will be used to establish the following mapping: Collective pitch input is primarily mapped to the same-direction variation of collective pitch between the two rotors to control altitude / lift.

[0058] The longitudinal periodic pitch input is mainly mapped to the same direction of the longitudinal periodic pitch of the two rotors to generate a pitching moment about the Y-axis.

[0059] The lateral periodic pitch input is mainly mapped to the same-direction change of the lateral periodic pitch of the two rotors to generate a rolling torque about the X-axis.

[0060] To achieve yaw control, differential adjustment of the torque of the front and rear rotors is required, which is usually achieved by introducing differential components into the collective pitch or lateral cyclic pitch.

[0061] Columnar configuration-specific control law: For the reconstructed tandem configuration, a pre-designed and stored nonlinear control law is enabled.

[0062] The core of adjusting the torque of the front and rear rotors lies in using the differential change of the counter-torque of the tandem front and rear rotors to generate yaw torque.

[0063] While each rotor generates lift during rotation, it is also subjected to a counter-torque from air resistance, which acts in the opposite direction to the rotor's rotation. In a tandem configuration, the two rotors are typically configured to rotate in opposite directions to counteract most of this counter-torque. Yaw control is achieved by intentionally disrupting the balance of the counter-torques between the two rotors.

[0064] Method 1: Differential Collective Pitch (Differential Torque). This is the most direct method. The rotor's counter-torque is roughly proportional to the square of its collective pitch (or rotational speed). By slightly increasing the collective pitch of the front rotor and slightly decreasing the collective pitch of the rear rotor (and vice versa), a difference in the magnitude of the counter-torque generated by the front and rear rotors can be created. This difference in counter-torque cannot be completely canceled out by the airframe structure, thus forming a net torque that causes the aircraft to rotate around its yaw axis.

[0065] Method 2: Differential Lateral Cyclic Pitch. The lateral cyclic pitch of the rotor changes the lateral tilt angle of the rotor disk plane, thus generating lateral forces. If the front rotor disk is tilted to the left while the rear rotor disk is tilted to the right, these two lateral forces will form a couple, directly generating a yaw moment. This method may have a faster response, but the control coupling is more complex.

[0066] In actual control law design, these two methods may be used in combination. The control algorithm calculates the required combination of front and rear rotor collective pitch and cyclic pitch commands in real time according to the current flight state to achieve coordinated control of the four channels of altitude, pitch, roll and yaw.

[0067] In other words, the control law fully utilizes the remaining two control surfaces (rotors) to achieve full attitude control: Altitude and collective pitch control: By synchronously adjusting the collective pitch of the two working rotors, combined with the aircraft's descent speed, the working rotors are kept within an acceptable rotational speed range, while the lift generated by the rotors can overcome gravity.

[0068] Pitch control: Pitch control torque is generated around the lateral axis by differentially changing the collective pitch (or speed) of the front and rear rotors, or by direct longitudinal cyclic pitch control. For example, increasing the collective pitch of the front rotor while decreasing the collective pitch of the rear rotor will generate a nose-down torque.

[0069] Roll control: Primarily achieved through direct lateral cyclic pitch control. Synchronous lateral cyclic pitch control of the front and rear rotors generates a roll moment around the longitudinal axis. For example, tilting both rotor disks to the left simultaneously generates a left roll moment.

[0070] Yaw control: This is achieved by differentially adjusting the counter-torque generated by the front and rear rotors. This can be achieved by fine-tuning the collective pitch of the two rotors (differential torque variation) or differential lateral periodic pitch variation to achieve precise control of the yaw torque.

[0071] Control command generation and execution: The Flight Control System (FCS), based on the deviation between the desired emergency landing trajectory and the current flight state (attitude, position, etc.), and in conjunction with the control law corresponding to the tandem configuration, calculates in real time the collective pitch, longitudinal cyclic pitch, and lateral cyclic pitch commands for the two working rotors, and drives the corresponding servo mechanisms to execute them, thereby stabilizing the aircraft's attitude and guiding it to a safe area. In the actual control law design, the control algorithm calculates the required combination of collective pitch and cyclic pitch commands for the front and rear rotors in real time based on the current flight state, to achieve coordinated control of the four channels: altitude, pitch, roll, and yaw.

[0072] Safe landing guidance: The ultimate goal of the control system is to achieve a safe landing. Throughout the reconfiguration and control process, the system prioritizes a stable vertical descent or a slow self-descent, continuously assesses landing site conditions, and ultimately guides the aircraft to a smooth touchdown.

[0073] The benefits of this application are: (1) Fundamental improvement in fault tolerance: This application breaks through the traditional approach of local adjustment and reconstructs the aircraft from an unstable tri-rotor mode to a tandem dual-rotor mode with mature and stable aerodynamic characteristics by actively stopping one healthy rotor. This is a selective reconstruction strategy that eliminates the instability factors caused by asymmetry from the mechanical root, and greatly improves the probability of survival under extreme failures.

[0074] (2) Optimization of control efficiency and stability: The control law designed specifically for the reconfiguration configuration fully explores and optimizes the control potential of the remaining two rotors, achieving efficient distribution of control energy. Compared to struggling to maintain stability in the three-rotor mode, the control method provided in this application responds more quickly, recovers attitude more smoothly, and is more robust.

