CN115964795A - Deformation control method for morphing aircraft based on disturbance observer - Google Patents

Deformation control method for morphing aircraft based on disturbance observer Download PDF

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CN115964795A
CN115964795A CN202210732930.0A CN202210732930A CN115964795A CN 115964795 A CN115964795 A CN 115964795A CN 202210732930 A CN202210732930 A CN 202210732930A CN 115964795 A CN115964795 A CN 115964795A
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aircraft
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沈庆成
吴楠
裔扬
朱政宇
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Jiangsu Qingya Electronic Technology Co ltd
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Abstract

The invention discloses a deformation control method of a variant aircraft based on an interference observer, which comprises the steps of establishing a longitudinal model of a conventional aircraft, introducing variable parameters related to wing camber into pneumatic parameters in the model, fitting the pneumatic parameters of the aircraft and the wing camber relationship, then carrying out small-disturbance linearization treatment on the pneumatic parameters, establishing a variable parameter model of the aircraft by utilizing the fitted functional relationship between the longitudinal pneumatic parameters of the aircraft and the wing camber, namely establishing a specific state space model, constructing the interference observer to estimate interference dynamics according to the constructed state space model, and combining an interference estimation value and a PI type state feedback controller to effectively control a controlled model so as to improve stability. And (3) combining a Lyapunov stability method and a convex optimization method, calculating to obtain a controller gain and an observer gain, and further ensuring that the controlled variant aircraft system obtains good dynamic performance.

Description

Deformation control method for morphing aircraft based on disturbance observer
Technical Field
The invention relates to the technical field of anti-interference control of aircrafts, in particular to a deformation control method of a variant aircraft based on an interference observer.
Background
In the traditional modeling method of the variant aircraft, multi-body modeling, analytical mechanics modeling and vector mechanics modeling are most common, and due to the influence of design factors of the variant aircraft, the variant aircraft has a complex deformation structure and an actuating mechanical structure, so that the modeling using the universal method is complex, and when the aerodynamic shape of the aircraft changes, the aerodynamic parameters of the aircraft can change along with the deformation structure, so that the aerodynamic parameters are not fixed and the accuracy is low.
At present, almost all systems have external interference, such as a motion control system, a robot control system, a complex chemical process, a terminal sliding mode system, a flight control system and the like. The aircraft has poor stability and poor control effect in the deformation process under the condition of input interference in the existing reaction system, so that the problem of anti-interference control attracts wide attention of academic and engineering circles.
Therefore, a deformation control method of a variant aircraft based on an interference observer is needed to be designed, in order to reduce the complexity of researching the problem of a system model, a theoretical model of the variant aircraft is constructed aiming at the longitudinal characteristic and the control problem of the variant aircraft with variable wing profile camber, a longitudinal model of a conventional aircraft is established, variable parameters related to the wing profile camber are introduced into pneumatic parameters in the model, namely, the pneumatic parameters of the aircraft are fitted with the relation of the wing profile camber, then small-disturbance linearization processing is carried out on the pneumatic parameters, a variable parameter model of the aircraft is established by utilizing the functional relation of the fitted longitudinal pneumatic parameters of the aircraft and the wing profile camber, namely, a specific state space model is established, and then a class of algorithm is designed according to a disturbance observer theory, so that the problem of system control with external disturbance is solved. A switching signal is constructed based on an average residence time method, and meanwhile, a PI type anti-interference composite controller is designed by combining state feedback information and interference estimation information. And then, the stability of the designed closed-loop system is proved based on a Lyapunov stability analysis method.
Disclosure of Invention
The invention aims to provide a deformation control method of a variant aircraft based on a disturbance observer, which is used for solving the problems that the conventional aircraft modeling method is complex and has low accuracy and the stability of the aircraft is poor in the deformation process under the condition that the existing reaction system has input disturbance.
