Fault-tolerant control method and device for aircraft, aircraft, and storage medium

Abstract

This invention belongs to the field of aircraft technology and discloses a fault-tolerant control method, device, aircraft, and storage medium for aircraft. The method includes: fitting the speed transition process during motor failure to obtain a time-varying speed curve; determining the disturbance amount generated by the motor failure on each control channel based on the time-varying speed curve; establishing an extended state observer based on the disturbance amount corresponding to each control channel; compensating the control input value based on the extended state observer to generate a target control quantity; and allocating the target control quantity according to a control allocation matrix. By considering the disturbance amount during the speed transition process during motor failure and compensating for this disturbance using an extended state observer, the stability and accuracy of aircraft control are improved.

Description

Technical Field

[0001] This invention relates to the field of aircraft technology, and in particular to a fault-tolerant control method, device, aircraft, and storage medium for aircraft. Background Technology

[0002] When an aircraft's motor fails, the speed does not drop to 0 rpm instantly, but rather decreases slowly over a period of time. Current fault tolerance methods for rotorcraft directly assume that the speed is 0 after failure, without considering the process of motor speed change, which affects the stability and accuracy of aircraft control.

[0003] The above content is only used to help understand the technical solution of the present invention and does not represent an admission that the above content is prior art. Summary of the Invention

[0004] The main objective of this invention is to provide a fault-tolerant control method, device, aircraft, and storage medium for aircraft, aiming to solve the technical problem that current fault-tolerant methods for rotorcraft do not consider the change process of motor speed, which affects the stability and accuracy of aircraft control.

[0005] To achieve the above objectives, the present invention provides a fault-tolerant control method for an aircraft, the method comprising the following steps:

[0006] The speed transition process when the motor fails is fitted to obtain the speed curve that changes with time;

[0007] The disturbance generated by motor failure in each control channel is determined based on the time-varying speed curve.

[0008] An extended state observer is established based on the disturbance amount corresponding to each control channel;

[0009] The target control quantity is generated by compensating the control input value based on the extended state observer.

[0010] The target control quantity is allocated according to the control allocation matrix.

[0011] Optionally, before allocating the target control quantity according to the control allocation matrix, the method further includes:

[0012] Calculate the manipulation efficiency matrix;

[0013] The control allocation matrix is ​​calculated based on the preset weight matrix and the manipulation efficiency matrix.

[0014] Optionally, the calculation of the manipulation efficiency matrix includes:

[0015] Determine the failure type of the motor, including upper propeller failure or lower propeller failure;

[0016] Select the corresponding matrix model based on the failure type;

[0017] Determine the linear proportionality coefficient between the lift generated by a single propeller and the lift generated by two propellers;

[0018] Substituting the linear scaling coefficient into the selected matrix model yields the manipulation efficiency matrix.

[0019] Optionally, determining the disturbance amount generated by the motor failure in each control channel based on the time-varying speed curve includes:

[0020] The control gain on each control channel is determined based on the maximum speed of the motor, the mass of the aircraft, the lift coefficient of the propeller, the distance between the two propellers on the same side, and the moment of inertia.

[0021] The disturbance amount generated by motor failure in each control channel is determined based on the control gain and the time-varying speed curve.

[0022] Optionally, before compensating the control input value based on the extended state observer to generate the target control quantity, the method further includes:

[0023] Establish system models for each control channel;

[0024] Based on the angular acceleration limit value, model predictive control is performed on the system model of each control channel to generate control input values ​​that meet the constraints.

[0025] Optionally, the step of performing model predictive control on the system model of each control channel based on the angular acceleration limit value to generate control input values ​​that satisfy the constraints includes:

[0026] The optimization objective is to minimize the sum of the squared angular error, the squared angular rate error, and the squared angular acceleration error. Based on the angular acceleration limit value, model predictive control is performed on the system model of each control channel to generate control input values ​​that meet the constraints.

[0027] Optionally, fitting the speed transition process when the motor fails to obtain a speed curve that varies with time includes:

[0028] Get the current speed of the motor when it fails;

[0029] The rotational speed curve is obtained by fitting the second-order linear transfer function based on the current rotational speed.

