Face symmetry rocket closed-loop fault detection method based on gap measurement technology

Abstract

The invention discloses a plane symmetry rocket closed-loop fault detection method based on a gap measurement technology, and belongs to the field of automatic control. According to the method, in the whole fault detection process, complexity caused by a system closed-loop control law and uncertain interference of a system model is considered. Firstly, a co-prime decomposition method is introduced into the first step of the detection process to obtain left co-prime decomposition of a linearized model of a system at a state working point; then, a closed-loop fault detection model of the plane-symmetric rocket servo mechanism is established based on left co-prime decomposition and gap measurement of the system, and fault detection signals on three channels are generated; moreover, thresholds of fault detection signals on three channels are obtained based on the left co-prime decomposition and control law of the system, and the thresholds consider the influence brought by the closed-loop control law of the system. According to the method, the problem that a fault signal is covered by a closed-loop control law can be well solved, so that the accuracy of a fault detection system is improved.

Description

Technical Field

[0001] The present invention relates to the technical field of automatic control, and particularly relates to a closed-loop fault detection method for a face-symmetric rocket based on gap measurement technology. Background Art

[0002] A face-symmetric launch vehicle is a rocket that has symmetry in a specific plane or a certain direction, and its design is different from that of traditional axisymmetric rockets. The unique structural layout of the face-symmetric rocket enables it to flexibly adapt to diverse mission requirements, such as uneven payload distribution or carrying multiple functional modules. However, the incomplete symmetry design also makes the rocket more vulnerable to aerodynamic interference during flight, increasing the complexity of attitude control and the possibility of faults. The actuators of the face-symmetric launch vehicle use two-stage, six engines (core stage and boosters) with bidirectional swing to provide three-channel control torque. The particularity of the face-symmetric rocket's structure places higher requirements on actuator fault detection and fault-tolerant control. Therefore, how to efficiently detect faults in the face-symmetric rocket actuators and ensure its reliable operation through a fault-tolerant control mechanism is a key issue in current research.

[0003] Currently, there have been some research results on fault detection technologies for closed-loop control systems. In reality, most control systems are closed-loop control systems. Detecting faults without considering the control law often brings difficulties to fault detection. For the detection of servo faults in face-symmetric rockets, the above methods have not been applied. Moreover, the attitude control system of face-symmetric rockets is a closed-loop control system, and the existing fault detection methods have not considered the detection method based on coprime factorization and gap measurement technology. Summary of the Invention

[0004] The present invention provides a closed-loop fault detection method for a face-symmetric rocket based on gap measurement technology, which considers the influence of the closed-loop control law on fault detection and can realize three-channel fault detection of the face-symmetric rocket under closed-loop control.

[0005] An embodiment of the present invention provides a closed-loop fault detection method for a face-symmetric rocket based on gap measurement technology, including the following steps:

[0006] Step 1, according to the characteristics of the bundled launch vehicle using the combined swing of the core stage and booster engines for attitude control, establish the kinematic and dynamic models of the face-symmetric rocket during the ascent stage;

[0007] Step 2, based on the kinematic and dynamic models of the face-symmetric rocket during the ascent stage, adopt an equal-proportion distribution of the swing control commands between the core stage and the boosters according to their respective maximum swing angle limits;

[0008] Step 3: According to the control allocation relationship of the swing angle control instruction between the core stage and the booster, convert the kinematic and dynamic model of the face-symmetric rocket in the ascending stage into an input-output model of the face-symmetric rocket with three-channel control swing angle instructions as the input and the attitude angle of the face-symmetric rocket as the output;

[0009] Step 4: Perform working point linearization at the state working point of the input-output model of the face-symmetric rocket to obtain the linearized model at the state working point;

[0010] Step 5: Considering the system model uncertainty existing after the linearization of the face-symmetric rocket at the state working point, perform a coprime factorization on the linearized model to obtain the left coprime factorization of the nominal model;

[0011] Step 6: According to the left coprime factorization of the nominal model, based on the gap metric technique, establish a closed-loop fault detection system model for the face-symmetric rocket against servo mechanism faults to obtain fault detection signals on three channels;

[0012] Step 7: For the fault detection signals on the three channels, determine the fault detection signal thresholds on the three channels in the closed-loop fault detection system model of the face-symmetric rocket according to the left coprime factorization and the closed-loop control law, and judge whether there is a fault in the actuator on the channel according to the fault detection signals and the fault detection signal thresholds.

