CN111679580A - Self-adaptive aircraft control system fault compensation and disturbance suppression method - Google Patents

Self-adaptive aircraft control system fault compensation and disturbance suppression method Download PDF

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CN111679580A
CN111679580A CN202010527352.8A CN202010527352A CN111679580A CN 111679580 A CN111679580 A CN 111679580A CN 202010527352 A CN202010527352 A CN 202010527352A CN 111679580 A CN111679580 A CN 111679580A
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fault
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interference
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CN111679580B (en
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姚雪莲
杨艺
朱燕
常瀚文
吴梦平
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Jiangsu University of Technology
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Abstract

The invention provides a self-adaptive aircraft control system fault compensation and disturbance suppression method, which comprises the following steps: acquiring interference items and fault items influencing an aircraft control system; respectively establishing an interference model and a fault model according to the interference item and the fault item; establishing a dynamic model influencing the input of the aircraft control system according to the interference model and the fault model; designing a basic controller according to the dynamic model; respectively designing adaptive controllers based on the failure modes; and designing the integrated controller based on the self-adaptive weighted fusion. The method can effectively inhibit gust disturbance and compensate faults when the gust disturbance and unknown actuator faults occur in the aircraft control system, so that the aircraft control system is ensured to have expected closed loop stability and output tracking performance, targets can be controlled stably and gradually in a tracking manner, the fault-tolerant control effect is improved, and the flight safety of the aircraft is ensured.

Description

Self-adaptive aircraft control system fault compensation and disturbance suppression method
Technical Field
The invention relates to the technical field of automatic control, in particular to a fault compensation and disturbance suppression method for an aircraft control system based on self-adaptation.
Background
With the development of science and technology, some novel application systems with performance key of large-scale system dynamics, such as aerospace vehicles, nuclear reactors, micro-electro-mechanical systems and smart grid systems, appear. The novel systems have the characteristics of large working change range, complex system dynamic range, more elements and the like, so that the novel systems are extremely easy to be influenced by external environment. The influence of an uncertain external environment on the system can be characterized by a plurality of uncertain external disturbances, such as gusts, turbulence, freezing and other adverse airborne conditions in the aerospace field. These external disturbances, which are not deterministic, can often cause uncertainties in the system itself, and in particular if the system still experiences actuator failures, it will further degrade the desired system performance and even cause system instability, thereby inducing catastrophic failures. With the rapid development of performance key systems represented by aircrafts, system models are more complex, and the characteristics of nonlinearity, multivariable, strong coupling, uncertainty and the like are more prominent, so that the existing control theory and technology cannot meet the requirements of safe flight of the aircrafts, and new control theory and method are urgently needed to improve the safety and reliability of the control system, thereby effectively avoiding catastrophic accidents and major economic loss.
Despite the numerous and feasible advances made in actuator fault compensation research for flight systems, there are still a number of open-ended problems, especially control problems not considering the simultaneous presence of unknown mismatch disturbances and multiple uncertain actuator faults in practice.
Disclosure of Invention
The present invention is directed to solving, at least to some extent, one of the technical problems in the art described above. Therefore, the invention aims to provide a fault compensation and disturbance suppression method for an aircraft control system based on self-adaptation, which can effectively suppress gust disturbance and compensation faults when the aircraft control system has gust disturbance and unknown actuator faults, so that the aircraft control system is ensured to have expected closed-loop stability and output tracking performance, a target can be controlled in a stable and gradual tracking manner, the fault-tolerant control effect is improved, and the flight safety of an aircraft is ensured.
In order to achieve the above object, an embodiment of the present invention provides an adaptive-based aircraft control system fault compensation and disturbance suppression method, including the following steps: acquiring interference items and fault items influencing an aircraft control system; respectively establishing an interference model and a fault model according to the interference item and the fault item; establishing a dynamic model influencing the input of the aircraft control system according to the interference model and the fault model; designing a base controller from the dynamic model, wherein the base controller is configured to suppress an effect of interference on the aircraft control system output; respectively designing adaptive controllers based on the failure modes, wherein the adaptive controllers are used for compensating the influence of the failure on the output of the aircraft control system; and designing a comprehensive controller based on the self-adaptive weighted fusion, wherein the comprehensive controller is used for compensating faults and inhibiting the influence of interference on the performance of the aircraft control system.