[0030] Furthermore, to achieve the above objectives, the present invention also proposes a fault-tolerant control device for an aircraft, the fault-tolerant control device comprising:

[0031] The fitting module is used to fit the speed transition process when the motor fails, and obtain the speed curve that changes over time.

[0032] The determination module is used to determine the disturbance amount generated by the motor failure in each control channel based on the time-varying speed curve;

[0033] An observer establishment module is used to establish an extended state observer based on the disturbance amount corresponding to each control channel.

[0034] The compensation module is used to compensate the control input value based on the extended state observer and generate the target control quantity;

[0035] The control module is used to control and allocate the target control quantity according to the control allocation matrix.

[0036] Furthermore, to achieve the above objectives, the present invention also proposes an aircraft comprising: a memory, a processor, and a fault-tolerant control program for the aircraft stored in the memory and executable on the processor, the fault-tolerant control program being configured to implement the fault-tolerant control method for the aircraft as described above.

[0037] Furthermore, to achieve the above objectives, the present invention also proposes a storage medium storing a fault-tolerant control program for an aircraft, wherein the fault-tolerant control program for the aircraft, when executed by a processor, implements the fault-tolerant control method for the aircraft as described above.

[0038] This invention obtains a time-varying speed curve by fitting the speed transition process during motor failure; it then determines the disturbances generated by motor failure in each control channel based on these curves; an extended state observer is established based on the disturbances corresponding to each control channel; the control input values ​​are compensated using the extended state observer to generate a target control quantity; and the target control quantity is allocated according to the control allocation matrix. By considering the disturbances during the speed transition process of motor failure and compensating for these disturbances using the extended state observer, the stability and accuracy of aircraft control are improved. Attached Figure Description

[0039] Figure 1 This is a schematic diagram of the structure of the aircraft in the hardware operating environment involved in the embodiments of the present invention;

[0040] Figure 2 This is a flowchart illustrating the first embodiment of the fault-tolerant control method for the aircraft of the present invention;

[0041] Figure 3 This is a second-order fitting curve diagram of the first embodiment of the fault-tolerant control method for the aircraft of the present invention;

[0042] Figure 4 This is a flowchart illustrating the second embodiment of the fault-tolerant control method for the aircraft of the present invention;

[0043] Figure 5 This is a schematic diagram of the structure of the quadcopter-eight-propeller aircraft in this invention;

[0044] Figure 6 This is a flowchart illustrating the third embodiment of the fault-tolerant control method for the aircraft of the present invention;

[0045] Figure 7 This is a schematic diagram of the controller structure according to an example of the present invention;

[0046] Figure 8 This is a structural block diagram of the first embodiment of the fault-tolerant control device for the aircraft of the present invention.

[0047] The realization of the objective, functional features and advantages of the present invention will be further explained in conjunction with the embodiments and with reference to the accompanying drawings. Detailed Implementation

[0048] It should be understood that the specific embodiments described herein are for illustrative purposes only and are not intended to limit the scope of the invention.

[0049] Reference Figure 1 , Figure 1 This is a schematic diagram of the aircraft structure in the hardware operating environment involved in the embodiments of the present invention.

[0050] like Figure 1 As shown, the aircraft may include: a processor 1001, such as a central processing unit (CPU), a communication bus 1002, a user interface 1003, a network interface 1004, and a memory 1005. The communication bus 1002 is used to enable communication between these components. The user interface 1003 may include a display screen or an input unit such as a keyboard; optionally, the user interface 1003 may also include a standard wired interface or a wireless interface. The network interface 1004 may optionally include a standard wired interface or a wireless interface (such as a Wireless-Fidelity (Wi-Fi) interface). The memory 1005 may be high-speed random access memory (RAM) or stable non-volatile memory (NVM), such as a disk drive. The memory 1005 may also optionally be a storage device independent of the aforementioned processor 1001.

