A method for controlling a rigid-flexible coupling rotating aircraft based on a sliding mode and an observer

Abstract

The application provides a rigid-elastic coupling rotating aircraft control method based on a sliding mode and an observer, which comprises the following steps: firstly, establishing a rigid-elastic coupling model, a rigid body dynamics model, an elastic deformation dynamics model of the rigid-elastic coupling rotating aircraft, and a conversion model between a measured angular velocity and a rigid body angular velocity; secondly, establishing a state equation of a control object of the rigid-elastic coupling rotating aircraft; and finally, establishing a control law of the rigid-elastic coupling rotating aircraft based on the sliding mode and the observer, so as to realize the stable control of the rigid body motion and the vibration suppression of the elastic deformation of the rotating aircraft through an active control technology. The application solves the problem that the elastic deformation of the rigid-elastic coupling rotating aircraft seriously affects the stability of the rigid body motion, the conventional wave trap or the adaptive wave trap is difficult to adapt to the case that the elastic vibration frequency is close to the rigid body motion frequency, and the attitude system of the rotating aircraft is unstable and the structure is damaged.

Description

TECHNICAL FIELD

[0001] The application relates to a rigid-elastic coupling rotating aircraft control method based on a sliding mode and an observer, which is used for realizing stable control of rigid body motion and vibration suppression of elastic deformation of a rigid-elastic coupling rotating aircraft and belongs to the technical field of overall design and control of aircrafts. BACKGROUND

[0002] The rotating aircraft refers to a kind of aircraft rotating around its longitudinal axis at a certain rotating speed during flight. The rotating aircraft has many advantages, such as reducing the influence of deviation in the machining process on control accuracy, simplifying the control system, reducing cost, reducing the influence of asymmetric factors on control accuracy during flight, etc. However, the influence of elastic deformation motion of the rotating aircraft on system stability cannot be ignored. Since the measurement unit is directly installed on the aircraft, the elastic vibration signal of the aircraft will enter the control loop through the measurement unit, so that the controller generates additional command output. The additional command output of the servo mechanism in response to the elastic vibration signal will further intensify the elastic vibration, causing instability of the attitude system and damage to the structure of the aircraft.

[0003] The common method for the influence of elastic deformation of the rotating aircraft on motion stability is a passive control method, such as optimal design of measurement unit layout, wave trap, adaptive wave trap, etc. The influence of the elastic vibration signal on the stability of the control system is reduced through self-compensation of the controller. However, the optimal design method of the measurement unit layout is difficult to adapt to the time-varying vibration mode in the flight process. Both the conventional wave trap and the adaptive wave trap work when the frequency of the elastic vibration is higher than that of the rigid body motion, that is, the elastic vibration is filtered as high-frequency noise. However, when the frequency of the elastic vibration is close to that of the rigid body motion, the conventional wave trap or the adaptive wave trap is no longer applicable.

[0004] Therefore, in view of the fact that the elastic deformation of the rigid-elastic coupling rotating aircraft seriously affects the stability of the rigid body motion, and the conventional wave trap or the adaptive wave trap is difficult to adapt to the case where the frequency of the elastic vibration is close to that of the rigid body motion, it is necessary to propose a rigid-elastic coupling rotating aircraft control method based on a sliding mode and an observer, so as to realize stable control of the rigid body motion and vibration suppression of the elastic deformation through active control technology. SUMMARY

[0005] The technical problem to be solved by the application is to overcome the shortcomings of the prior art, solve the problem that the elastic deformation of the rigid-elastic coupling rotating aircraft seriously affects the stability of the rigid body motion, and the conventional wave trap or the adaptive wave trap is difficult to adapt to the case where the frequency of the elastic vibration is close to that of the rigid body motion, thereby causing instability of the attitude system of the rotating aircraft and structural damage.

[0006] The object of the application is achieved by the following technical scheme.

[0007] A rigid-flexible coupling rotary aircraft control method based on a sliding mode and an observer, comprising:

[0008] Given a rigid-flexible coupling model of a rigid-flexible coupling rotary aircraft;

[0009] A rigid body dynamics model of the rigid-flexible coupling rotary aircraft in a quasi-aircraft body coordinate system is established;

[0010] An elastic deformation dynamics model of the rigid-flexible coupling rotary aircraft in an aircraft body coordinate system is established;

[0011] A conversion model between the measured angular velocity and the rigid body angular velocity of the rigid-flexible coupling rotary aircraft is established;

[0012] A state equation of the control object of the rigid-flexible coupling rotary aircraft is established;

[0013] An observer of the rigid-flexible coupling rotary aircraft is established;

[0014] A sliding mode control law of the rigid-flexible coupling rotary aircraft is established;

[0015] A rigid-flexible coupling rotary aircraft control law based on a sliding mode and an observer is established, and stable control of the rotary aircraft is realized.