According to the fault compensation and disturbance suppression method for the aircraft control system based on the self-adaptation provided by the embodiment of the invention, the interference model and the fault model are respectively established based on the interference item and the fault item, the dynamic model which influences the input of the aircraft control system is further established according to the interference model and the fault model, the uncertain influence of the interference and the fault on the input of the aircraft control system is expressed in a mathematical model mode, then the basic controller can be designed according to the dynamic model to suppress the influence of the interference on the output of the aircraft control system, meanwhile, the self-adaptation controllers can be respectively designed based on the fault model to compensate the influence of the fault on the output of the aircraft control system, and finally, the integrated controller can be designed based on the self-adaptation weighting fusion to compensate the influence of the fault and the disturbance on the performance of the aircraft control system, therefore, when the aircraft control system has gust disturbance and unknown, the method effectively inhibits gust disturbance and compensates faults, thereby ensuring that an aircraft control system has expected closed loop stability and output tracking performance, realizing stable and progressive tracking control of targets, improving fault-tolerant control effect and ensuring flight safety of the aircraft.
In addition, the fault compensation and disturbance suppression method based on the adaptive aircraft control system proposed according to the above embodiment of the present invention may also have the following additional technical features:
further, the method for fault compensation and disturbance suppression of the self-adaptive aircraft control system further comprises the step of verifying the integrated controller by using a Lyapunov function.
Further, the interference term is gust disturbance, and the fault term is uncertain actuator faults of the aircraft.
Further, the interference model is:
Figure BDA0002534041320000031
wherein the content of the first and second substances,
Figure BDA0002534041320000032
in order for the interference parameters to be unknown,
Figure BDA0002534041320000033
is a known basis function of the known functions of, among others,
Figure BDA0002534041320000034
and
Figure BDA0002534041320000035
respectively as follows:
Figure BDA0002534041320000036
wherein j is 1,2, 3.
Further, the fault model is:
Figure BDA0002534041320000037
where σ (t) is the corresponding actuator failure mode matrix, v (t) is the control input signal to be designed,
Figure BDA0002534041320000038
is a fault parameter vector, where σ (t), v (t), and
Figure BDA0002534041320000039
respectively as follows:
σ(t)＝diag{σ1(t),σ2(t),σ3(t),σ4(t)}，
v(t)＝[v1(t),v2(t),v3(t),v4(t)]T，
Figure BDA00025340413200000310
further, the dynamic model is:
Figure BDA0002534041320000041
where V is aircraft speed, α is angle of attack, θ is pitch angle, q is pitch angle rate, m is mass, IyIs moment of inertia, M is pitching moment, d1，d2And d3Is a turbulent perturbation signal.
Further, designing a base controller according to the dynamic model includes: estimating interference parameters according to the interference model; obtaining an interference design self-adaptive linear control law according to the estimation result; setting an interference parameter self-adaptive law; setting a Lyapunov function; and verifying the parameter self-adaption law according to the Lyapunov function and the interference parameter self-adaption law.
Further, the failure modes include an actuator no failure mode and an actuator failure mode, wherein a first adaptive controller is designed based on the actuator no failure mode, and a second adaptive controller is designed based on the actuator failure mode.
Further, the design comprehensive controller based on the adaptive weighted fusion comprises: setting a fault indication function; and obtaining an integrated controller according to the fault indication function, the first self-adaptive controller and the second self-adaptive controller.