In order to achieve the purpose, the specific technical scheme of the method for controlling the deformation of the morphing aircraft based on the disturbance observer is as follows:
a deformation control method of a variant aircraft based on a disturbance observer comprises the following steps:
step (1): based on the conventional dynamics modeling of a general aircraft, obtaining a dynamics equation and a kinematics equation of the general aircraft; obtaining model equations of the aircraft in different coordinate shafting according to the aircraft kinetic equation and the kinematics equation; decoupling a motion equation of the aircraft model to obtain a longitudinal motion equation irrelevant to the transverse state quantity;
step (2): based on a longitudinal motion equation and a wing switching principle, fitting to obtain aerodynamic parameters of wings with a camber relation, then substituting into a longitudinal model of the morphing aircraft and balancing to obtain balance points changing along with the camber of the wings, and combining the balance points to construct a longitudinal small disturbance linearization equation and a parameter-containing model of the morphing aircraft, namely establishing a specific state space model;
and (3): respectively designing a PI type controller and an interference observer according to the established state space model and in combination with the existence of interference, so as to realize estimation of unknown interference and effective control of output;
and (4): and combining a Lyapunov stability analysis method to obtain corresponding controller gain and observer gain, and bringing the controller gain and observer gain into a state space model to be applied to complete the anti-interference control of the deformation of the aircraft.
Compared with the prior art, the invention has the beneficial effects that:
(1) Compared with the conventional aircraft modeling method, the variable parameter modeling method is simpler and higher in accuracy, and provides another idea for the control research of the variable aircraft.
(2) The invention carries out control analysis aiming at the camber of the variable wing of the aircraft, and designs reasonable average residence time to ensure that the control problem of the deformation process has better effect.
(3) The interference observer method of the invention can estimate and compensate the interference existence condition, effectively inhibit the interference and improve the stability of the variant process.
Drawings
FIG. 1 is a flow chart of the present invention.
Detailed Description
A method for controlling deformation of a morphing aircraft based on a disturbance observer, as shown in fig. 1, comprises the following steps:
the method comprises the following steps that (1) a dynamic equation and a kinematic equation of the general aircraft are obtained based on the conventional dynamic modeling of the general aircraft; obtaining model equations of the aircraft in different coordinate shafting according to the aircraft kinetic equation and the kinematics equation; decoupling the motion equation of the aircraft model to obtain a longitudinal motion equation irrelevant to the transverse state quantity, which specifically comprises the following steps: the kinetic and kinematic equations of the aircraft body coordinate system are given below, according to newton's law of motion, as follows:
force equation set
Figure BDA0003714539260000031
System of equations of motion
Figure BDA0003714539260000041
System of moment equations
Figure BDA0003714539260000042
Navigation equations set
Figure BDA0003714539260000043
According to the aircraft dynamic equation and the kinematics equation, model equations of different coordinate shafting of the aircraft can be obtained:
(1) Body coordinate system model equation
Assuming an engine thrust offset angle of alpha T ＝β T =0, i.e. thrust on the x-axis of the body coordinate system, T = T x . The body coordinate system may be converted to the air flow coordinate system as follows:
Figure BDA0003714539260000044
wherein the conversion matrix
Figure BDA0003714539260000045
Is defined as follows:
Figure BDA0003714539260000046
substituting the forces in the above equation can yield:
Figure BDA0003714539260000051
unfolding the above formula yields:
Figure BDA0003714539260000052
substituting the above equation into the force equation set equation to obtain a model equation of a body coordinate system:
Figure BDA0003714539260000053
where α represents the angle of attack of the aircraft and β represents the angle of sideslip.
(2) Air flow coordinate system model equation
Assuming that the total external force of the aircraft is F W Including engine thrust T, total aerodynamic R ∑ And the component of gravity G on each axis of the airflow coordinate system, namely:
Figure BDA0003714539260000054
and substituting the aircraft force equation set formula to obtain an airflow coordinate system model equation:
Figure BDA0003714539260000055
in this context, we assume that the aircraft is in an ideal flight environment, i.e. ideal flight conditions without horizontal and side-slip (phi = beta ≡ 0, p =r ≡ 0,
Figure BDA0003714539260000056
and->
Figure BDA0003714539260000057
) According to the model equation of each coordinate system obtained in the previous step, the motion equation of the aircraft model is decoupled, and a longitudinal motion equation irrelevant to the transverse and lateral state quantities can be obtained:
Figure BDA0003714539260000061
in the formula, V is the speed of the aircraft, alpha represents the attack angle of the aircraft, m is the mass of the aircraft, and g is the gravity acceleration; theta and q are a pitch angle and a pitch angle speed respectively; I.C. A y Is the moment of inertia about the shaft; lift L, resistance D, pitching moment M, thrust T do respectively:
Figure BDA0003714539260000062
Figure BDA0003714539260000063
Figure BDA0003714539260000064
Figure BDA0003714539260000065
wherein, C L Is a coefficient of lift, C D Is a coefficient of resistance, C M Is a moment coefficient, T δt Is the thrust coefficient of the engine, delta t Is the aircraft throttle opening.