[0013] Optionally, in an embodiment of the present invention, in Step 1, the kinematic and dynamic model of the face-symmetric rocket in the ascending stage is:

[0014]

[0015] where, represents the state variables, which are respectively the pitch angle, yaw angle, roll angle, pitch rate, yaw rate, and roll rate of the rocket; represents the system output; is the combined swing angle equivalent to the core stage and the booster on the three channels, representing the system input; d1, d 3xj , d 3zt are the system parameters before the state variables and control variables in the face-symmetric rocket system model, d3″ xj , d3″ zt are the system parameters before the second-order derivative of the control variable in the face-symmetric rocket system model; α, α w , β, β w are respectively the angle of attack, additional angle of attack, sideslip angle, and additional sideslip angle; represents the interference torque on the three channels; the dots on the state variable and control variable parameters represent the first derivative and the second derivative.

[0016] Optionally, in an embodiment of the present invention, in step 2, the swing angle control instruction between the core stage and the booster is:

[0017]

[0018] k1 = k2 = 0.5

[0019] where, respectively represent the swing angle control instructions of the core stage in the pitch, yaw, and roll channels; respectively represent the swing angle control instructions of the booster in the pitch, yaw, and roll channels; respectively represent the swing angle instructions output by the basic controllers in the pitch, yaw, and roll channels, and k1 and k2 are the swing angle distribution coefficients of the core stage and the booster respectively.

[0020] Optionally, in an embodiment of the present invention, in step 3, the input-output model of the face-symmetric rocket is:

[0021]

[0022] The non-linear system representation form of the input-output model of the face-symmetric rocket is:

[0023] x = f(x, u, t)

[0024] y = g(x, u, t)

[0025] where, is the equivalent synthetic swing angle obtained according to the control allocation relationship on the three channels, representing the system input.

[0026] Optionally, in an embodiment of the present invention, in step 4, the state-space model described by A, B, C, D is used as the linearized model of the input-output model of the face-symmetric rocket at the state operating point:

[0027]

[0028] Δy = CΔx + DΔu

[0029]

[0030] where, x0 and u0 are the state operating points of the input-output model of the face-symmetric rocket, Δx = x - x0, Δy = y - x0, and Δu = u - u0.

[0031] Optionally, in an embodiment of the present invention, in step 5, considering the system model uncertainty existing after the face-symmetric rocket is linearized at the state operating point, the linearized model is factorized into left and right relatively prime factors to obtain the left relatively prime factorization of the nominal model as:

[0032] G0(s) = Ml0(s) -1 Nl0(s)

[0033]

[0034] Ml(s) = (A Ml , B Ml , C Ml , D Ml )

[0035] Nl(s) = (A Nl , B Nl , C Nl , D Nl )

[0036] where Ml(s) -1 Nl(s) is the left - coprime factorization of the nominal model G(s), satisfying Ml(s)Ml -1 (s) + Nl(s)Nl -1 (s) = I, where I is the identity transfer function matrix, and Ml0(s) -1 Nl0(s) is the left - coprime factorization of the linearized model G0(s), satisfying Ml0(s)Ml0 -1 (s) + Nl0(s)Nl0 -1 (s) = I, and are parameters describing the uncertainty of the system model. (A Ml , B Ml , C Ml , D Ml ) is the state - space model of the left - coprime factorization sub - model Ml(s), and (A Nl , B Nl , C Nl , D Nl ) is the state - space model of the left - coprime factorization sub - model Nl(s).

[0037] Optionally, in an embodiment of the present invention, in step 6, the fault detection signals on the three channels are:

[0038]

[0039] Optionally, in an embodiment of the present invention, in step 7, the fault detection signal threshold on the three channels is J th :

[0040]

[0041] Among them, K is the transfer function represented by the system control law, G is the nominal model, J(K, G) is the transfer function matrix with respect to G and K, v is the system closed-loop control command, used to describe the system model uncertainty, δ Δ is the parameter for describing the system model uncertainty, || || ∞ represents the infinity norm, is the two-norm.

[0042] The closed-loop fault detection method for a face-symmetric rocket based on the gap metric technology in the embodiment of the present invention generates a fault detection signal based on the coprime factorization and gap metric technology. The fault detection signal characterizes the distance between the nominal system of the face-symmetric rocket and the faulty system, and can effectively detect early or minor faults. The present invention generates a fault detection signal threshold based on the coprime factorization and gap metric technology. This threshold takes into account the system model uncertainty and the system closed-loop control law, and can better handle the problem that the closed-loop system control law masks the fault signal.