[0051] Those skilled in the art will understand that Figure 1The structure shown does not constitute a limitation on the aircraft and may include more or fewer components than shown, or combine certain components, or have different component arrangements.

[0052] like Figure 1 As shown, the memory 1005, which serves as a storage medium, may include an operating system, a network communication module, a user interface module, and the aircraft's fault-tolerant control program.

[0053] exist Figure 1 In the aircraft shown, the network interface 1004 is mainly used for data communication with the network server; the user interface 1003 is mainly used for data interaction with the user; the processor 1001 and the memory 1005 in the aircraft of the present invention can be set in the aircraft. The aircraft calls the fault-tolerant control program of the aircraft stored in the memory 1005 through the processor 1001 and executes the fault-tolerant control method of the aircraft provided in the embodiment of the present invention.

[0054] This invention provides a fault-tolerant control method for an aircraft, referring to... Figure 2 , Figure 2 This is a flowchart illustrating the first embodiment of the fault-tolerant control method for the aircraft of the present invention.

[0055] In this embodiment, the fault-tolerant control method for the aircraft includes the following steps:

[0056] Step S10: Fit the speed transition process when the motor fails to obtain the speed curve that changes with time.

[0057] It is understood that the execution entity in this embodiment can be a controller installed on the aircraft, and this embodiment is not limited to this. When the motor fails, the motor speed does not drop to 0 rpm instantaneously, but gradually decreases from the current speed to 0. Optionally, this embodiment performs linear or nonlinear fitting on the speed transition process. This embodiment does not limit the specific methods of linear and nonlinear fitting. For example, the speed deceleration process can be fitted as a second-order transition process, and more complex nonlinear fitting methods can also be used.

[0058] It should be noted that in the time-varying speed curve graph, the horizontal axis represents time and the vertical axis represents speed, showing the process of the motor speed dropping to 0 over a period of time. In the specific implementation, the time it takes for the motor speed to drop from the current speed to 0 when the motor fails is determined experimentally. When performing curve fitting, this pre-stored time is used as the basis for fitting, which improves the efficiency of curve fitting.

[0059] Optionally, step S10 includes: obtaining the current speed when the motor fails; and fitting the current speed based on a second-order linear transfer function to obtain a speed curve that changes over time.

[0060] It should be understood that the maximum speed of the time-varying speed curve is the current speed when the motor fails, and the expression of the second-order linear transfer function is as follows:

[0061]

[0062] Where G(s) is the second-order linear transfer function, s is the Laplace transform factor, ε is the damping factor, and w is the corner frequency.

[0063] The second-order fitting effect of the speed change is as follows Figure 3 As shown, Figure 3 This is a second-order fitting curve diagram of the first embodiment of the fault-tolerant control method for the aircraft of the present invention, where the horizontal axis represents time and the vertical axis represents rotational speed. At 5 seconds, the motor fails, and the motor speed decays from 750 rpm to 0 in a second-order manner. Optionally, the time required for the motor speed to drop to 0 is pre-stored, and a second-order linear transfer function is fitted based on the current speed and the pre-stored time to obtain a speed curve that changes over time, thus improving curve fitting efficiency.

[0064] Step S20: Determine the disturbance amount generated by the motor failure in each control channel based on the time-varying speed curve.

[0065] It should be noted that the control channels include pitch, roll, yaw, and altitude channels. Based on the current time after motor failure, the motor speed corresponding to the current time is retrieved from the time-varying speed curve, and the disturbance amount generated by the motor failure in each control channel is determined based on the motor speed.

[0066] Specifically, step S20 includes: determining the control gain on each control channel based on the maximum motor speed, aircraft mass, propeller lift coefficient, distance between two propellers on the same side, and moment of inertia; and determining the disturbance amount generated on each control channel by motor failure based on the control gain and the time-varying speed curve.