And (2) fitting to obtain aerodynamic parameters of the wing with a camber relation based on a longitudinal motion equation and a wing switching principle, then substituting the aerodynamic parameters into a longitudinal model of the morphing aircraft and balancing to obtain balance points changing along with the camber of the wing, and constructing a longitudinal small disturbance linearization equation and a parameter-containing model of the morphing aircraft by combining the balance points, namely establishing a specific state space model, wherein the specific process is as follows:
firstly, giving aerodynamic parameters of wings:
(1) Aircraft lift parameter
Aircraft fly over the sky by relying on lift provided by movement in the atmosphere. The lift of the aircraft is mainly derived from wing lift L ω Lift L of the fuselage b And horizontal tail lift force L t Namely:
L＝L ω +L b +L t
the lift force of each part is substituted to obtain:
Figure BDA0003714539260000071
the lift force mainly comes from wings, the full-motion horizontal tail control surface aircraft model is used as the model used in the chapter, the aerodynamic force generated by the control surface is extremely small compared with the lift force of the aircraft body, therefore, the influence of the lift force on the overall moment of the aircraft can be ignored, and then a lift force formula is optimized to be simplified into the sum of the lift force of the wings and the lift force of the full-motion horizontal tail control surface.
Based on the above analysis, the aircraft lift coefficient may be modified to:
Figure BDA0003714539260000072
wherein, a zero attack angle lift coefficient C of the whole aircraft is defined L0 Directly taking zero attack angle lift coefficient C after fitting of deformed wing l0 Aerodynamic derivative of the total lift C Lα Taking the aerodynamic derivative C of the lift force of the deformed wing lα Then, there are:
C L0 ＝0.1044f+0.02036
C Lα ＝5.73/rad
(2) Aircraft resistance parameter
The drag factors that make up an aircraft are complex, typically consisting of zero lift drag D L0 And lift-induced resistance D t Constituting aircraft drag. The zero lift resistance comprises friction resistance, pressure difference resistance and zero lift wave resistance; the lift-induced drag comprises induced drag and lift-induced wave drag, and therefore, the overall drag coefficient of the aircraft is as follows:
Figure BDA0003714539260000073
wherein a zero-lift drag coefficient is defined
Figure BDA0003714539260000074
Defining a rising resistance factor->
Figure BDA0003714539260000075
When the aircraft enters a subsonic flight state, the induced resistance is a main constituent factor of the lift-induced resistance:
Figure BDA0003714539260000076
wherein ε is the down wash angle of the aircraft. When the aircraft is in a supersonic flight state, the lift-induced wave drag is a main component factor of the lift-induced drag:
Figure BDA0003714539260000077
according to the relationship between the aircraft resistance and the attack angle, the resistance coefficient can be written as:
C D ＝C D0 +C Dα α
the resistance coefficient is related to the lift coefficient through the composition factors of the aircraft resistance, the resistance is mainly derived from the action of airflow on the lift component and the aircraft body, the aircraft body resistance and the attack angle can be approximately regarded as a linear relation, and the C is used Db K is expressed as k α, k is a suitably constant value, and the aircraft drag can be approximated as the sum of the fuselage and wing drag. Aircraft integral zero angle of attack drag coefficient C D0 And aerodynamic derivative of bulk drag C Dα Comprises the following steps:
C D0 ＝0.0008774f+0.007328
C Dα ＝(-0.8634f+3.2506)α+k
(3) Aircraft longitudinal pitching moment parameter
The moment can affect the flight performance and attitude of the aircraft, and is also an important parameter in the aircraft, and generally, the moment is mainly generated by the lift force and the aerodynamic force of a control surface of the aircraft, and can be described as follows:
M a ＝C M QS ω c A ＝(C m +C mb +C mt )QS ω c A
according to the formula, the overall moment of the aircraft is composed of moments generated by wings, a fuselage and a horizontal tail.