[0043] Additional aspects and advantages of the present invention will be given in part in the following description, become apparent in part from the following description, or be learned through the practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] The above and / or additional aspects and advantages of the present invention will become apparent and be readily understood from the following description of the embodiments in conjunction with the drawings, wherein:

[0045] Figure 1 is a flowchart of a closed-loop fault detection method for a face-symmetric rocket based on the gap metric technology according to an embodiment of the present invention;

[0046] Figure 2 is a framework diagram of the closed-loop fault detection method for a face-symmetric rocket based on the gap metric technology in an embodiment of the present invention;

[0047] Figure 3 is a diagram of the command value and actual output value of the control system in an embodiment of the present invention when the No. 1 servo of the core stage is stuck;

[0048] Figure 4 is a diagram of the fault signal detection value and threshold of the three channels of the control system in an embodiment of the present invention without a stuck fault;

[0049] Figure 5 is a diagram of the fault signal detection value and threshold of the control system in an embodiment of the present invention when the No. 1 servo of the core stage is stuck. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0050] Embodiments of the present invention will be described in detail below. Examples of the embodiments are shown in the accompanying drawings, where the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The embodiments described below by referring to the accompanying drawings are exemplary and are intended to explain the present invention, and should not be construed as limiting the present invention.

[0051] As Figure 1 and Figure 2 shown, the face-symmetric rocket closed-loop fault detection method based on the gap measurement technology includes the following steps:

[0052] Step 1: According to the characteristics of the bundled launch vehicle using the combined swing of the core stage and booster engines for attitude control, establish the kinematic and dynamic models of the face-symmetric rocket in the ascending stage.

[0053] Optionally, in an embodiment of the present invention, in Step 1, the kinematic and dynamic models of the face-symmetric rocket in the ascending stage are as follows:

[0054]

[0055] Among them, represents the state variables, which are the pitch angle, yaw angle, roll angle, pitch rate, yaw rate, and roll rate of the rocket respectively; represents the system output; is the synthetic swing angle equivalent to the core stage and booster on the three channels, representing the system input; d1, d 3xj , d 3zt are the system parameters before the state variables and control variables in the face-symmetric rocket system model, d3″ xj , d3″ zt are the system parameters before the second-order derivative of the control variable in the face-symmetric rocket system model; α, α w , β, β w are the angle of attack, additional angle of attack, sideslip angle, and additional sideslip angle respectively; represents the interference torque on the three channels; the dots on the state variable and control variable parameters represent the first-order derivative and second-order derivative.

[0056] The servo mechanism limit considers a single unit limit of 6°, a rate limit of 10° / s, and an acceleration limit of 10 rad / s. The transfer function of the actuator is a standard second-order transfer function, and its natural frequency and damping ratio are [32 0.3].

[0057] Step 2: Based on the kinematic and dynamic models of the face-symmetric rocket in the ascending stage, adopt an equal-proportion distribution of the swing angle control commands between the core stage and the booster according to their respective maximum swing angle limits.

[0058] Optionally, in an embodiment of the present invention, in step 2, the swing angle control instructions between the core stage and the booster are:

[0059]

[0060] k1 = k2 = 0.5

[0061] wherein, respectively represent the swing angle control instructions of the core stage in the pitch, yaw, and roll channels; respectively represent the swing angle control instructions of the booster in the pitch, yaw, and roll channels; respectively represent the swing angle instructions output by the basic controllers in the pitch, yaw, and roll channels, and k1 and k2 are the swing angle distribution coefficients of the core stage and the booster respectively.

[0062] Step 3: According to the control allocation relationship of the swing angle control instructions between the core stage and the booster, convert the kinematic and dynamic model of the symmetric rocket in the ascending stage into an input-output model of the symmetric rocket with three-channel control swing angle instructions as the input and the attitude angle of the symmetric rocket as the output.

[0063] Optionally, in an embodiment of the present invention, in step 3, the input-output model of the symmetric rocket is:

[0064]

[0065] The non-linear system representation form of the input-output model of the symmetric rocket is:

[0066] x = f(x, u, t)

[0067] y = g(x, u, t)

[0068] wherein, is the equivalent synthetic swing angle obtained according to the control allocation relationship on the three channels, representing the system input.