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FIG. 1 is a flow chart of a method for fault compensation and disturbance rejection for an adaptive-based aircraft control system in accordance with an embodiment of the present invention;
FIG. 2 is a flow chart of a method for fault compensation and disturbance rejection for an adaptive-based aircraft control system according to an embodiment of the invention;
FIG. 3 is a schematic illustration of a comparison between reference signals corresponding to actual outputs of an aircraft control system in accordance with one embodiment of the present invention;
FIG. 4 is a schematic illustration of a tracking error of an aircraft control system in accordance with an embodiment of the present invention;
FIG. 5 is a schematic representation of control input signals applied to a control system by an aircraft actuator according to one embodiment of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
FIG. 1 is a method for fault compensation and disturbance suppression of an adaptive-based aircraft control system according to an embodiment of the present invention, including the following steps:
and S1, acquiring interference items and fault items influencing the aircraft control system.
And S2, respectively constructing an interference model and a fault model according to the interference item and the fault item.
In one embodiment of the present invention, interference terms and fault terms affecting an aircraft control system may be collected by a sensor, where the interference terms, i.e., interference signal d (t), do not match a control signal u (t) of the aircraft control system, and an interference vector of the interference terms, i.e., interference signal d (t), may be:
d(t)＝[d1(t),d2(t),d3(t)]3∈R3。
further, the components of the interference vector may be:
Figure BDA0002534041320000051
wherein d isj0And djkUnknown parameter, fjk(t) known parameters.
Further, the interference model can be characterized by a parameterized model:
Figure BDA0002534041320000052
wherein the content of the first and second substances,
Figure BDA0002534041320000053
in order for the interference parameters to be unknown,
Figure BDA0002534041320000054
is a known basis function of the known functions of, among others,
Figure BDA0002534041320000055
and
Figure BDA0002534041320000056
respectively as follows:
Figure BDA0002534041320000061
Figure BDA0002534041320000062
wherein j is 1,2, 3.
In one embodiment of the invention, an actuator fault function σ may be introducedj(t) and an indicator index j equal to 1,2,3,4, where σ is the case if the jth actuator failsj(t) 1, otherwise σj(t)＝0。
Further, the fault model may be characterized by a parameterized model:
Figure BDA0002534041320000063
wherein j ∈ {1,2,3,4} is an indeterminate fault index, tj>0 is the uncertain fault time and the fault time,
Figure BDA0002534041320000064
and
Figure BDA0002534041320000065
for uncertain fault parameters, fji(t) and qjThe known content of more than or equal to 1,
Figure BDA0002534041320000066
for the unknown fault parameter vector to be,
Figure BDA0002534041320000067
is a known basis function.
Further, an actuator fault function σ may be incorporatedj(t) and indicator index j ═ 1,2,3,4 represent the fault model as:
Figure BDA0002534041320000068
where σ (t) is the corresponding actuator failure mode matrix, v (t) is the control input signal to be designed,
Figure BDA0002534041320000069
is a fault parameter vector.
Specifically, the actuator failure mode matrix is:
σ(t)＝diag{σ1(t),σ2(t),σ3(t),σ4(t)}，
the control input signal v (t) to be designed is:
v(t)＝[v1(t),v2(t),v3(t),v4(t)]T，
fault parameter vector
Figure BDA0002534041320000071
Comprises the following steps:
Figure BDA0002534041320000072
and S3, establishing a dynamic model influencing the input of the aircraft control system according to the interference model and the fault model.
In one embodiment of the invention, the interference model and the fault model may be passed, i.e.
Figure BDA0002534041320000073
And
Figure BDA0002534041320000074
and representing the influence of disturbance items and fault items, namely gust disturbance and uncertain actuator faults on the input of the aircraft control system in a mathematical model manner so as to establish a dynamic model influencing the input of the aircraft control system.
Wherein, the dynamic model is as follows:
Figure BDA0002534041320000075
where V is aircraft speed, α is angle of attack, θ is pitch angle, q is pitch angle rate, m is mass, IyIs moment of inertia, M is pitching moment, d1，d2And d3For turbulent disturbance signals, Fx，FzAnd M is respectively:
Figure BDA0002534041320000076
wherein the content of the first and second substances,
Figure BDA0002534041320000077
for dynamic pressure, ρ is the air density, S is the wing density, c is the average chord, T1And T2As a thruster, Cx，CzAnd CmAre respectively as：
Figure BDA0002534041320000081
Wherein the content of the first and second substances,e1ande2two actuators designed for fault compensation.