[0075] (3) High adaptability and flexibility: The decision logic can be extended to handle multiple fault modes (such as single failures in different locations). The system can automatically select the optimal reconstruction path according to the specific fault location, demonstrating good versatility.

[0076] (4) Fully automated emergency response: It realizes full-process automation from fault detection, configuration decision, system reconfiguration to control law switching, without human intervention, which saves valuable time for emergency response and effectively reduces the risk of accidents. It is especially suitable for unmanned aerial vehicles.

[0077] In one exemplary embodiment, such as Figure 5 As shown, an aircraft fault reconfiguration and fault-tolerant control device is provided, comprising: The information acquisition module is used to acquire the operating status information of each rotor; the operating status information is based on the multi-source sensors integrated on each rotor of the four tiltrotor aircraft; the operating status information includes: rotational speed, motor current, drive torque and fuselage vibration acceleration.

[0078] The determination module is used to determine the complete failure fault based on the operating status information of each rotor and obtain the determination result.

[0079] The fault location module is used to locate the physical location of the faulty rotor and obtain fault location information when the determination result is a complete failure fault.

[0080] The reconstructed execution module is used to reconstruct and execute the target configuration based on the fault location information and according to the preset decision logic to obtain the reconstructed configuration.

[0081] The attitude control module is used to perform full attitude control on the reconstructed configuration using nonlinear control laws to stabilize the aircraft's attitude and achieve a safe landing.

[0082] In one exemplary embodiment, a computer device is provided, which may be a server or a terminal, and its internal structure diagram may be as follows. Figure 6 As shown, this computer device includes a processor, memory, input / output (I / O) interfaces, and a communication interface. The processor, memory, and I / O interfaces are connected via a system bus, and the communication interface is also connected to the system bus via the I / O interfaces. The processor provides computational and control capabilities. The memory includes non-volatile storage media and internal memory. The non-volatile storage media stores the operating system, computer programs, and a database. The internal memory provides the environment for the operating system and computer programs stored in the non-volatile storage media. The database stores aircraft fault reconfiguration and fault-tolerant control data. The I / O interfaces are used for information exchange between the processor and external devices. The communication interface is used for communication with external terminals via a network connection. When the computer program is executed by the processor, it implements aircraft fault reconfiguration and fault-tolerant control methods.

[0083] Those skilled in the art will understand that Figure 6 The structure shown is merely a block diagram of a portion of the structure related to the present application and does not constitute a limitation on the computer device to which the present application is applied. Specific computer devices may include more or fewer components than those shown in the figure, or combine certain components, or have different component arrangements.

[0084] In one exemplary embodiment, a computer device is also provided, including a memory and a processor, wherein the memory stores a computer program, and the processor executes the computer program to implement the steps in the above-described method embodiments.

[0085] In one exemplary embodiment, a computer-readable storage medium is provided storing a computer program that, when executed by a processor, implements the steps in the above-described method embodiments.

[0086] In one exemplary embodiment, a computer program product is provided, including a computer program that, when executed by a processor, implements the steps in the above-described method embodiments.

[0087] It should be noted that the user information (including but not limited to user device information, user personal information, etc.) and data (including but not limited to data used for analysis, data stored, data displayed, etc.) involved in this application are all information and data authorized by the user or fully authorized by all parties, and the collection, use and processing of the relevant data must comply with relevant regulations.

[0088] Those skilled in the art will understand that all or part of the processes in the methods of the above embodiments can be implemented by a computer program instructing related hardware. The computer program can be stored in a non-volatile computer-readable storage medium, and when executed, it can include the processes of the embodiments of the above methods. Any references to memory, databases, or other media used in the embodiments provided in this application can include at least one of non-volatile and volatile memory. Non-volatile memory can include read-only memory (ROM), magnetic tape, floppy disk, flash memory, optical memory, high-density embedded non-volatile memory, resistive random access memory (ReRAM), magnetic random access memory (MRAM), ferroelectric random access memory (FRAM), phase change memory (PCM), graphene memory, etc. Volatile memory can include random access memory (RAM) or external cache memory, etc. By way of illustration and not limitation, RAM can take many forms, such as Static Random Access Memory (SRAM) or Dynamic Random Access Memory (DRAM).

[0089] The databases involved in the embodiments provided in this application may include at least one type of relational database and non-relational database. Non-relational databases may include, but are not limited to, blockchain-based distributed databases. The processors involved in the embodiments provided in this application may be general-purpose processors, central processing units, graphics processing units, digital signal processors, programmable logic devices, data processing logic devices, etc., and are not limited to these.

[0090] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0091] This document uses specific examples to illustrate the principles and implementation methods of this application. The descriptions of the above embodiments are only for the purpose of helping to understand the methods and core ideas of this application. Furthermore, those skilled in the art will recognize that, based on the ideas of this application, there will be changes in the specific implementation methods and application scope. Therefore, the content of this specification should not be construed as a limitation of this application.