[0067] It should be understood that this embodiment uses a quadcopter with eight rotors as an example for illustration: assuming the square of the rotational speed is proportional to the lift, and the lift coefficients of the upper and lower rotors are equal, the control gain b on each control channel can be expressed by the following formula:

[0068]

[0069]

[0070]

[0071]

[0072] Where, ωmax Where m is the maximum speed of the motor, and C is the mass of the quadcopter-eight-propeller aircraft. t I is the lift coefficient of the propeller, l is the distance between the two propellers on the same side, and I xx I yy I zz These are the moments of inertia of the aircraft's roll, pitch, and yaw channels, respectively.

[0073] When a motor fails, the torque disturbance generated at different motor speeds is as follows:

[0074]

[0075]

[0076]

[0077]

[0078] Where ω is the motor speed at the current time on the speed curve that changes with time.

[0079] The disturbance generated during this speed transition process can be expressed by the following formula:

[0080]

[0081]

[0082]

[0083]

[0084] Among them, b roll b pitch b yaw and b height Here, f represents the control gain on each control channel, ω represents the motor speed at the current time on the time-varying speed curve, and f represents the speed gain on each channel. droll f dpitch f dyaw and f dheight This refers to the disturbance generated in each control channel due to motor failure.

[0085] Step S30: Establish an extended state observer based on the disturbance amount corresponding to each control channel.

[0086] It should be noted that when the motor fails, the extended state observer can be used to observe and compensate for the disturbance caused by the failure. The current expression for the extended state observer is:

[0087]

[0088]

[0089] Where z1 represents the estimated pitch, roll, and yaw rates, y represents the actual pitch, roll, and yaw rates, b represents the control gain, u represents the virtual control variable, and β1 and β2 represent the virtual control variables. f is an adjustable parameter, f is the disturbance estimate. The extended state observer can estimate the magnitude of the disturbance in the system using the above formula, but this method does not take into account the disturbance effect caused by the gradual change in speed during motor failure.

[0090] In this embodiment, the improved extended state observer is represented by the following formula:

[0091]

[0092]

[0093] Optionally, f d Let f be respectively droll f dpitch f dyaw and f dheight This can compensate for disturbances in the pitch, roll, yaw, and altitude channels caused by motor failure transients. Preferably, yaw channel disturbances are not considered to reduce the computational load on the controller. d Let f be respectively droll f dpitch and f dheight It compensates for the pitch, roll, and height channel disturbances caused during the transition process of motor failure.

[0094] Step S40: Compensate the control input value based on the extended state observer to generate the target control quantity.

[0095] It should be understood that the control input value is the state quantity of the aircraft system. Preferably, the system state quantity over a certain period of time is predicted using a model predictive control method as the control input value. In specific implementation, an active disturbance rejection controller (ADRC) is constructed based on the improved extended state observer of this embodiment, and the ADRC generates the target control quantity. The ADRC includes a tracking differentiator, an extended state observer, and an error feedback controller. The tracking differentiator is used to arrange the transient process for the control input value and extract the input signal containing random noise and its derivative signal. The extended state observer is used to estimate the real-time effect value of internal and external disturbances to the system and compensate for them in the feedback to eliminate the influence of disturbances and make the system anti-interference. The error feedback controller is used to suppress disturbances. Based on the input signal and its derivative signal given by the tracking differentiator and the system state and its derivative obtained by the extended state observer, the error is calculated, and the control quantity is calculated using a nonlinear combination method to compensate for the disturbance.

[0096] Step S50: Perform control allocation on the target control quantity according to the control allocation matrix.

[0097] It should be noted that the control allocation matrix is ​​used to map the virtual control output to the output of the specific actuator, that is, to obtain the mapping matrix from the expected lift and expected torque to the expected speed of each motor. This embodiment does not limit the solution process of the control allocation matrix.

[0098] This embodiment fits the speed transition process during motor failure to obtain a time-varying speed curve; based on the time-varying speed curve, it determines the disturbance amount generated by motor failure in each control channel; based on the disturbance amount corresponding to each control channel, it establishes an extended state observer; based on the extended state observer, it compensates for the control input value to generate a target control quantity; and based on the control allocation matrix, it allocates the target control quantity. Through this method, the disturbance amount in the speed transition process during motor failure is considered, and this disturbance is compensated for by the extended state observer, thus improving the stability and accuracy of aircraft control.