Further, the interference model, the fault model, the dynamic model, and F may be based on the abovex，FzAnd equation M, and Cx，CzAnd CmEquality and selecting the state variable x1＝V,x2＝α,x3＝θ,x4Q, input variable u ═ u1,u2,u3,u4]T＝[e1,e2,T1,T2]TThe method can obtain a mathematical model of the fault of the uncertain actuator of the aircraft control system under the gust disturbance, and the state space representation of the mathematical model is as follows:
Figure BDA0002534041320000082
wherein the content of the first and second substances,
Figure BDA0002534041320000083
Figure BDA0002534041320000084
g2(x)＝-a1x1sin(x2)+a2x1cos(x2)，g31(x)＝cos r1cos(x2)+sin r2sin(x2)，
Figure BDA0002534041320000085
g41(x)cosr1cos(x2)+sin r2sin(x2)，
Figure BDA0002534041320000086
Figure BDA0002534041320000087
B,k1,k2,c1,c2,p1,a1,a2,r1,r2,b1,b2is a known constant.
S4, designing a basic controller according to the dynamic model, wherein the basic controller is used for restraining influence of interference on output of the aircraft control system.
Specifically, step S4 includes: estimating interference parameters according to the interference model; obtaining an interference design self-adaptive linear control law according to an estimation result; setting an interference parameter self-adaptive law; setting a Lyapunov function; and verifying the parameter self-adaptive law according to the Lyapunov function and the interference parameter self-adaptive law.
More specifically, the interference model may be estimated as:
Figure BDA0002534041320000091
wherein the content of the first and second substances,
Figure BDA0002534041320000092
as a disturbance parameter, i.e.
Figure BDA0002534041320000093
J-1, 2,3,4 is a known function.
Further, an interference design adaptive linear control law can be obtained according to the estimation result:
Figure BDA0002534041320000094
wherein the content of the first and second substances,
Figure BDA0002534041320000095
is uLIs estimated, and
Figure BDA0002534041320000096
the estimated components of (a) are:
Figure BDA0002534041320000097
Figure BDA0002534041320000098
wherein i is 1,2, 3.
Further, an interference parameter adaptation law, i.e. interference parameters, can be designed
Figure BDA0002534041320000099
The adaptation law of (2):
Figure BDA00025340413200000910
meanwhile, the control equation adaptation law can be designed:
Figure BDA00025340413200000911
further, the lyapunov function can be designed:
Figure BDA00025340413200000912
wherein the content of the first and second substances,djis an adaptive gain matrix, and
Figure BDA0002534041320000101
p is a positive definite symmetric matrix and satisfies the equation
Figure BDA0002534041320000102
Q＝QT＞0。
Further, the derivative may be performed on the designed lyapunov function:
Figure BDA0002534041320000103
wherein Z isP＝[ZP1,ZP2,ZP3,ZP4]T，ZPi,i＝1,2,3,4Is zTP，
Figure BDA0002534041320000104
To (1) a
Figure BDA0002534041320000105
And (4) a component.
Further, the interference parameter may be determined
Figure BDA0002534041320000106
Substituting the self-adaptive law formula into a designed Lyapunov function derivation formula to obtain:
Figure BDA0002534041320000107
further in accordance with
Figure BDA0002534041320000108
Validating parameters of a design
Figure BDA0002534041320000109
The self-adaptive law formula can ensure the stability of the aircraft control system and ensure that the aircraft control system obtains expected performance.
And S5, respectively designing adaptive controllers based on the fault modes, wherein the adaptive controllers are used for compensating the influence of the fault on the output of the aircraft control system.
In one embodiment of the invention, the failure modes may include an actuator no failure mode and an actuator failure mode, wherein the first adaptive controller may be designed based on the actuator no failure mode and the second adaptive controller may be designed based on the actuator failure mode.