Claims

1. A method for fault reconstruction and fault-tolerant control of an aircraft, characterized in that, The aircraft fault reconstruction and fault-tolerant control method is applied to a scenario in which a single rotor completely fails in the helicopter mode of a four-tilt rotor aircraft. The aircraft fault reconstruction and fault-tolerant control method includes: The system acquires the operating status information of each rotor; the operating status information is based on multi-source sensors integrated on each rotor of the four tiltrotor aircraft; the operating status information includes: rotational speed, motor current, drive torque, and fuselage vibration acceleration; Based on the operating status information of each rotor, a complete failure fault is determined, and the determination result is obtained; If the determination result is a complete failure, then locate the physical location of the faulty rotor to obtain the fault location information; Based on the fault location information, the target configuration is reconstructed and processed according to the preset decision logic to obtain the reconstructed configuration. A nonlinear control law is used to perform full attitude control on the reconstructed configuration in order to stabilize the aircraft's attitude and achieve a safe landing.

2. The aircraft fault reconstruction and fault-tolerant control method according to claim 1, characterized in that, Based on the operating status information of each rotor, a complete failure fault is determined, and the determination result is obtained, specifically including: The speed, motor current, and fuselage vibration acceleration in the operating status information of each rotor are compared using a threshold comparison method to obtain the judgment result. The driving torque in the operating status information of each rotor is compared using a state estimation method to obtain a judgment result; if the deviation obtained from the comparison continues to exceed the preset range, the judgment result is that a complete failure has occurred.

3. The aircraft fault reconstruction and fault-tolerant control method according to claim 1, characterized in that, The pre-defined decision-making logic specifically includes: The four rotors are divided into two diagonal combinations. When a fault is detected in one of the rotors, the other rotor in the same combination is actively controlled to stop, in order to eliminate the inherent unbalanced torque caused by the asymmetrical structure. The remaining two rotors in another combination were identified as the reconfigured power unit.

4. The aircraft fault reconstruction and fault-tolerant control method according to claim 1, characterized in that, The target configuration adopts a tandem twin-rotor configuration.

5. The aircraft fault reconstruction and fault-tolerant control method according to claim 3, characterized in that, Based on the fault location information, the target configuration is reconstructed and processed according to preset decision logic to obtain the reconstructed configuration, specifically including: By using electrical signals or mechanical actuators, the power supply to the faulty rotor and the stopped rotor is cut off, and all power resources are concentrated on the reconfigured power unit. The aircraft's body coordinate system is reconstructed, and the control relationship mapping is reconstructed. The reconstruction of the body coordinate system redefines the heading reference as the direction from the front rotor to the rear rotor. The reconstruction of the control relationship mapping includes: mapping the collective pitch of the control input to the same direction of the collective pitch of the two rotors to control altitude / lift; mapping the longitudinal periodic pitch of the control input to the same direction of the longitudinal periodic pitch of the two rotors to generate pitch moment about the Y-axis; and mapping the lateral periodic pitch of the control input to the same direction of the lateral periodic pitch of the two rotors to generate roll moment about the X-axis.

6. The aircraft fault reconstruction and fault-tolerant control method according to claim 1, characterized in that, The nonlinear control laws include: altitude and collective pitch control, pitch control, roll control, and yaw control; The altitude and collective pitch control is achieved by synchronously adjusting the collective pitch of the two working rotors, combined with the aircraft's descent speed, so that the working rotors maintain their rotational speed within a preset acceptable range, while the lift generated by the working rotors overcomes gravity. The pitch control is achieved by differentially changing the collective pitch or speed of the front and rear rotors, or by direct longitudinal cyclic pitch control, to generate a pitch control torque around the lateral axis. The roll control is achieved by synchronously and periodically varying the lateral pitch of the front and rear rotors to generate a roll torque around the longitudinal axis. The yaw control is achieved by differentially adjusting the counter-torque generated by the front and rear rotors to control the yaw torque.

7. An aircraft fault reconstruction and fault-tolerant control device, characterized in that, include: The information acquisition module is used to acquire the operating status information of each rotor. The operational status information is obtained based on multi-source sensors integrated on each rotor of the four tiltrotor aircraft. The operating status information includes: rotational speed, motor current, drive torque, and chassis vibration acceleration; The determination module is used to determine the complete failure fault based on the operating status information of each rotor and obtain the determination result; The fault location module is used to locate the physical location of the faulty rotor and obtain fault location information when the determination result is a complete failure fault. The reconstructed execution module is used to reconstruct and execute the target configuration based on the fault location information and according to the preset decision logic to obtain the reconstructed configuration. The attitude control module is used to perform full attitude control on the reconstructed configuration using nonlinear control laws to stabilize the aircraft's attitude and achieve a safe landing.

8. A computer device, comprising: A memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that the processor executes the computer program to implement the aircraft fault reconfiguration and fault-tolerant control method according to any one of claims 1-6.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the computer program is executed by the processor, it implements the aircraft fault reconfiguration and fault-tolerant control method as described in any one of claims 1-6.

10. A computer program product, comprising a computer program, characterized in that, When the computer program is executed by the processor, it implements the aircraft fault reconfiguration and fault-tolerant control method as described in any one of claims 1-6.

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