The aircraft used herein is a full-motion horizontal tail, ignoring the lift of the fuselage, so the total static pitching moment coefficient can be simplified as:
Figure BDA0003714539260000081
wherein the content of the first and second substances,
Figure BDA0003714539260000082
and &>
Figure BDA0003714539260000083
Respectively, defined as the relative position of the aerodynamic focus of the aircraft and its center of gravity on the average geometric chord. Subsequently, the moment coefficients are simplified: the integral attack angle moment coefficient C of the aircraft M0 Zero angle of attack moment coefficient C defined as a deformed wing m0 Aerodynamic derivative of the overall moment C Mα Defined as the moment starting derivative C of the deformed wing mα Then, there are:
C M0 ＝0.0166f+0.0126
C Mα ＝-0.1962/rad
in conclusion, the aerodynamic parameters of the wing with the camber relation are obtained by fitting after being substituted into the functional relation between the aerodynamic parameters of the morphing aircraft and the camber of the wing, and the specific expression is as follows:
Figure BDA0003714539260000091
C D ＝C D0 +C Dα α＝0.0008774f+(3.2506-0.8634f)α 2 +0.007328
Figure BDA0003714539260000092
wherein, C L Is the coefficient of lift, C D Is a coefficient of resistance, C M Is the moment coefficient, δ e Is the elevator declination.
And after the pneumatic parameters of the aircraft are obtained, substituting the function relation into a system model of the variant aircraft to further research the control problem.
The second is the trim of the aircraft, i.e. the determination of its balance point. The basic principle is as follows: the earth is taken as a coordinate system, so that the resultant force of the aircraft on the two axes of X and Z in the coordinate system is 0, at this time, the force is in a balanced state, and the longitudinal moment is balanced, namely, the trim is generally performed for the longitudinal direction. Therefore, the method is used for carrying out balancing on the variant aircrafts under different wing section bending degrees f states to obtain an attack angle alpha and an elevator deflection angle delta under a balanced state e And thrust T of the engine
From the above principles, a simple equation for the trim is obtained:
Figure BDA0003714539260000093
the method is substituted into pneumatic parameters under various camber conditions to obtain balance points under different camber, and is combined with a physical parameter model of a variant aircraft to be substituted into a correlation formula of lift L, resistance D, moment M and thrust T of the aircraft to obtain the balance points changing along with the camber.
Further, a functional relationship of the fitted balance point to the change in camber can be found:
α trim ＝-1.242f+5.62
Figure BDA0003714539260000094
Figure BDA0003714539260000095
wherein alpha is trim Is an incidence angle balance point of the aircraft,
Figure BDA0003714539260000101
for the point of equilibrium of the deflection angle of the elevator, is>
Figure BDA0003714539260000102
Is the balance point of the opening degree of the accelerator of the aircraft.
After a longitudinal model equation of the aircraft is obtained, the system model equation is subjected to small-disturbance linearization at a balance point, and a state equation of the aircraft can be obtained:
Figure BDA0003714539260000103
in the above formula, the first and second carbon atoms are,
Figure BDA0003714539260000104
state variable X = [. DELTA.V.DELTA.alpha.DELTA.theta.DELTA.q] T Input variable (i.e., control variable) U = [. DELTA.. Delta. ] e △δ T ] T 。
Carrying out linearization processing on the balance point to obtain a matrix E, a state matrix A and a control matrix B:
Figure BDA0003714539260000105
Figure BDA0003714539260000106
Figure BDA0003714539260000107
then, the model of the aircraft is simplified, considering that the flight condition of the aircraft is a sideslip-free horizontal flight, and therefore the aerodynamic dynamic derivative related to the incidence angle change rate in the matrix E is set
Figure BDA0003714539260000109
And if the value is 0, the E matrix just becomes an identity matrix, and further, a linearized form of the state equation of the aircraft is obtained:
Figure BDA0003714539260000108
as in the above, the above-mentioned,pitch rate dependent aerodynamic derivatives in the state matrix A
Figure BDA0003714539260000111
Aerodynamic derivative dependent on velocity V
Figure BDA0003714539260000112
And the thrust derivative pick-up>
Figure BDA0003714539260000113
All are approximately 0, the a matrix can be simplified to:
Figure BDA0003714539260000114
in addition, the angle of attack α due to the balance point trim Small and the product of the mass of the aircraft and the initial speed is very large, so in control matrix B
Figure BDA0003714539260000115
This term can be approximated as 0, and the matrix B equation can be simplified as: />
Figure BDA0003714539260000116
Available to this end
Figure BDA0003714539260000117
V in equation, state matrix A and control matrix B expressed as simplified model after linearization of morphing aircraft 0 ，C L0 ，Q 0 ，C M0 ，C D0 The equal components respectively represent the flight speed, lift coefficient, dynamic pressure, pitching moment coefficient and drag coefficient when the aircraft is at the balance point. Substituting the state matrix A and the control matrix B into a function relation of the fitted pneumatic parameters and the camber to obtain a variable parameter matrix model containing a variable parameter (camber f):
Figure BDA0003714539260000118
Figure BDA0003714539260000119
a (f) is a state matrix of a reference model of a variant aircraft with an airfoil camber change parameter f, the aircraft increases the span camber f by increasing the thickness of the upper surface of an airfoil, and the flight state is changed by the change of state quantity caused by the span camber f, wherein the relative camber increasing amplitude of the variant aircraft is 0.25%, the relative camber change range of the airfoil is 1-2.5%, namely, the value of f is between [1,2.5 ]. The numerical simulation is respectively carried out on the changed wing profiles. The influence of the aerodynamic characteristics of the morphing wing with changes in camber can be obtained.
