Method for morphing aircraft implicit fault tolerant control

Abstract

The application discloses a morphing aircraft implicit fault-tolerant control method. The method specifically comprises the following steps: constructing an attitude dynamics model of the morphing aircraft; constructing a bias and efficiency degradation fault model of the morphing aircraft; constructing an attitude decoupling model of the morphing aircraft; designing a linear extended state observation model; and determining equivalent control instructions of the morphing aircraft according to the attitude decoupling model and the linear extended state observation model. The application avoids the dependence on an explicit fault detection and isolation module by constructing the attitude dynamics model, the bias and efficiency degradation fault model and the attitude decoupling model, and maps the actuator fault and the aerodynamic uncertainty into total disturbances for real-time estimation and feedforward compensation, thereby significantly reducing the structural and computational complexity of the system. Meanwhile, the application can still generate smooth and continuous actual rudder deflection instructions in the case of serious efficiency degradation and bias fault of the actuator, thereby guaranteeing the robustness and stability of the closed-loop system as a whole.

Description

Technical Field

[0001] This application relates to an implicit fault-tolerant control method, electronic equipment, and computer-readable storage medium for deformable aircraft, belonging to the field of aircraft control and fault-tolerant technology. Background Technology

[0002] Deformable aircraft are currently attracting significant attention due to their ability to maintain superior flight performance across a wide flight envelope. By actively adjusting their structural configuration, these aircraft can alter their aerodynamic characteristics to meet diverse mission requirements. However, the drastic and continuous changes in aerodynamic parameters during deformation pose a significant challenge to accurate aerodynamic modeling, while the complex unsteady aerodynamic environment increases the probability of actuator failure.

[0003] Currently, fault-tolerant control methods for morphing aircraft typically rely on explicit fault detection and isolation modules (MDI modules), requiring the pre-establishment of accurate fault diagnosis models to identify the location and severity of faults in real time. However, this approach is difficult to adapt to the rapidly changing dynamic environment of morphing aircraft, and is prone to delays, false alarms, and missed alarms. Summary of the Invention

[0004] This application discloses an implicit fault-tolerant control method for deformable aircraft, electronic equipment, and computer-readable storage medium.

[0005] The implicit fault-tolerant control method for deformable aircraft in this application includes the following steps: Based on the physical property parameters of the deformable aircraft, an attitude dynamics model of the deformable aircraft is constructed; Based on the aerodynamic control surface deflection of the deformable aircraft, a fault model for the bias and efficiency degradation of the deformable aircraft is constructed. Based on the physical property parameters of the deformable aircraft, the attitude dynamics model, and the preset aerodynamic moment derivative, an attitude decoupling model of the deformable aircraft is constructed. Based on the attitude decoupling model, a linear extended state observation model is designed to determine the attitude angle, attitude angular velocity and comprehensive disturbance estimate of the deformable aircraft in multiple attitude control dimensions. Based on the attitude decoupling model and the linear extended state observation model, the equivalent control command of the deformable aircraft is determined to achieve implicit fault-tolerant control of the deformable aircraft in case of failure.

[0006] In some embodiments, constructing the attitude dynamics model of the deformable aircraft based on its physical property parameters includes: Based on the angle of attack, sideslip angle, and roll angle of the deformable aircraft, an attitude state vector is constructed. The attitude dynamics model is constructed based on the attitude state vector, the angular velocity and control torque of the deformable aircraft in multiple coordinate axis directions, and the inertial tensor of the deformable aircraft. The physical property parameters of the deformable aircraft include the angle of attack, the sideslip angle, the roll angle, the angular velocity, the control torque, and the inertial tensor.

[0007] In some implementations, the attitude dynamics model is a nonlinear attitude dynamics differential equation, with the following formula:

[0008] in Let be the attitude state vector. The transformed angular velocity variable is determined based on the angular velocities of the deformable aircraft along multiple coordinate axes. The characterization matrix, used to characterize nonlinear torques and Coriolis coupling effects, is determined based on the attitude state vector, the angular velocities of the deformable aircraft along multiple coordinate axes, and the inertial tensor. The control effectiveness matrix is ​​determined based on the inertia tensor and the attitude state vector. This is a vector formed by the control torques of the deformable aircraft along multiple coordinate axes. The total external disturbance acting on the deformable aircraft.