Specifically, designing the first adaptive controller based on the actuator failure-free mode includes setting a desired control signal for the aircraft control system, i.e., setting a desired control signal for the aircraft control system
Figure BDA00025340413200001010
Meanwhile, according to the failure-free mode of the actuator, the output signal of the actuator can be obtained as sigma-diag {0,0,0, 0), and further, the fault parameter, i.e., the fault parameter, can be obtained
Figure BDA00025340413200001011
To zero, the resulting fault parameter, i.e.
Figure BDA00025340413200001012
Substituting for zero the control signal desired by the aircraft control system set above, i.e.
Figure BDA0002534041320000111
So that the control signal w can be obtainedd(t)＝Aσ(x)v。
Further, a suitable matrix equation h may be selected21(x) Obtaining:
Figure BDA0002534041320000112
further, the equation may be solved
Figure BDA0002534041320000113
Obtaining a first self-adaptive controller under the actuator failure-free mode:
Figure BDA0002534041320000114
similarly, designing the second adaptive controller based on actuator failure mode includes, from the actuator failure mode, obtaining:
Figure BDA0002534041320000115
A(x)＝[A1,A2,A3,A4]＝[A1,A(2)]∈R12，A(2)＝[A2,A3,A4]∈R9，
Figure BDA0002534041320000116
wherein, i is 2,3,4,.
Further, a suitable matrix equation h may be selected22Obtaining:
Figure BDA0002534041320000117
further, the equation may be solved
Figure BDA0002534041320000118
Obtaining a second self-adaptive controller under the actuator failure mode:
Figure BDA0002534041320000119
and S6, designing the comprehensive controller based on the self-adaptive weighted fusion, wherein the comprehensive controller is used for compensating faults and inhibiting the influence of interference on the performance of the aircraft control system.
Specifically, step S6 includes: setting a fault indication function; obtaining a comprehensive controller according to the fault indication function, the first self-adaptive controller and the second self-adaptive controller; and designing a fault parameter self-adaptive updating law based on a projection algorithm.
More specifically, the fault indication function may be set to:
Figure BDA0002534041320000121
further, the integrated controller may be obtained according to the set fault indication function, the first adaptive controller expression and the second adaptive controller obtained integrated controller expression:
Figure BDA0002534041320000122
wherein the content of the first and second substances,
Figure BDA0002534041320000123
and
Figure BDA0002534041320000124
are respectively as
Figure BDA0002534041320000125
And
Figure BDA0002534041320000126
an estimate of (d).
Further, the comprehensive controller expression can be parameterized to obtain:
Figure BDA0002534041320000127
Figure BDA0002534041320000128
Figure BDA0002534041320000129
wherein, χj,iAnd theta1(i)Are respectively as
Figure BDA00025340413200001210
And
Figure BDA00025340413200001211
is further specified, wherein
Figure BDA00025340413200001212
Further, the above parameters may be aligned based on error and projection algorithms
Figure BDA00025340413200001213
And
Figure BDA00025340413200001214
designing an adaptive updating law:
Figure BDA00025340413200001215
Figure BDA0002534041320000131
Figure BDA0002534041320000132
wherein the content of the first and second substances,
Figure BDA0002534041320000133
γ1i> 0 and gamma2iThe adaptive gain is more than 0, and the adaptive gain is more than 0,
Figure BDA0002534041320000134
is a projection algorithm.
It should be further noted that the method for fault compensation and disturbance suppression of an aircraft control system based on self-adaptation according to the embodiment of the present invention further includes checking the integrated controller by using the lyapunov function.
Specifically, the time period T ∈ [ T ] may be0,T1),T1＝∞,σ＝σ(1)I.e. actuator u, {0,0, 0}, i.e. diag {0,0, 0}1The design method of the integrated controller is checked by using the Lyapunov function under the condition of no fault, and the time period T ∈ [ T [ [ T ]1,T2),T2＝∞,σ＝σ(2)I.e. actuator u, diag {1,0,0,0}1And (3) in case of failure, the design method of the integrated controller is checked by adopting a Lyapunov function.