[0069] Step 4: Perform working point linearization at the state working point of the input-output model of the symmetric rocket to obtain the linearized model at the state working point.

[0070] Optionally, in an embodiment of the present invention, in step 4, the linearized model G0(s) of the input-output model of the symmetric rocket at the state working point, described by A, B, C, and D, is:

[0071]

[0072] Δy = CΔx + DΔu

[0073]

[0074] where x0 = [1.5135 0.0586 0.0509 0 0 0] T and u0 = [-0.0166 0.0166 0.0144] T are the state operating points of the input-output model of the axisymmetric rocket, Δx = x - x0, Δy = y - x0, and Δu = u - u0.

[0075] Step 5: Considering the system model uncertainty existing after the axisymmetric rocket is linearized at the state operating point, perform a coprime factorization on the linearized model to obtain the left coprime factorization of the nominal model.

[0076] In the embodiment of the present application, the system refers to the system after the axisymmetric rocket is linearized at the state operating point, and it has uncertainty.

[0077] Optionally, in an embodiment of the present invention, in Step 5, considering the system model uncertainty existing after the axisymmetric rocket is linearized at the state operating point, perform a coprime factorization on the linearized model to obtain the left coprime factorization of the nominal model as:

[0078] G0(s) = Ml0(s) -1 Nl0(s)

[0079]

[0080] Ml(s) = (A Ml , B Ml , C Ml , D Ml )

[0081] Nl(s) = (A Nl , B Nl , C Nl , D Nl )

[0082] where Ml(s) -1 Nl(s) is the left coprime factorization of the nominal model G(s), satisfying Ml(s)Ml -1 (s) + Nl(s)Nl -1 (s) = I, I is the unit transfer function matrix, Ml0(s) -1 Nl0(s) is the left coprime factorization of the linearized model G0(s), satisfying Ml0(s)Ml0 -1 (s) + Nl0(s)Nl0 -1 (s) = I, and are the parameters describing the system model uncertainty, (A Ml , B Ml , C Ml , DMl ) is the state - space model of the left - relatively - prime factorization sub - model \(M_l(s)\), \((A Nl , B Nl , C Nl , D Nl ) is the state - space model of the left - relatively - prime factorization sub - model \(N_l(s)\).

[0083] Step 6: Based on the left - relatively - prime factorization of the nominal model and the gap metric technique, establish a closed - loop fault - detection system model for the symmetric rocket facing servo - mechanism faults, and obtain the fault - detection signals on the three channels.

[0084] Optionally, in an embodiment of the present invention, in Step 6, the fault - detection signals on the three channels are:

[0085]

[0086] Step 7: For the fault - detection signals on the three channels, according to the left - relatively - prime factorization and the closed - loop control law, determine the fault - detection signal thresholds on the three channels in the closed - loop fault - detection system model of the symmetric rocket, and judge whether the actuators on the channels are faulty according to the fault - detection signals and the fault - detection signal thresholds.

[0087] In the embodiment of the present invention, first, the left - relatively - prime factorization of the linearized model at the operating point is obtained. Then, based on the left - relatively - prime factorization and the gap metric technique, a closed - loop fault - detection system model for the symmetric rocket is established. According to the left - relatively - prime factorization and the closed - loop control law, the fault - detection signal thresholds are obtained. When a fault occurs in the system actuator, the fault - detection signal exceeds the threshold, and it is detected that the actuator on the channel has a fault.

[0088] Optionally, in an embodiment of the present invention, in Step 7, the fault - detection signal thresholds on the three channels are \(J th :

[0089]

[0090] where \(K\) is the transfer function represented by the system control law, \(G\) is the nominal model, \(J(K,G)\) is the transfer - function matrix with respect to \(G\) and \(K\), \(v\) is the system closed - loop control command, used to describe the system model uncertainty, \(\delta Δ is the parameter describing the system model uncertainty, \(\|\cdot\|\) ∞ represents the infinity norm, is the two - norm.

[0091] Next, the effectiveness of the present invention is verified by simulation through the accompanying drawings and specific embodiments. The simulation parameters are as follows:

[0092] Symmetric rocket model parameters:

[0093] Initial state value: x0 = [1.5135 0.0586 0.0509 0 0 0] T , u0 = [-0.0166 0.0166 0.0144] T , and the controller gain is: (The same for three channels), and the reference signal is v = [1.5707 0 0] T .

[0094] Consider a possible servo jamming fault model that may be suffered.