[0009] In some embodiments, constructing a bias and efficiency degradation fault model for the morphing vehicle based on the aerodynamic control surface deflection yaw rate of the morphing vehicle includes: The actual actuator commands of the deformable aircraft are determined based on the aerodynamic control surface deflection of the deformable aircraft and the preset control allocation matrix. The bias and efficiency degradation fault model is constructed based on the actual actuator instructions, the preset efficiency degradation matrix, and the preset bias fault vector.

[0010] In some embodiments, constructing the attitude decoupling model of the deformable aircraft based on the physical property parameters of the deformable aircraft, the attitude dynamics model, the bias and efficiency degradation fault model, and the preset aerodynamic moment derivative includes: Based on the characterization matrix, the bias and efficiency degradation fault model, and the total external disturbance, determine the comprehensive disturbance matrix; The control gain parameters are determined based on the inertia tensor, the dynamic pressure of the deformable aircraft, and the corresponding aerodynamic torque derivative, wherein the physical property parameters of the deformable aircraft also include the dynamic pressure and the aerodynamic torque derivative; Based on the preset channel equivalent control command, the control gain parameter, and the integrated disturbance matrix, the attitude decoupling model of the deformable aircraft is constructed.

[0011] In some implementations, the step of designing a linear extended state observation model based on the attitude decoupling model to determine the attitude angles, attitude angular velocities, and comprehensive disturbance estimates of the deformable aircraft in multiple attitude control dimensions includes: Based on the channel equivalent control command, the attitude decoupling model, the control gain parameters, and the observer gain based on bandwidth parameter tuning, the linear extended state observation model is designed and constructed.

[0012] In some implementations, the formula for the linearly extended state observation model is:

[0013] in Representing dimension, Indicates the roll angle dimension. Indicates the yaw angle dimension. Indicates the pitch angle dimension. Indicates the first The observed values ​​of the attitude angles in the given dimension. Indicates the first The observed values ​​of the attitude angular velocity in the given dimension. Indicates the first The observed values ​​of the comprehensive perturbation estimate under the given dimension, For the first The control gain parameter in the dimension, For the first The channel equivalent control command in the dimension. For the first The output of the attitude decoupling model in the given dimension. The preset observer bandwidth parameters, For the first The observer gain corresponding to the attitude angle in the given dimension. For the first The observer gain corresponding to the attitude angular velocity in the given dimension. For the first The observer gain corresponding to the comprehensive perturbation estimation under the given dimension.

[0014] In some implementations, determining the equivalent control command for the morphing vehicle based on the attitude decoupling model and the linear extended state observation model to achieve implicit fault-tolerant control of the morphing vehicle in case of a fault includes: Based on the multiple observations output by the linear extended state observation model, the channel equivalent control command, the control gain parameter, and the preset controller bandwidth parameter, the error feedback control law is determined. Based on the error feedback control law, the equivalent control command of the deformable aircraft is determined to achieve implicit fault-tolerant control for the deformable aircraft in case of failure.

[0015] The electronic device in this application includes a memory and a processor. The memory stores a computer program, and when the computer program is executed by the processor, the implicit fault-tolerant control method for deformable aircraft in the above-described embodiments is implemented.

[0016] The computer-readable storage medium in the embodiments of this application stores a computer program that, when executed by one or more processors, implements the implicit fault-tolerant control method for deformable aircraft described above.

[0017] The beneficial effects of this application are as follows: The implicit fault-tolerant control method for deformable aircraft in the embodiments of this application avoids the dependence on explicit fault detection and isolation modules (MDI modules) by constructing attitude dynamics models, bias and efficiency degradation fault models and attitude decoupling models. It maps actuator faults and aerodynamic uncertainties into total disturbances for real-time estimation and feedforward compensation, which significantly reduces the structural and computational complexity of the system. At the same time, even under severe operating conditions such as serious efficiency degradation and bias faults in the actuators, it can still generate smooth and continuous actual control surface deflection commands, ensuring the overall robustness and stability of the closed-loop system. Attached Figure Description

[0018] Figure 1 This is one of the flowcharts illustrating the implicit fault-tolerant control method for deformable aircraft in the embodiments of this application; Figure 2 This is the second flowchart illustrating the implicit fault-tolerant control method for deformable aircraft in the embodiments of this application; Figure 3 This is the third flowchart illustrating the implicit fault-tolerant control method for deformable aircraft in the embodiments of this application; Figure 4 This is the fourth flowchart illustrating the implicit fault-tolerant control method for deformable aircraft in the embodiments of this application; Figure 5 This is the fifth flowchart illustrating the implicit fault-tolerant control method for deformable aircraft in the embodiments of this application. Detailed Implementation

[0019] Please see Figure 1 The implicit fault-tolerant control method for deformable aircraft in this application includes the following steps: Step 01: Construct the attitude dynamics model of the deformable aircraft based on its physical property parameters.