More specifically, during time period T ∈ [ T0,T1),T1＝∞,σ＝σ(1)I.e. actuator u, {0,0, 0}, i.e. diag {0,0, 0}1The method for testing the design method of the integrated controller by adopting the Lyapunov function under the condition of no fault comprises the following steps:
first, actuator u may be listed1Lyapunov function equation for aircraft control systems in the absence of faults, i.e. when σ ═ σ(1)When the value is "diag {0,0,0,0 }:
Figure BDA0002534041320000135
which in turn may incorporate the above equality interference parameters
Figure BDA0002534041320000136
Adaptive law of (a) for the above parameters based on error and projection algorithms
Figure BDA0002534041320000137
And
Figure BDA0002534041320000138
designing a self-adaptive updating law to conduct derivation on a Lyapunov function equation to obtain:
Figure BDA0002534041320000139
likewise, at time period T ∈ [ T ]1,T2),T2＝∞,σ＝σ(2)I.e. actuator u, diag {1,0,0,0}1The method for testing the design method of the integrated controller by adopting the Lyapunov function under the condition of the fault comprises the following steps:
first, actuator u may be listed1Lyapunov function equation for aircraft control systems in the absence of faults, i.e. when σ ═ σ(2)When 1,0,0,0 ═ diag:
Figure BDA0002534041320000141
which in turn may incorporate the above equality interference parameters
Figure BDA0002534041320000142
Adaptive law of (a) for the above parameters based on error and projection algorithms
Figure BDA0002534041320000143
And
Figure BDA0002534041320000144
design adaptationThe lyapunov equation is derived by applying the update law to obtain:
Figure BDA0002534041320000145
the Lyapunov function equation after derivation can prove that the integrated controller can be used for system stabilization and asymptotic tracking in an aircraft control system.
Based on the above steps, fault compensation and disturbance suppression for the aircraft control system can be realized, and in order to further illustrate the control process of the method for fault compensation and disturbance suppression for the adaptive aircraft control system according to the embodiment of the present invention, the following description will be made with reference to the control process diagram shown in fig. 2.
Specifically, as shown in fig. 2, an interference item affecting the flight control system may be set as gust disturbance, a fault item is multiple uncertain actuator faults, a mathematical model of the aircraft based on the gust disturbance and the multiple uncertain actuator faults may be established, and the inputs of the gust disturbance and the multiple uncertain actuator faults to the aircraft control system are described by the mathematical model, that is, ym(t) and establishing a dynamic model of the fault of the uncertain actuator of the aircraft control system under the gust disturbance; further, a basic controller can be designed according to a dynamic model of the aircraft control system uncertain actuator fault under gust disturbance to inhibit the influence of gust disturbance on the aircraft control system output, namely y (t), and specifically, adaptive parameter estimation can be carried out to obtain an estimation item of an interference parameter
Figure BDA0002534041320000146
And obtaining a self-adaptive controller set, namely a basic controller, a first self-adaptive controller and a second self-adaptive controller, according to the multiple uncertain executor faults so as to compensate the influence of the faults on the output of an aircraft control system, namely y (t); further, a fault indication function can be set, and a first adaptive controller and a second adaptive controller are combined to be designed in a fusion mode to obtain a comprehensive controller so as to obtain a desired control signal of an aircraft control system, namely v (t).
According to the fault compensation and disturbance suppression method for the aircraft control system based on the self-adaptation provided by the embodiment of the invention, the interference model and the fault model are respectively established based on the interference item and the fault item, the dynamic model which influences the input of the aircraft control system is further established according to the interference model and the fault model, the uncertain influence of the interference and the fault on the input of the aircraft control system is expressed in a mathematical model mode, then the basic controller can be designed according to the dynamic model to suppress the influence of the interference on the output of the aircraft control system, meanwhile, the self-adaptation controllers can be respectively designed based on the fault model to compensate the influence of the fault on the output of the aircraft control system, and finally, the integrated controller can be designed based on the self-adaptation weighting fusion to compensate the influence of the fault and the disturbance on the performance of the aircraft control system, therefore, when the aircraft control system has gust disturbance and unknown, the method effectively inhibits gust disturbance and compensates faults, thereby ensuring that an aircraft control system has expected closed loop stability and output tracking performance, realizing stable and progressive tracking control of targets, improving fault-tolerant control effect and ensuring flight safety of the aircraft.