According to the system matrix obtained by calculation, the following state space model is established:
Figure BDA0003714539260000121
Figure BDA0003714539260000122
Figure BDA0003714539260000123
wherein d (t) represents unknown interference and satisfies:
Figure BDA0003714539260000124
where w (t) is the interference state and H, Y represent parameters related to the interference type.
Respectively designing a PI type controller and a disturbance observer according to the established state space model and in combination with the existence of disturbance, so as to realize estimation of unknown disturbance and effective control of output; utensil for cleaning buttockThe body is as follows: to obtain good tracking performance, new augmentation variables are defined
Figure BDA0003714539260000125
Wherein z (t) is the aircraft state, e y The difference between the system output and the desired output;
the augmentation system may be described as:
Figure BDA0003714539260000131
wherein
Figure BDA0003714539260000132
A disturbance observer is then constructed to estimate the disturbance, in the following form:
Figure BDA0003714539260000133
in the formula
Figure BDA0003714539260000134
Is an estimate of the interference; the interference estimation error is defined as: />
Figure BDA0003714539260000135
Further, designing a PI state feedback controller by combining interference estimation comprises the following steps:
Figure BDA0003714539260000136
and (4) solving by combining a Lyapunov stability analysis method to obtain corresponding controller gain and observer gain, and bringing the controller gain and observer gain into a state space model to be applied to complete the anti-interference control of the aircraft deformation.
The lyapunov analysis method used to solve for the controller and observer gains is expressed as:
theorem 1: for a given constant T f >0，α>0, mu is more than or equal to 1. Here, if there is a positive definite symmetric matrix P i ∈R m×m ，Q i ∈R k×k Satisfies the following conditions:
Figure BDA0003714539260000137
P i ≤μP j ，Q i ≤μQ j ，
Figure BDA0003714539260000138
j∈Z，i≠j/>
Figure BDA0003714539260000139
it can be deduced that the variant aircraft system is stable and the disturbance estimation error system is convergent and good tracking performance can be obtained, wherein the controller and disturbance observer gains can be gained by
Figure BDA00037145392600001310
V i ＝K i ρ i And are each selected from
Figure BDA00037145392600001311
And (6) calculating to obtain.
And (3) proving that: the following Lyapunov function was chosen: phi is a i ＝φ 1,i +φ 2,i ，φ 2,i ＝e w T (t)Q i e w (t)，
Figure BDA00037145392600001312
/>
The method is easy to obtain according to the step (3):
Figure BDA0003714539260000141
based on the Schur complement theory, the diag { rho is multiplied on two sides of the first matrix inequality simultaneously -1 i I I, one can obtain:
Figure BDA0003714539260000142
when phi is i = h and->
Figure BDA0003714539260000143
At time, there is->
Figure BDA0003714539260000144
From this, it can be seen that the initial condition is phi σ(0) When = h, has phi σ(t) ≤h,/>
Figure BDA0003714539260000145
Is paired and/or matched>
Figure BDA0003714539260000146
In [0,t]By integrating above, we can get:
Figure BDA0003714539260000147
at this time, if there is phi σ(t) (ζ (t)) > α/h, there are
Figure BDA0003714539260000148
Next, we aim at phi σ(t) Two cases of (ζ (t)) are discussed:
(1) When the system is running, the system does not satisfy the condition
Figure BDA0003714539260000149
I.e. each subsystem satisfies the condition
Figure BDA00037145392600001410
At this point, the switching system is bounded.