[0020] Specifically, the implicit fault-tolerant control method for deformable aircraft in this application first constructs an attitude dynamics model of the deformable aircraft based on its physical property parameters. The main function of this attitude dynamics model is to integrate fault torque loss during the subsequent establishment of the bias and efficiency degradation fault model, and to provide the input state data foundation for subsequent processes. For further details, please refer to... Figure 2 In some embodiments, step 01 further includes: Step 011: Construct the attitude state vector based on the angle of attack, sideslip angle, and roll angle of the deformable aircraft; Step 012: Based on the attitude state vector, the angular velocities and control torques of the morphing aircraft along multiple coordinate axes, and the inertial tensor of the morphing aircraft, construct an attitude dynamics model. The physical properties of a transforming aircraft include angle of attack, sideslip angle, roll angle, angular velocity, control torque, and inertial tensor.

[0021] Specifically, firstly, based on the attitude sensors installed on the deformable aircraft itself, or based on measuring instruments set in the external environment, the physical property parameters of the deformable aircraft are obtained, including the angle of attack in the global coordinate system. Sideslip angle and roll angle The three angle parameters mentioned above are used as attitude state variables to construct an attitude state vector. , Considering the uncertainties in modeling and factors such as gusts of wind in the external environment, the above attitude state vector is used. Based on the data, an attitude dynamics model is constructed by combining the angular velocities and control torques of the morphing aircraft along multiple coordinate axes and the inertial tensor of the morphing aircraft. The attitude dynamics model is shown in Equation 1.

[0022] ... Formula 1 in This is the attitude state vector. The transformed angular velocity variable is determined based on the angular velocities of the deformable aircraft along multiple coordinate axes. The characterization matrix, used to characterize nonlinear torques and Coriolis coupling effects, is determined based on the attitude state vector, the angular velocities of the deformable aircraft along multiple coordinate axes, and the inertial tensor. The control effectiveness matrix is ​​determined based on the inertia tensor and the attitude state vector. This is a vector formed by the control moments of the deformable aircraft along multiple coordinate axes. This is a unified representation of the total external disturbances acting on a deformable aircraft. Furthermore... This represents the first-order derivative matrix of the attitude state vector with respect to time. This represents the first derivative matrix of the angular velocity variable with respect to time.

[0023] For formula 1 , , , For the definition of , please refer to Formula 2 and Formula 3.

[0024] ... Formula 2 ... Formula 3 in Here is the kinematic transformation matrix. The first derivative of the kinematic transformation matrix with respect to time is given by the matrix. Let be the angular velocity vector of the deformable aircraft in its body coordinate system. Its three components are the three sets of angular velocity components of the deformable aircraft along the three coordinate axes in its body coordinate system (corresponding to...). ), The inertial tensor of the deformable aircraft includes multiple components of the inertial tensor in the three coordinate axes of the body coordinate system (corresponding to...). , , , , In response to x Moment of inertia of the axis, In response to x shaft and y The product of inertia of the axis, In response to y Moment of inertia of the axis, In response to z Moment of inertia of the axis. It is a skew-symmetric operator based on the aforementioned angular velocity vector. The control torque vector applied to the deformable aircraft comprises three components of the control torque in the three coordinate axes of the body coordinate system (corresponding to the above). ).

[0025] Please continue reading. Figure 1 The implicit fault-tolerant control method for deformable aircraft in this application also includes: Step 02: Based on the aerodynamic control surface deflection of the deformable aircraft, construct the bias and efficiency degradation fault model of the deformable aircraft.