The effectiveness of the adaptive aircraft control system fault compensation and disturbance suppression method according to the embodiment of the present invention will be further described with specific simulation conditions given below, where the simulation conditions include:
aircraft parameters:
Figure BDA0002534041320000151
ρ＝0.7377kg/m3，Iy＝31027kg·m2，Cx1＝0.39，Cx2＝2.9099，Cx3＝-0.0758，Cx4＝0.0961，Cz1＝-7.0186，Cz2＝4.1109，Cz3＝-0.3112，Cz4＝-0.2340，Cz5＝-0.1023，Cm1＝-0.8789，Cm2＝-3.8520，Cm3＝-0.0108，Cm4＝-1.8987，Cm5＝-0.6266。
interference parameters:
d(t)＝[0.3cos(0.1t)+0.1,0.15sin(0.2t)+0.3cos(0.1t),0.03sin(0.1t)+0.1]TN·m/(kg·m2)。
in the simulation verification, the following fault conditions are considered, namely the fault modes meeting the set requirements of fault compensation are as follows: diag {1,0,0,0}, diag {0,1,0,0}, diag {0,0,1,0}, diag {0,0,0,1}, diag {0,0,0,0 }.
(i) When t is less than 150s, actuator u1No fault condition, ui(t)＝vi(t)，i＝1,2,3,4,t＜150s；
(ii) When t is more than or equal to 150s and less than or equal to 300s, the actuator u1The occurrence of a stuck-at fault,
u1(t)＝0.04rad,ui(t)＝vi(t),i＝2,3,4，for150s≤t≤300s；
(iii) when t is more than or equal to 300s and less than or equal to 400s, the actuator u1Return to normal, ui(t)＝vi(t)，i＝1,2,3,4；
(iv) When t is more than or equal to 400s, the actuator u4Occurrence of a stuck-at fault u4(t)＝300N,ui(t)＝vi(t),i＝1,2,3。
Simulation parameters: integrated controller
Figure BDA0002534041320000164
Has a parameter state of α11＝α21＝α31＝α32Other design parameters are 0.75: x is the number of0＝[0.008,0.72,0.008,0.008]T。
In addition, the basis functions, initial interference parameters and adaptive gains in the interference model are:
Figure BDA0002534041320000161
Figure BDA0002534041320000162
d1＝d2＝d3＝10I4。
and, the basis functions, initial interference parameters and adaptive gains in the fault model are:
Figure BDA0002534041320000163
based on the simulation experiment, a comparison relationship between reference signals corresponding to actual outputs of the aircraft control system shown in fig. 3, a tracking error of the aircraft control system shown in fig. 4, and a control input signal of the control system acted by the aircraft actuator shown in fig. 5 can be obtained.
Furthermore, it can be seen from fig. 3 and 4 that the method for compensating and suppressing the fault of the aircraft control system based on the adaptive control method of the embodiment of the invention can always realize the stable and asymptotically output and tracked control target of the closed-loop system no matter in the actual operation process, or under the condition that the fault occurrence time, the fault value and the fault mode are unknown, and it can be seen from fig. 5 that, in the time period of t ∈ [0,150s), when the actuator of the aircraft control system has external interference, the transient response will appear in the process of the control system outputting and tracking the given command, and the transient response will gradually decrease along with the time change, thereby verifying that the method for compensating and suppressing the fault of the aircraft control system based on the adaptive control method of the embodiment of the invention has robustness, and when t is 150s, respectively, the actuator u is implemented1Actuator u with fault sum t 400s4In case of a fault, the simulation result verifies the effectiveness of the fault compensation and disturbance suppression method of the aircraft control system based on self-adaptation in the embodiment of the invention.