[0095] Fault 1: A jamming fault occurs in the core stage servo at t = 40 s, and the jamming angle is -2°.

[0096]

[0097] Among them, is the command value of the core stage No. 1 servo mechanism, is the actual value of the core stage No. 1 servo mechanism.

[0098] Determine the threshold: According to the calculation formula, select the parameter δ used to describe the uncertainty of the system model Δ = 0.05, and calculate to get J th = [0.0033 0.0266 0.0042].

[0099] The jamming fault is as Figure 3 shown. For the symmetric rocket, a jamming fault occurs in the core stage No. 1 servo at 40 s, and the jamming angle is -2°.

[0100] The fault signal detection diagram under the fault-free condition is as Figure 4 shown. It can be seen that when there is no jamming fault, the fault signal detection values of the three channels do not exceed the threshold.

[0101] The fault signal detection diagram under the faulty condition is as Figure 5 shown. It can be seen that for the symmetric rocket, a jamming fault occurs in the core stage No. 1 servo at 40 s. According to the servo synthesis relationship, the core stage No. 1 servo affects the pitch and roll channels and has no effect on the yaw channel. From the simulation result diagram, the fault detection signal values of the pitch and roll channels exceed the threshold, and it is detected that there is a servo fault on this channel. The simulation results verify the effectiveness of the proposed fault detection strategy.

[0102] The closed-loop fault detection method for face-symmetric rockets based on gap metric technology in the embodiments of the present invention is directed to the six-degree-of-freedom model of face-symmetric rockets. A method based on closed-loop fault detection technology is proposed under the premise of considering the closed-loop control law of the system to achieve three-channel fault detection for face-symmetric rockets. In the whole fault detection process, the complexity brought by the closed-loop control law of the system and the uncertain interference of the system model is considered. First, the right-coprime factorization method is introduced in the first step of the detection process to obtain the right-coprime factorization of the linearized model of the system at the state operating point. Then, based on the right-coprime factorization of the system and the gap metric, a closed-loop fault detection model of the servo mechanism of the face-symmetric rocket is established to generate fault detection signals on three channels. Moreover, based on the right-coprime factorization of the system and the control law, the thresholds of the fault detection signals on three channels are obtained, and this threshold considers the influence brought by the closed-loop control law of the system. Finally, for the stuck fault of the servo mechanism of the face-symmetric rocket. The present invention can better handle the problem of masking of fault signals due to the closed-loop control law, thereby improving the accuracy of the fault detection system.

[0103] In the description of this specification, the descriptions referring to the terms "one embodiment", "some embodiments", "example", "specific example", or "some examples", etc. mean that the specific features, structures, materials, or characteristics described in connection with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials, or characteristics described can be combined in any one or N embodiments or examples in a suitable manner. In addition, without contradiction, those skilled in the art can combine and combine the different embodiments or examples described in this specification and the features of different embodiments or examples.

[0104] In addition, the terms "first" and "second" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly specifying the quantity of the indicated technical features. Thus, the features defined with "first" and "second" can explicitly or implicitly include at least one of these features. In the description of the present invention, the meaning of "N" is at least two, such as two, three, etc., unless otherwise specifically defined.

[0105] Any process or method description shown in the flowchart or described in other ways herein can be understood as representing a module, segment, or part of code including one or N executable instructions for implementing a customized logical function or process. The scope of the preferred embodiments of the present invention includes additional implementations, where the functions can be executed in a substantially simultaneous manner or in the reverse order according to the involved functions, rather than in the order shown or discussed, which should be understood by those skilled in the art of the embodiments of the present invention.