[0026] Specifically, the actual actuators of the morphing aircraft in related technologies include four sets of rudders. However, in the design of the control system, three virtual control variables are generally considered to be used in the roll, yaw, and pitch channels, which are then converted by the equivalent control surface deflection allocation matrix. In essence, the morphing aircraft generates control torque through aerodynamic control surface deflection. The constructed bias and efficiency degradation fault model is the core objective of subsequent observation. Furthermore, in some implementations, please refer to... Figure 3 Step 02 further includes: Step 021: Determine the actual actuator commands of the deformable aircraft based on the aerodynamic control surface deflection and the preset control allocation matrix; Step 022: Construct a bias and efficiency degradation fault model based on the actual actuator instructions, the preset efficiency degradation matrix, and the preset bias fault vector.

[0027] Specifically, regarding the actual construction method of the bias and efficiency degradation fault model, for example, based on the control logic described above that generates control torque through aerodynamic control surface deflection, let the aerodynamic control surface deflection magnitude be... ,in These are the virtual control variables for the roll, yaw, and pitch control channels, respectively. Additionally, actual actuator commands are defined. ,in These are the four sets of aerodynamic rudders included in the actual deformable aircraft system, and the command corresponds to the deflection of the aerodynamic control surfaces. The relationship between them is controlled by the allocation matrix. The decision is as shown in Formula 4.

[0028] ... Formula 4 Based on Equation 4, with a known aerodynamic control surface deflection... Under these conditions, combined with the preset control allocation matrix That is, the actual execution mechanism instructions can be calculated. .

[0029] Furthermore, when the actuator malfunctions, the actual control command carrying the fault... (That is, the bias and efficiency degradation failure model mentioned above) is represented by the case shown in Equation 5.

[0030] ... Formula 5 in This is the efficiency degradation matrix of the actual actuators of the deformable aircraft, which includes four sets of efficiency degradation parameters corresponding to the four sets of aerodynamic rudders (corresponding to the above). The value range of each of the above efficiency degradation parameters is (0,1]. This is the actual actuator bias fault vector for the deformable aircraft, which includes the bias degree parameters corresponding to the four sets of aerodynamic rudders (corresponding to the above). Therefore, based on the above bias and efficiency degradation fault model, including the efficiency degradation parameter and the bias parameter, combined with the total external disturbance in Formula 1... The above three factors are then mapped and absorbed into the total disturbance of the three control channels of roll, yaw, and pitch in the subsequent process for implicit processing.

[0031] Please continue reading. Figure 1 The implicit fault-tolerant control method for deformable aircraft in this application also includes: Step 03: Based on the physical property parameters, attitude dynamics model, and preset aerodynamic moment derivatives of the deformable aircraft, construct the attitude decoupling model of the deformable aircraft.

[0032] Specifically, based on the above implementation method, and using the attitude dynamics model determined in step 01, the bias and efficiency degradation fault model determined in step 02, and the physical property parameters of the deformable aircraft itself, an attitude decoupling model is constructed to decouple the deformable aircraft in the roll, yaw, and pitch control channels. The aforementioned physical property parameters also include the dynamic pressure of the environment surrounding the deformable aircraft and the time derivative of the aerodynamic torque. In some implementation methods, please refer to... Figure 4 Step 03 further includes: Step 031: Determine the comprehensive disturbance matrix based on the characterization matrix, the bias and efficiency degradation fault model, and the total external disturbance; Step 032: Determine the control gain parameters based on the inertia tensor, the dynamic pressure of the deformable aircraft, and the corresponding derivative of the aerodynamic torque. The physical property parameters of deformable aircraft also include dynamic pressure and the derivative of aerodynamic torque; Step 033: Construct the attitude decoupling model of the deformable aircraft based on the preset channel equivalent control command, control gain parameters and integrated disturbance matrix.

[0033] Specifically, regarding the specific determination method of the attitude decoupling model, firstly, the index labels corresponding to the three sets of control channels—roll, yaw, and pitch—are set as follows: , as well as The representation matrix in the above implementation method Actual control commands with faults and total external disturbances The disturbances are uniformly mapped and absorbed into the total disturbances of the three control channels: roll, yaw, and pitch, thus defining the integrated disturbance matrix. , that is The comprehensive disturbance matrix corresponding to the roll channel, The integrated disturbance matrix corresponding to the yaw channel, This is the integrated disturbance matrix corresponding to the pitch channel.