In the present invention, unless otherwise expressly specified or limited, the term "coupled" is to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral part; can be mechanically or electrically connected; either directly or indirectly through intervening media, either internally or in any other relationship. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In the description herein, references to the description of the term "one embodiment," "some embodiments," "an example," "a specific example," or "some examples," etc., mean that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the invention. In this specification, the schematic representations of the terms used above are not necessarily intended to refer to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples. Furthermore, various embodiments or examples and features of different embodiments or examples described in this specification can be combined and combined by one skilled in the art without contradiction.
Although embodiments of the present invention have been shown and described, it will be appreciated by those skilled in the art that changes, modifications, substitutions and alterations can be made in these embodiments without departing from the principles and spirit of the invention, the scope of which is defined in the appended claims and their equivalents.

Claims (9)

1. An adaptive-based aircraft control system fault compensation and disturbance suppression method is characterized by comprising the following steps:
acquiring interference items and fault items influencing an aircraft control system;
respectively establishing an interference model and a fault model according to the interference item and the fault item;
establishing a dynamic model influencing the input of the aircraft control system according to the interference model and the fault model;
designing a base controller from the dynamic model, wherein the base controller is configured to suppress an effect of interference on the aircraft control system output;
respectively designing adaptive controllers based on the failure modes, wherein the adaptive controllers are used for compensating the influence of the failure on the output of the aircraft control system;
and designing a comprehensive controller based on the self-adaptive weighted fusion, wherein the comprehensive controller is used for compensating faults and inhibiting the influence of interference on the performance of the aircraft control system.
2. The adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 1, further comprising verifying the integrated controller using a lyapunov function.
3. The adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 2, wherein the disturbance term is a gust disturbance and the fault term is an indeterminate actuator fault of the aircraft.
4. The adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 3, wherein the disturbance model is:
Figure FDA0002534041310000011
wherein the content of the first and second substances,
Figure FDA0002534041310000012
in order for the interference parameters to be unknown,
Figure FDA0002534041310000013
is a known basis function of the known functions of, among others,
Figure FDA0002534041310000014
and
Figure FDA0002534041310000015
respectively as follows:
Figure FDA0002534041310000016
Figure FDA0002534041310000021
wherein j is 1,2, 3.
5. The adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 4, wherein the fault model is:
Figure FDA0002534041310000022
where σ (t) is the corresponding actuator failure mode matrix, v (t) is the control input signal to be designed,
Figure FDA0002534041310000023
is a fault parameter vector, where σ (t), v (t), and
Figure FDA0002534041310000024
respectively as follows:
σ(t)＝diag{σ1(t),σ2(t),σ3(t),σ4(t)}，
v(t)＝[v1(t),v2(t),v3(t),v4(t)]T，
Figure FDA0002534041310000025
6. the adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 5, wherein the dynamic model is:
Figure FDA0002534041310000026
where V is the aircraft speed, α is the angle of attack, θ is the pitch angle,qfor pitch angular rate, m for mass, IyIs moment of inertia, M is pitching moment, d1，d2And d3Is a turbulent perturbation signal.
7. The adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 6, wherein designing a base controller from the dynamic model comprises:
estimating interference parameters according to the interference model;
obtaining an interference design self-adaptive linear control law according to the estimation result;
setting an interference parameter self-adaptive law;
setting a Lyapunov function;
and verifying the parameter self-adaption law according to the Lyapunov function and the interference parameter self-adaption law.
8. The adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 7, wherein the fault modes include an actuator no fault mode and an actuator fault mode, wherein a first adaptive controller is designed based on the actuator no fault mode and a second adaptive controller is designed based on the actuator fault mode.
9. The adaptive-based aircraft control system fault compensation and disturbance rejection method of claim 8, wherein designing the integrated controller based on adaptive weighted fusion comprises:
setting a fault indication function;
and obtaining an integrated controller according to the fault indication function, the first self-adaptive controller and the second self-adaptive controller.
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