[0099] refer to Figure 4 , Figure 4 This is a flowchart illustrating the second embodiment of the fault-tolerant control method for the aircraft of the present invention.

[0100] Based on the first embodiment described above, the fault-tolerant control method for the aircraft in this embodiment further includes, before step S50:

[0101] Step S501: Calculate the manipulation efficiency matrix.

[0102] Further, step S501 includes: determining the failure type of the motor failure, the failure type including upper propeller failure or lower propeller failure; selecting a corresponding matrix model according to the failure type; determining the linear proportionality coefficient between the lift generated by the single propeller and the lift generated by both propellers; and substituting the linear proportionality coefficient into the selected matrix model to obtain the control efficiency matrix.

[0103] It should be understood that current fault-tolerant methods for rotorcraft are mostly designed for the failure of the motor portion of a quadcopter, with very little research on fault-tolerant control for the complete motor failure of a quadcopter with eight rotors. While there are similarities, the motor failure problem in quadcopters with eight rotors has its own unique characteristics. Furthermore, it is often assumed that the lift generated by a single rotor after motor failure is half that of a dual rotor. However, the lift generated by a single rotor and a dual rotor is not a simple linear relationship, causing the calculated control efficiency matrix and control allocation matrix to not accurately reflect the actual situation. This embodiment, however, considers the difference in control efficiency between a single rotor and a dual rotor, adjusts the control allocation matrix, and eliminates the impact of the lift difference between single and dual rotors on control performance.

[0104] It should be noted that, referring to Figure 5 , Figure 5 This is a schematic diagram of the quadcopter with eight rotors in this invention. Rotors 1, 3, 5, and 7 rotate counterclockwise, while rotors 2, 4, 6, and 8 rotate clockwise. Because the downwash airflow generated by the coaxial upper rotor affects the lower rotor, the lift generated by the lower rotor is slightly reduced. Let k be the linear proportionality coefficient between the lift generated by a single rotor and the lift generated by two rotors, k∈(0,1), then its control efficiency matrix is:

[0105]

[0106] If the upper propeller fails, the lower propeller's lift will no longer be affected by the upper propeller; conversely, if the lower propeller fails, it will not affect the lift generated by the upper propeller. Therefore, the control allocation matrix is ​​considered in two cases: upper propeller failure and lower propeller failure.

[0107] Reference Figure 5 Assuming the No. 1 motor of the upper propeller fails, the control efficiency matrix becomes:

[0108]

[0109] Assuming the lower propeller motor #6 fails, the control efficiency matrix becomes:

[0110]

[0111] Substituting the linear scaling factor k into the selected matrix model yields the manipulation efficiency matrix.

[0112] Step S502: Calculate the control allocation matrix based on the preset weight matrix and the manipulation efficiency matrix.

[0113] It should be noted that the control allocation matrix is ​​calculated using the following formula:

[0114] B + =B T (B T ) -1 ;

[0115] Where W is a preset weight matrix, which is a diagonal matrix, and B + B is the control allocation matrix, and B is the manipulation efficiency matrix. T Let B be the transpose of B. By setting different control efficiency matrices, the disturbances caused by the inconsistency in lift efficiency when the single propeller and the twin propeller are working can be canceled out.

[0116] This embodiment takes into account the difference in control efficiency between single-propeller and dual-propeller systems. By selecting a failure type matrix model and adjusting the control allocation matrix, the impact of lift differences between single-propeller and dual-propeller systems on control performance is eliminated, further improving the stability and accuracy of aircraft control.

[0117] refer to Figure 6 , Figure 6 This is a flowchart illustrating the third embodiment of the fault-tolerant control method for the aircraft of the present invention.

[0118] Based on the first embodiment described above, the fault-tolerant control method for the aircraft in this embodiment further includes, before step S40:

[0119] Step S401: Establish the system model for each control channel.