[0034] Furthermore, based on the dynamic characteristics of the aircraft, the control gain parameters corresponding to the three sets of control channels can be determined using the method shown in Formula 6. .

[0035] ... Formula 6 in The current dynamic pressure of the deformable drone, For deformable drones in body coordinate system x The derivative of the aerodynamic torque of the shaft, For deformable drones in body coordinate system y The derivative of the aerodynamic torque of the shaft, For deformable drones in body coordinate system z The derivative of the aerodynamic torque of the shaft.

[0036] Furthermore, based on the aforementioned multiple integrated disturbance matrices, control gain parameters, and preset channel equivalent control commands, an attitude decoupling model as shown in Equation 7 can be constructed.

[0037] ... Formula 7 in, For the first The attitude angle state corresponding to the control channel of the dimension. This is the first derivative of the aforementioned attitude angle state with respect to time. For the first The attitude angular velocity state corresponding to the control channel of the dimension. This is the first derivative of the aforementioned attitude angular velocity state with respect to time. For the first The attitude angle acceleration state corresponding to the control channel of the dimension. This is the first derivative of the aforementioned attitude angular acceleration state with respect to time. For the first The equivalent control command for the control channel of a dimension. This represents the desired attitude angle command, and also the output of the attitude decoupling model. For the first The first-order derivative of the integrated disturbance matrix with respect to time for the control channel of the dimension.

[0038] Therefore, for the three control channels of roll, yaw, and pitch, the attitude decoupling models corresponding to the three control channels can be obtained according to Formula 7.

[0039] Please continue reading. Figure 1 The implicit fault-tolerant control method for deformable aircraft in this application also includes: Step 04: Based on the attitude decoupling model, design a linear extended state observation model to determine the attitude angle, attitude angular velocity, and comprehensive disturbance estimate of the deformable aircraft in multiple attitude control dimensions.

[0040] Specifically, based on the above implementation methods, in order to obtain the attitude decoupling model without relying on explicit fault isolation... , and the corresponding integrated perturbation matrix Based on the estimated value, a linear expansion state observation model is designed. Specifically, in some embodiments, step 04 further includes: Based on the channel equivalent control command, attitude decoupling model, control gain parameters, and observer gain based on bandwidth parameter tuning, a linear extended state observation model is designed and constructed.

[0041] Specifically, the control gain parameters are taken from the attitude decoupling model. Channel equivalent control commands and the desired attitude angle command By combining the preset observer gain based on bandwidth parameter tuning, a linear extended state observation model is designed as shown in Equation 8.

[0042] ... Formula 8 in Representing dimension, Indicates the roll angle dimension. Indicates the yaw angle dimension. Indicates the pitch angle dimension. Indicates the first Observed values ​​of attitude angles in dimension 1. Indicates the first Observed values ​​of attitude angular velocity in dimensionality. Indicates the first Observations for comprehensive perturbation estimation under different dimensions. For the first Control gain parameters in dimensionality, For the first Channel equivalent control commands in a given dimension For the first Output of the pose decoupling model in dimensionality The preset observer bandwidth parameters, For the first The observer gain corresponding to the attitude angle in the dimension. For the first The observer gain corresponding to the attitude angular velocity in the dimension. For the first The observer gain corresponding to the comprehensive disturbance estimation in the dimension. In this way, based on the linear extended state observation model shown in Equation 8, the estimated values ​​of attitude angle, attitude angular velocity, and comprehensive disturbance estimation in the attitude control dimension corresponding to the three control channels of roll, yaw, and pitch can be obtained.

[0043] Please continue reading. Figure 1 The implicit fault-tolerant control method for deformable aircraft in this application also includes: Step 05: Based on the attitude decoupling model and the linear extended state observation model, determine the equivalent control command for the deformable aircraft to achieve implicit fault-tolerant control for the deformable aircraft in case of failure.

[0044] Specifically, based on the above implementation method, the output of the linear extended state observation model and the control gain parameters in the attitude decoupling model will be used next. and channel equivalent control commands Based on the data, control commands for four sets of aerodynamic rudders for the deformable aircraft were gradually determined, thereby realizing implicit fault-tolerant control for the deformable aircraft in case of failure.