[0016] Compared with the prior art, the present application has the following beneficial effects:

[0017] (1) Compared with the conventional notch filter or adaptive notch filter, the rigid-flexible coupling rotary aircraft control method based on a sliding mode and an observer can realize stable control of rigid body motion and vibration suppression of elastic deformation in the case where the elastic vibration and the rigid body motion frequency are close.

[0018] (2) Compared with the conventional notch filter or adaptive notch filter, the rigid-flexible coupling rotary aircraft control method based on a sliding mode and an observer can realize stable control of rigid body motion and vibration suppression of elastic deformation by actively estimating the vibration information.

[0019] (3) For the elastic deformation rotary aircraft with strong uncertainty characteristics, the observer can estimate the system state variables and internal and external disturbances, better handle the strong uncertainty caused by elastic deformation, effectively eliminate the interference caused by elastic deformation on the system, and improve the robustness of the controller of the rotary aircraft.

[0020] (4) For the elastic deformation of the rotating aircraft with strong nonlinear characteristics, the sliding mode control is a robust nonlinear control method, and the sliding mode control law based on the non-singular terminal sliding mode surface and the double power approaching law has the finite time convergence characteristic, which ensures that the attitude angle error of the rotating aircraft can converge to zero in a limited time, and is beneficial to improve the performance of the controller of the rotating aircraft. BRIEF DESCRIPTION OF DRAWINGS

[0021] Figure 1 The method flowchart of the application.

[0022] Figure 2 The control block diagram of the rigid-flexible coupling rotating aircraft. DETAILED DESCRIPTION

[0023] In order to make the purpose, technical scheme and advantages of the application clearer, the embodiments of the application will be further described in detail below with reference to the drawings.

[0024] The elastic deformation of the rigid-flexible coupling rotating aircraft seriously affects the stability of the rigid body motion, and the conventional notch filter or adaptive notch filter is difficult to solve the case that the elastic vibration frequency is close to the rigid body motion frequency, thereby causing the instability of the attitude system of the rotating aircraft and the structural damage, in order to solve this problem, the application provides a rigid-flexible coupling rotating aircraft control method based on sliding mode and observer, first, the rigid-flexible coupling model, the rigid body dynamics model and the elastic deformation dynamics model of the rigid-flexible coupling rotating aircraft are established, and the conversion model between the measured angular velocity and the rigid body angular velocity is established, then the state equation of the control object of the rigid-flexible coupling rotating aircraft is established, and finally the control law of the rigid-flexible coupling rotating aircraft based on sliding mode and observer is established, and the stable control of the rigid body motion of the rotating aircraft and the vibration suppression of the elastic deformation are realized through active control technology.

[0025] Specifically includes:

[0026] (1) The rigid-flexible coupling model of the rigid-flexible coupling rotating aircraft is given.

[0027] (2) The rigid body dynamics model of the rigid-flexible coupling rotating aircraft in the quasi-aircraft body coordinate system is established.

[0028] (3) The elastic deformation dynamics model of the rigid-flexible coupling rotating aircraft in the aircraft body coordinate system is established.

[0029] (4) The conversion model between the measured angular velocity and the rigid body angular velocity of the rigid-flexible coupling rotating aircraft is established.

[0030] (5) The state equation of the control object of the rigid-flexible coupling rotating aircraft is established.

[0031] (6) The observer of the rigid-flexible coupling rotating aircraft is established.

[0032] (7) Establish a sliding mode control law for rigid-flexible coupling rotary aircraft.

[0033] (8) Establish a sliding mode and observer-based control law for rigid-flexible coupling rotary aircraft to achieve stable control of rotary aircraft.

[0034] More specifically, the present application proposes a sliding mode and observer-based control method for rigid-flexible coupling rotary aircraft, as shown in the accompanying drawings, comprising the following steps: Figure 1

[0035] (1) Given the rigid-flexible coupling model of the rigid-flexible coupling rotary aircraft.