[0120] Understandably, after a motor failure, the control reachability of the aircraft will decrease. Generally, there is no control input limitation processing after the failure, which may lead to control capability saturation. This embodiment uses model prediction to limit the maximum angular acceleration of the control system model, thereby providing sufficiently smooth angle and angular velocity commands under limiting conditions. Taking the pitch channel as an example, let's assume the pitch channel system is a simple dual-integral system:

[0121]

[0122] Where θ is the pitch angle, q is the pitch rate, and α is the pitch acceleration. The dual-integral system model is merely an example and does not constitute a limitation on the system model of each control channel in this embodiment.

[0123] Step S402: Perform model predictive control on the system model of each control channel based on the angular acceleration limit value to generate control input values ​​that meet the constraints.

[0124] Specifically, step S402 includes: minimizing the sum of the squared angular error, the squared angular rate error, and the squared angular acceleration error as the optimization objective; performing model predictive control on the system model of each control channel based on the angular acceleration limit value; and generating control input values ​​that meet the constraints.

[0125] It should be noted that the constraint condition in this embodiment is that the angular acceleration is less than a preset angular acceleration limit value. The angular acceleration limit value is set based on the safe angular acceleration value that the evaluation system can provide after motor failure. Taking the pitch channel as an example, the predicted model in MPC control (Model Predictive Control) used in this embodiment can be expressed by the following formula:

[0126] minω1e T e+ω2q T q+ω3αT α

[0127] st:

[0128]

[0129] - max <α< max ;

[0130] Where e is the error between the desired pitch angle and the actual pitch angle, and r max This is the angular acceleration limit value.

[0131] In practical implementation, minimizing the sum of the squared pitch angle error, squared pitch rate error, and squared pitch acceleration error of the aircraft is taken as the optimization objective of the problem model. After obtaining the optimal cost function, the system state variables over a period of time can be predicted in advance through the system model, thereby generating the optimal trajectory that meets the constraints and obtaining the control input value. The disturbance is compensated by the active disturbance rejection controller, and then mapped to the output of the specific actuator according to the control allocation matrix to realize the fault-tolerant control of the aircraft motor failure.

[0132] Reference Figure 7 , Figure 7 This is a schematic diagram of the controller structure of an example of the present invention; where r_cmd refers to the feedforward command, α is the angular acceleration, and the simplified system model is optimized using the model predictive control method, without directly controlling the measured values ​​of the actual system, which can greatly reduce the calculation frequency of model predictive control; then, after the ideal optimal trajectory under the limit of angular acceleration α is given by the model predictive control, the control input value x is output to the active disturbance rejection controller, the active disturbance rejection controller estimates the effect value of the disturbance f_d, and outputs the virtual control quantity u, which is mapped to the actual system through the control allocation matrix. The specific actuator of the actual system executes the control quantity y, thereby controlling the real system and eliminating the disturbance.

[0133] In this embodiment, model predictive control is used to limit the magnitude of the angular acceleration of the feedforward after motor failure. This makes it less likely for the system to enter a control saturation state after motor failure, and generates control input values ​​that meet the constraints, thereby further improving the stability and accuracy of aircraft control.

[0134] Furthermore, this embodiment of the invention also proposes a storage medium storing a fault-tolerant control program for an aircraft. When the fault-tolerant control program is executed by a processor, it implements the fault-tolerant control method for the aircraft as described above.

[0135] Since this storage medium adopts all the technical solutions of all the above embodiments, it has at least all the beneficial effects brought about by the technical solutions of the above embodiments, which will not be repeated here.

[0136] Reference Figure 8 , Figure 8 This is a structural block diagram of the first embodiment of the fault-tolerant control device for the aircraft of the present invention.

[0137] like Figure 8 As shown, the fault-tolerant control device for an aircraft proposed in this embodiment of the invention includes:

[0138] The fitting module 10 is used to fit the speed transition process when the motor fails, and obtain the speed curve that changes over time.

[0139] The determination module 20 is used to determine the disturbance amount generated by the motor failure on each control channel based on the time-varying speed curve.