[0045] Further, please refer to Figure 5 In some implementations, step 05 includes: Step 051: Determine the error feedback control law based on the multiple observations output by the linear extended state observation model, the channel equivalent control command, the control gain parameter, and the preset controller bandwidth parameter; Step 052: Determine the equivalent control command for the deformable aircraft based on the error feedback control law, so as to realize implicit fault-tolerant control for the deformable aircraft in case of failure.

[0046] Specifically, for the process of determining the control command, firstly, based on the real-time output of each observation value of the linear extended state observation model, an error feedback control law is designed using a proportional-differential structure, as shown in Equation 9.

[0047] ... Formula 9 in For the first The original control command without disturbance compensation in the dimension, For the first Disturbance rejection control commands after excluding disturbance observations in the dimension. For the first feedback controller gain, For the gain of the second feedback controller, For the first Controller bandwidth parameters in the dimension For the first Preset channel reference attitude commands in the dimension.

[0048] Finally, based on the error feedback control law shown in Formula 9, the disturbance rejection control commands corresponding to the three control channels of roll, yaw, and pitch can be calculated. , as well as Then, further combining the pseudo-inverse matrix of the control allocation matrix in the above embodiments... This means that the anti-disturbance control commands of the three control channels mentioned above can be converted back into equivalent control commands for the four sets of aerodynamic rudders (i.e., actuators) of the deformable aircraft. These equivalent control commands are generally expressed as actual deflection angle commands. When the aerodynamic rudder executes the command, it adjusts its own deflection angle according to the command, thereby realizing implicit fault-tolerant control for the deformable aircraft in the event of a fault.

[0049] Thus, the implicit fault-tolerant control method for deformable aircraft in this application avoids dependence on explicit fault detection and isolation modules (MDI modules) by constructing attitude dynamics models, bias and efficiency degradation fault models, and attitude decoupling models. It maps actuator faults and aerodynamic uncertainties into total disturbances for real-time estimation and feedforward compensation, significantly reducing the structural and computational complexity of the system. At the same time, even under severe operating conditions such as serious efficiency degradation and bias faults in the actuators, it can still generate smooth and continuous actual control surface deflection commands, ensuring the overall robustness and stability of the closed-loop system.

[0050] The electronic device in this application includes a memory and a processor. The memory stores a computer program, and when the computer program is executed by the processor, the implicit fault-tolerant control method for deformable aircraft in the above-described embodiments is implemented.

[0051] The computer-readable storage medium in the embodiments of this application stores a computer program that, when executed by one or more processors, implements the implicit fault-tolerant control method for deformable aircraft described above.

[0052] The above description is merely a preferred embodiment of this application and is not intended to limit this application in any way. Although this application has disclosed the preferred embodiment as above, it is not intended to limit this application. Any person skilled in the art can make some modifications or alterations to the above-disclosed technical content to create equivalent embodiments without departing from the scope of the technical solution of this application. Any simple modifications, equivalent substitutions, and improvements made to the above embodiments without departing from the technical solution of this application, based on the technical essence of this application and within the spirit and principles of this application, shall still fall within the protection scope of the technical solution of this application.

Claims

1. An implicit fault-tolerant control method for a deformable aircraft, characterized in that, The method includes: Based on the physical property parameters of the deformable aircraft, an attitude dynamics model of the deformable aircraft is constructed; Based on the aerodynamic control surface deflection of the deformable aircraft, a fault model for the bias and efficiency degradation of the deformable aircraft is constructed. Based on the physical property parameters of the deformable aircraft, the attitude dynamics model, the bias and efficiency decay fault model, and the preset aerodynamic moment derivative, an attitude decoupling model of the deformable aircraft is constructed. Based on the attitude decoupling model, a linear extended state observation model is designed to determine the attitude angle, attitude angular velocity and comprehensive disturbance estimate of the deformable aircraft in multiple attitude control dimensions. Based on the attitude decoupling model and the linear extended state observation model, the equivalent control command of the deformable aircraft is determined to achieve implicit fault-tolerant control of the deformable aircraft in case of failure.

2. The method according to claim 1, characterized in that, The step of constructing the attitude dynamics model of the deformable aircraft based on its physical property parameters includes: Based on the angle of attack, sideslip angle, and roll angle of the deformable aircraft, an attitude state vector is constructed. The attitude dynamics model is constructed based on the attitude state vector, the angular velocity and control torque of the deformable aircraft in multiple coordinate axis directions, and the inertial tensor of the deformable aircraft. The physical property parameters of the deformable aircraft include the angle of attack, the sideslip angle, the roll angle, the angular velocity, the control torque, and the inertial tensor.