(2) When the system is running, the system can satisfy the condition
Figure BDA00037145392600001411
When the system is running, we can get->
Figure BDA00037145392600001412
For the formula from t k Integration by t yields:
Figure BDA00037145392600001413
after simplifying the above inequalities, the following forms can be obtained:
Figure BDA00037145392600001414
suppose that at the switching time t k Existence of
Figure BDA00037145392600001415
When inequality in the theorem is satisfied
In time, there are:
Figure BDA00037145392600001416
in combination with the above formula, one obtains:
Figure BDA0003714539260000151
based on the above conditions and the iterative theorem, we can get:
Figure BDA0003714539260000152
according to the theory, the method can be known,
Figure BDA0003714539260000153
thus, the switching system is bounded.
From the above two conditions, phi σ(t) (ζ (t)) is ultimately achievable to be bounded. And can know phi σ(t) (ζ (t)) can eventually converge to a boundary π and reachTo a stable state in which
Figure BDA0003714539260000154
As phi σ(t) A component of (ζ (t)), based on the comparison result>
Figure BDA0003714539260000155
And eventually converge to within a boundary. It can be seen that when t → ∞ is present, the alkyl radical is present in the alkyl radical>
Figure BDA0003714539260000156
Exist and are bounded. Through the above proof, it can be proved that the designed error tracking system is effective.
The present invention is not limited to the above-mentioned embodiments, and based on the technical solutions disclosed in the present invention, those skilled in the art can make some substitutions and modifications to some technical features without creative efforts according to the disclosed technical contents, and these substitutions and modifications are all within the protection scope of the present invention.

Claims (8)

1. A deformation control method of a variant aircraft based on a disturbance observer is characterized by comprising the following steps:
step (1): based on the conventional dynamics modeling of a general aircraft, obtaining a dynamics equation and a kinematics equation of the general aircraft; obtaining model equations of the aircraft in different coordinate shafting according to the aircraft kinetic equation and the kinematics equation; decoupling the motion equation of the aircraft model to obtain a longitudinal motion equation irrelevant to the transverse state quantity;
step (2): based on a longitudinal motion equation and a wing switching principle, fitting to obtain aerodynamic parameters of wings with a camber relation, substituting the aerodynamic parameters into a longitudinal model of the morphing aircraft, balancing to obtain balance points changing along with the camber of the wings, and constructing a longitudinal small disturbance linearization equation and a parameter-containing model of the morphing aircraft by combining the balance points, namely establishing a specific state space model;
and (3): respectively designing a PI type controller and an interference observer according to the established state space model and in combination with the existence of interference, so as to realize estimation of unknown interference and effective control of output;
and (4): and combining a Lyapunov stability analysis method to obtain corresponding controller gain and observer gain, and bringing the controller gain and observer gain into a state space model to be applied to complete the anti-interference control of the deformation of the aircraft.
2. The disturbance observer-based morphing aircraft control method according to claim 1, wherein the model equations of the aircraft in different coordinate axis systems in the step (1) are a body coordinate system model equation and an air flow coordinate system model equation respectively.
3. The method for controlling deformation of a morphing aircraft based on a disturbance observer according to claim 1, wherein the equations of motion of the general aircraft model in step (1) are decoupled, and the longitudinal equations of motion independent of the lateral state quantity are obtained as follows:
Figure FDA0003714539250000021
in the formula, theta and q are a pitch angle and a pitch angle speed respectively; m is the aircraft mass, g is the gravitational acceleration; I.C. A y Is the moment of inertia about the shaft; lift D, pitching moment M, thrust T are respectively:
Figure FDA0003714539250000022
Figure FDA0003714539250000023
Figure FDA0003714539250000024
Figure FDA0003714539250000025
wherein, C L Is a coefficient of lift, C D Is a coefficient of resistance, C M In order to be the moment coefficient,
Figure FDA0003714539250000026
is the thrust coefficient of the engine, delta t Is the aircraft throttle opening.