[0140] The observer establishment module 30 is used to establish an extended state observer based on the disturbance amount corresponding to each control channel.

[0141] The compensation module 40 is used to compensate the control input value based on the extended state observer to generate the target control quantity.

[0142] The control module 50 is used to control and allocate the target control quantity according to the control allocation matrix.

[0143] It should be understood that the above are merely illustrative examples and do not constitute any limitation on the technical solutions of the present invention. In specific applications, those skilled in the art can make settings as needed, and the present invention does not impose any restrictions on this.

[0144] This embodiment fits the speed transition process during motor failure to obtain a time-varying speed curve; based on the time-varying speed curve, it determines the disturbance amount generated by motor failure in each control channel; based on the disturbance amount corresponding to each control channel, it establishes an extended state observer; based on the extended state observer, it compensates for the control input value to generate a target control quantity; and based on the control allocation matrix, it allocates the target control quantity. Through this method, the disturbance amount in the speed transition process during motor failure is considered, and this disturbance is compensated for by the extended state observer, thus improving the stability and accuracy of aircraft control.

[0145] It should be noted that the workflow described above is merely illustrative and does not limit the scope of protection of this invention. In practical applications, those skilled in the art can select some or all of the workflow to achieve the purpose of this embodiment according to actual needs, and no restrictions are imposed here.

[0146] In addition, for technical details not described in detail in this embodiment, please refer to the fault-tolerant control method for aircraft provided in any embodiment of the present invention, which will not be repeated here.

[0147] In one embodiment, the control module 50 is further configured to calculate the manipulation efficiency matrix and calculate the control allocation matrix based on the preset weight matrix and the manipulation efficiency matrix.

[0148] In one embodiment, the control module 50 is further configured to determine the failure type of the motor failure, the failure type including upper propeller failure or lower propeller failure; select a corresponding matrix model according to the failure type; determine the linear proportionality coefficient between the lift generated by the single propeller and the lift generated by both propellers; and substitute the linear proportionality coefficient into the selected matrix model to obtain the control efficiency matrix.

[0149] In one embodiment, the determining module 20 is further configured to determine the control gain on each control channel based on the maximum motor speed, aircraft mass, propeller lift coefficient, distance between two propellers on the same side, and moment of inertia; and to determine the disturbance amount generated on each control channel by motor failure based on the control gain and the time-varying speed curve.

[0150] In one embodiment, the fault-tolerant control device of the aircraft further includes a prediction module;

[0151] The prediction module is used to establish system models for each control channel; and to perform model predictive control on the system models of each control channel based on the angular acceleration limit value, thereby generating control input values ​​that meet the constraints.

[0152] In one embodiment, the prediction module is further configured to take the minimization of the sum of the squared angular error, the squared angular rate error, and the squared angular acceleration error as the optimization objective, and perform model predictive control on the system model of each control channel according to the angular acceleration limit value to generate control input values ​​that meet the constraints.

[0153] In one embodiment, the fitting module 10 is further configured to obtain the current speed when the motor fails; and to perform fitting based on the current speed using a second-order linear transfer function to obtain a speed curve that changes over time.

[0154] Furthermore, it should be noted that, in this document, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or system that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or system. Unless otherwise specified, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or system that includes that element.

[0155] The sequence numbers of the above embodiments of the present invention are for descriptive purposes only and do not represent the superiority or inferiority of the embodiments.

[0156] Through the above description of the embodiments, those skilled in the art can clearly understand that the methods of the above embodiments can be implemented by means of software plus necessary general-purpose hardware platforms. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation method. Based on this understanding, the technical solution of the present invention, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product is stored in a storage medium (such as read-only memory (ROM) / RAM, magnetic disk, optical disk) and includes several instructions to cause a terminal device (which may be a mobile phone, computer, server, or network device, etc.) to execute the methods described in the various embodiments of the present invention.

[0157] The above are merely preferred embodiments of the present invention and do not limit the scope of the patent. Any equivalent structural or procedural transformations made based on the description and drawings of the present invention, or direct or indirect applications in other related technical fields, are similarly included within the scope of patent protection of the present invention.