3. The method according to claim 2, characterized in that, The attitude dynamics model is a nonlinear attitude dynamics differential equation, and the formula is: in Let be the attitude state vector. The transformed angular velocity variable is determined based on the angular velocities of the deformable aircraft along multiple coordinate axes. The characterization matrix, used to characterize nonlinear torques and Coriolis coupling effects, is determined based on the attitude state vector, the angular velocities of the deformable aircraft along multiple coordinate axes, and the inertial tensor. The control effectiveness matrix is ​​determined based on the inertia tensor and the attitude state vector. This is a vector formed by the control torques of the deformable aircraft along multiple coordinate axes. The total external disturbance acting on the deformable aircraft.

4. The method according to claim 1, characterized in that, The process of constructing an attitude decoupling model for the deformable aircraft based on its physical property parameters, attitude dynamics model, bias and efficiency degradation fault model, and preset aerodynamic moment derivatives includes: The actual actuator commands of the deformable aircraft are determined based on the aerodynamic control surface deflection of the deformable aircraft and the preset control allocation matrix. The bias and efficiency degradation fault model is constructed based on the actual actuator instructions, the preset efficiency degradation matrix, and the preset bias fault vector.

5. The method according to claim 3, characterized in that, The step of constructing the attitude decoupling model of the deformable aircraft based on the physical property parameters of the deformable aircraft, the attitude dynamics model, the bias and efficiency decay fault model, and the preset aerodynamic moment derivative includes: Based on the characterization matrix, the bias and efficiency degradation fault model, and the total external disturbance, determine the comprehensive disturbance matrix; The control gain parameters are determined based on the inertia tensor, the dynamic pressure of the deformable aircraft, and the corresponding aerodynamic torque derivative, wherein the physical property parameters of the deformable aircraft also include the dynamic pressure and the aerodynamic torque derivative; Based on the preset channel equivalent control command, the control gain parameter, and the integrated disturbance matrix, the attitude decoupling model of the deformable aircraft is constructed.

6. The method according to claim 5, characterized in that, The step of designing a linear extended state observation model based on the attitude decoupling model to determine the attitude angles, attitude angular velocities, and comprehensive disturbance estimates of the deformable aircraft in multiple attitude control dimensions includes: Based on the channel equivalent control command, the attitude decoupling model, the control gain parameters, and the observer gain based on bandwidth parameter tuning, the linear extended state observation model is designed and constructed.

7. The method according to claim 6, characterized in that, The formula for the linear expansion state observation model is: in Representing dimension, Indicates the roll angle dimension. Indicates the yaw angle dimension. Indicates the pitch angle dimension. Indicates the first The observed values ​​of the attitude angles in the given dimension. Indicates the first The observed values ​​of the attitude angular velocity in the given dimension. Indicates the first The observed values ​​of the comprehensive perturbation estimate under the given dimension, For the first The control gain parameter in the dimension, For the first The channel equivalent control command in the dimension. For the first The output of the attitude decoupling model in the given dimension. The preset observer bandwidth parameters, For the first The observer gain corresponding to the attitude angle in the given dimension. For the first The observer gain corresponding to the attitude angular velocity in the given dimension. For the first The observer gain corresponding to the comprehensive perturbation estimation under the given dimension.

8. The method according to claim 7, characterized in that, The step of determining the equivalent control command for the deformable aircraft based on the attitude decoupling model and the linear extended state observation model to achieve implicit fault-tolerant control for the deformable aircraft in case of failure includes: Based on the multiple observations output by the linear extended state observation model, the channel equivalent control command, the control gain parameter, and the preset controller bandwidth parameter, the error feedback control law is determined. Based on the error feedback control law, the equivalent control command of the deformable aircraft is determined to achieve implicit fault-tolerant control for the deformable aircraft in case of failure.

9. An electronic device, characterized in that, The electronic device includes a memory and a processor. The memory stores a computer program that, when executed by the processor, implements the implicit fault-tolerant control method for deformable aircraft as described in any one of claims 1-8.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that, when executed by one or more processors, implements the implicit fault-tolerant control method for deformable aircraft as described in any one of claims 1-8.

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