Claims

1. A method for detecting closed-loop faults of a plane-symmetric rocket based on gap measurement technology, characterized in that: The following steps are involved: Step 1: Based on the characteristics of the strapped launch vehicle using the core stage and booster engine to perform attitude control by joint swinging, a kinematic and dynamic model of the ascent phase of the plane-symmetric rocket is established; Step 2, based on the kinematics and dynamics model of the ascent phase of the plane-symmetric rocket, the swing angle control instructions between the core stage and the booster are distributed in equal proportion according to the respective maximum swing angle limits; Step 3, according to the control allocation relationship of the swing angle control instructions between the core stage and the booster, the kinematic and dynamic model of the ascent stage of the plane-symmetric rocket is converted into a plane-symmetric rocket input-output model whose input is a three-channel control swing angle instruction and whose output is a plane-symmetric rocket attitude angle; Step 4, performing operating point linearization at the state operating point of the plane-symmetric rocket input-output model to obtain a linearized model at the state operating point; Step 5, considering the system model uncertainty existing after the linearization of the plane-symmetric rocket at the state operating point, the linearized model is coprime decomposed to obtain the left coprime decomposition of the nominal model; Step 6, according to the left coprime decomposition of the nominal model, based on the gap measurement technology, for the servo mechanism fault, a plane-symmetric rocket closed-loop fault detection system model is established to obtain the fault detection signals on the three channels; Step 7, for the fault detection signals on the three channels, determine the fault detection signal thresholds on the three channels in the plane-symmetric rocket closed-loop fault detection system model based on the left coprime decomposition and the closed-loop control law, and determine whether a fault occurs in the actuator on the channel based on the fault detection signal and the fault detection signal threshold.

2. The method according to claim 1, characterized in that In step 1, the kinematic and dynamic model of the ascent phase of the plane-symmetric rocket is: in, Represents the state quantity, which are the rocket's pitch angle, yaw angle, roll angle, pitch angle rate, yaw angle rate, and roll angle rate; Indicates system output; is the synthetic swing angle equivalent to the core stage and booster on the three channels, indicating the system input; d1,d 3xj ,d 3zt is the system parameter before the state and control quantities in the symmetric rocket system model, are the system parameters before the second-order derivative of the control quantity in the symmetric rocket system model; α, α w ,β,β w They are wind attack angle, additional wind attack angle, sideslip angle, and additional sideslip angle; Represents the disturbance torque on the three channels; the points on the state quantity and control quantity parameters represent the first-order derivative and the second-order derivative.

3. The method according to claim 2, characterized in that In step 2, the swing angle control instruction between the core stage and the booster is: in, Respectively represent the core stage swing angle control instructions of the pitch, yaw and roll channels; Respectively represent the pitch, yaw, and roll channel thrust swing angle control instructions; They represent the swing angle commands output by the basic controllers of the pitch, yaw and roll channels respectively; k1 and k2 are the core stage and booster swing angle allocation coefficients respectively.

4. The method according to claim 2, characterized in that: In step 3, the plane-symmetric rocket input-output model is: The nonlinear system expression of the plane-symmetric rocket input-output model is: x=f(x,u,t) y=g(x,u,t) in, It is the equivalent synthetic swing angle obtained on the three channels according to the control distribution relationship, which represents the system input.

5. The method according to claim 4, characterized in that In step 4, the state space model described by A, B, C, and D is used as the linearized model of the plane symmetric rocket input-output model at the state operating point: Among them, x0 and u0 are the state operating points of the plane-symmetric rocket input-output model, Δx=x-x0, Δy=y-x0, Δu=u-u0.

6. The method according to claim 5, characterized in that In step 5, considering the system model uncertainty of the plane-symmetric rocket after linearization at the state operating point, the linearized model is coprime decomposed to obtain the left coprime decomposition of the nominal model: G0(s)=Ml0(s) -1 Nl0(s) Ml(s)=(A Ml ,B Ml ,C Ml ,D Ml ) Nl(s)=(A Nl ,B Nl ,C Nl ,D Nl ) Among them, Ml(s) -1 Nl(s) is the left coprime decomposition of the nominal model G(s), satisfying Ml(s)Ml -1 (s)+Nl(s)Nl -1 (s) = I, I is the unit transfer function matrix, Ml0(s) -1 Nl0(s) is the left coprime decomposition of the linearized model G0(s), satisfying Ml0(s)Ml0 -1 (s)+Nl0(s)Nl0 -1 (s) = I, and is the parameter describing the uncertainty of the system model, (A Ml ,B Ml ,C Ml ,D Ml ) is the state space model of the left coprime decomposition submodel Ml(s), (A Nl ,B Nl ,C Nl ,D Nl ) is the state space model of the left coprime decomposition submodel Nl(s).

7. The method according to claim 6, characterized in that In step 6, the fault detection signals on the three channels are:

8. The method according to claim 7, characterized in that In step 7, the fault detection signal threshold on the three channels is J th : Where K is the transfer function represented by the system control law, G is the nominal model, J(K,G) is the transfer function matrix about G and K, and v is the closed-loop control instruction of the system. Used to describe the system model uncertainty, δ Δ To describe the uncertainty parameters of the system model, || || ∞ represents the infinite norm, is the two-norm.

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