Claims

1. A fault-tolerant control method for an aircraft, characterized in that, The fault-tolerant control method for the aircraft includes: The speed transition process when the motor fails is fitted to obtain a speed curve that changes with time. The speed transition process is the process by which the motor speed gradually changes from the current speed at the time of failure to zero. The disturbance generated by motor failure in each control channel is determined based on the time-varying speed curve. An extended state observer is established based on the disturbance amount corresponding to each control channel; The target control quantity is generated by compensating the control input value based on the extended state observer. The target control quantity is allocated according to the control allocation matrix.

2. The fault-tolerant control method for an aircraft as described in claim 1, characterized in that, Before allocating the target control quantity according to the control allocation matrix, the method further includes: Calculate the manipulation efficiency matrix; The control allocation matrix is ​​calculated based on the preset weight matrix and the manipulation efficiency matrix.

3. The fault-tolerant control method for an aircraft as described in claim 2, characterized in that, The computational manipulation efficiency matrix includes: Determine the failure type of the motor, including upper propeller failure or lower propeller failure; Select the corresponding matrix model based on the failure type; Determine the linear proportionality coefficient between the lift generated by a single propeller and the lift generated by two propellers; Substituting the linear scaling coefficient into the selected matrix model yields the manipulation efficiency matrix.

4. The fault-tolerant control method for an aircraft as described in claim 1, characterized in that, The determination of the disturbance generated by motor failure in each control channel based on the time-varying speed curve includes: The control gain on each control channel is determined based on the maximum speed of the motor, the mass of the aircraft, the lift coefficient of the propeller, the distance between the two propellers on the same side, and the moment of inertia. The disturbance amount generated by motor failure in each control channel is determined based on the control gain and the time-varying speed curve.

5. The fault-tolerant control method for an aircraft as described in claim 1, characterized in that, Before compensating the control input value based on the extended state observer to generate the target control quantity, the method further includes: Establish system models for each control channel; Based on the angular acceleration limit value, model predictive control is performed on the system model of each control channel to generate control input values ​​that meet the constraints.

6. The fault-tolerant control method for an aircraft as described in claim 5, characterized in that, The step of performing model predictive control on the system model of each control channel based on the angular acceleration limit value to generate control input values ​​that satisfy the constraints includes: The optimization objective is to minimize the sum of the squared angular error, the squared angular rate error, and the squared angular acceleration error. Based on the angular acceleration limit value, model predictive control is performed on the system model of each control channel to generate control input values ​​that meet the constraints.

7. The fault-tolerant control method for an aircraft as described in any one of claims 1-6, characterized in that, The fitting of the speed transition process during motor failure to obtain a speed curve changing over time includes: Get the current speed of the motor when it fails; The rotational speed curve is obtained by fitting the second-order linear transfer function based on the current rotational speed.

8. A fault-tolerant control device for an aircraft, characterized in that, The fault-tolerant control device of the aircraft includes: The fitting module is used to fit the speed transition process when the motor fails, and obtain the speed curve that changes over time. The speed transition process is the process in which the motor speed gradually changes from the current speed at the time of failure to zero. The determination module is used to determine the disturbance amount generated by the motor failure in each control channel based on the time-varying speed curve; An observer establishment module is used to establish an extended state observer based on the disturbance amount corresponding to each control channel. The compensation module is used to compensate the control input value based on the extended state observer and generate the target control quantity; The control module is used to control and allocate the target control quantity according to the control allocation matrix.

9. An aircraft, characterized in that, The aircraft includes: a memory, a processor, and a fault-tolerant control program for the aircraft stored in the memory and executable on the processor, the fault-tolerant control program being configured to implement the fault-tolerant control method for the aircraft as described in any one of claims 1 to 7.

10. A storage medium, characterized in that, The storage medium stores a fault-tolerant control program for the aircraft, which, when executed by a processor, implements the fault-tolerant control method for the aircraft as described in any one of claims 1 to 7.

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