4. The method for controlling deformation of a morphing aircraft based on a disturbance observer according to claim 1, wherein in the step (2), aerodynamic parameters of the wing with a camber relationship are obtained by fitting based on a longitudinal motion equation and a wing switching principle, and the specific expression is as follows:
Figure FDA0003714539250000027
C D ＝C D0 +C Dα α＝0.0008774f+(3.2506-0.8634f)α 2 +0.007328
Figure FDA0003714539250000028
wherein, C L Is a coefficient of lift, C D Is a coefficient of resistance, C M Is the moment coefficient, δ e Is the elevator declination.
5. The method for controlling deformation of a morphing aircraft based on a disturbance observer according to claim 4, wherein the specific expression of the balance point of the airfoil curvature change in the step (2) is as follows:
Figure FDA0003714539250000029
when the balance equation is satisfied, the balance points under different curvatures can be obtained, and further the functional relation between the balance points and the curvatures is obtained:
α trim ＝-1.242f+5.62
Figure FDA0003714539250000031
Figure FDA0003714539250000032
wherein alpha is trim Is an incidence angle balance point of the aircraft,
Figure FDA0003714539250000033
for the point of equilibrium of the deflection angle of the elevator, is>
Figure FDA0003714539250000034
Is the balance point of the opening degree of the accelerator of the aircraft.
6. The method for controlling deformation of a morphing aircraft based on a disturbance observer according to claim 5, wherein in the step (2), a longitudinal small disturbance linearization equation and a parameter-containing model of the morphing aircraft are constructed by combining balance points according to the functional relation of the wing profiles corresponding to the aerodynamic parameters of the aircraft, and are as follows:
Figure FDA0003714539250000035
wherein X represents the flight state of the aircraft, U = [. DELTA.. Delta.) e ,△δ t ] T Representing the control input, namely the deflection angle of an aircraft elevator, and A and B represent the parameter model matrix containing the bending function of the system.
According to the system matrix obtained by calculation, establishing a state space model:
Figure FDA0003714539250000036
Figure FDA0003714539250000037
Figure FDA0003714539250000038
a (f) is a state matrix containing a reference model of the variant aircraft with an airfoil camber change parameter f, and the value of f is between [1,2.5 ]. d (t) represents unknown interference and satisfies:
Figure FDA0003714539250000039
/>
where w (t) is the interference state and H, Y represent parameters related to the interference type.
7. The method for controlling deformation of a morphing aircraft based on a disturbance observer according to claim 1, wherein the step (3) is specifically:
defining new augmented variables
Figure FDA0003714539250000041
Wherein z (t) is the aircraft state, e y For the difference between the system output and the desired output, the augmentation system is described as:
Figure FDA0003714539250000042
wherein
Figure FDA0003714539250000043
e y ＝r(t)-r d ，r d Is the desired output. The following construction of a disturbance observer is proposed to estimate the disturbance, in the following form:
Figure FDA0003714539250000044
in the formula
Figure FDA0003714539250000045
Is an estimate of the interference, v (t) is a set auxiliary variable, L i Is the observer gain. The interference estimation error is defined as: />
Figure FDA0003714539250000046
Designing a PI state feedback controller by combining interference estimation as follows:
Figure FDA0003714539250000047
wherein K i Representing the controller gain that needs to be solved for.
8. The disturbance observer-based morphing aircraft control method according to claim 1, wherein the step (4) of lyapunov analysis for solving the controller and observer gains is expressed as:
for a given constant T f >0，α>0, mu is more than or equal to 1. Here, if there is a positive definite symmetric matrix P i ∈R m×m ，Q i ∈R k×k Satisfies the following conditions:
Figure FDA0003714539250000048
Figure FDA0003714539250000051
it can be deduced that the variant aircraft system is stable and the disturbance estimation error system is convergent and good tracking performance can be obtained, where the controller and disturbance observer gains can be given by P i -1 ＝ρ i ，V i ＝K i ρ i And are each selected from
Figure FDA0003714539250000052
And (6) calculating to obtain. />
CN202210732930.0A 2022-06-27 2022-06-27 Deformation control method for morphing aircraft based on disturbance observer Pending CN115964795A (en)

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN116909307A (en) * 2023-09-12 2023-10-20 中国人民解放军32806部队 High-maneuvering motion control method for aircraft
CN116909307B (en) * 2023-09-12 2023-12-19 中国人民解放军32806部队 High-maneuvering motion control method for aircraft

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