[0036] The rigid-flexible coupling model of the rigid-flexible coupling rotary aircraft in the present application refers to that the rigid motion and elastic deformation are decoupled in the motion degrees of freedom, the coupling of rigidity and elasticity lies in the calculation of force and torque, and the elastic vibration signal of the aircraft enters the control loop through the measurement unit.

[0037] (2) Establish the rigid body dynamics model of the rigid-flexible coupling rotary aircraft in the quasi-aircraft body coordinate system.

[0038] The rigid body dynamics model of the rigid-flexible coupling rotary aircraft in the quasi-aircraft body coordinate system is established

[0039]

[0040] In the formula, α * is the quasi-attack angle, β * is the quasi-sideslip angle, m is the mass of the rotary aircraft, J x ,J y ,J z are the moments of inertia of the rotary aircraft in the OX4 axis, the OY4 axis and the OZ4 axis of the quasi-aircraft body coordinate system, V is the speed of the rotary aircraft, ω x4 ,ω y4 ,ω z4 are the angular velocities of the rotary aircraft in the OX4 axis, the OY4 axis and the OZ4 axis of the quasi-aircraft body coordinate system, θ is the pitch angle of the mass of the rotary aircraft, γ is the roll angle of the rotary aircraft, ψ is the yaw angle of the rotary aircraft, M yc ,M zc are the control torques of the rotary aircraft in the OY4 axis and the OZ4 axis of the quasi-aircraft body coordinate system, M x4 ,M y4 ,M z4 are the non-control torques of the rotary aircraft in the OX4 axis, the OY4 axis and the OZ4 axis of the quasi-aircraft body coordinate system, F y4 ,F z4 are the external forces of the rotary aircraft in the OY4 axis and the OZ4 axis of the quasi-aircraft body coordinate system. ​

[0041] (3) The elastic deformation dynamics model of the rigid-elastic coupling rotating aircraft in the body coordinate system of the aircraft is established.

[0042]

[0043] In the formula, η i ,ε i are the generalized coordinates describing the deformation of the rotating aircraft in the y direction and the z direction, ω i is the natural frequency of the i-th mode of the aircraft, φ i (x) is the mode function, M i is the generalized mass of the i-th mode of the aircraft, f y (x), f z (x) are the distributed forces acting on the microelement of the shaft section in the y direction and the z direction, respectively, x is the axial coordinate of the microelement of the shaft section, μ i is the damping ratio of the i-th mode of the aircraft, ω x1 , ω y1 , ω z1 are the rotational angular velocities of the rotating aircraft in the Ox1 axis, the Oy1 axis, and the Oz1 axis in the body coordinate system of the aircraft, respectively. x1 , ω y1 , ω z1 , and ω x4 , ω y4 , ω z4 can be transformed into each other through a conversion matrix.

[0044] (4) The conversion model between the measured angular velocity and the rigid body angular velocity of the rigid-elastic coupling rotating aircraft is established.

[0045] In the flight control system of the rigid-flexible coupling rotating aircraft, the attitude angular velocity directly measured by the sensor is composed of the rigid body motion attitude angular velocity and the additional angular velocity caused by the elastic deformation of the aircraft

[0046]

[0047] In the formula, ω y ′, ω z ′ are the y direction and the z direction attitude angular velocities directly measured by the sensor, ω y , ω z are the y direction and the z direction rigid body motion attitude angular velocities, x ω is the sensor installation position. ω y , ω z , and ω y4 , ω z4 are transformed into each other through a conversion matrix.

[0048] (5) Establish the state equation of the control object of the rigid-flexible coupled rotary aircraft.

[0049] The roll channel of the rotary aircraft is usually not controlled, only the pitch channel and the yaw channel are considered. Take [α * ,β * ,θ,ψ,ω y4 ,ω z4 ,η i ,ε i ,ω y ',ω z '] as the state variables, in which ω y ',ω z ' are directly measured by sensors, and the rest are state variables that need to be estimated. Based on the rigid body dynamics model of the rigid-flexible coupled rotary aircraft in the quasi-aircraft body coordinate system in step (2), the elastic deformation dynamics model of the rigid-flexible coupled rotary aircraft in the aircraft body coordinate system in step (3), and the conversion model between the measured angular velocity and the rigid body angular velocity of the rigid-flexible coupled rotary aircraft in step (4), the state equation of the control object of the rigid-flexible coupled rotary aircraft with [α * ,β * ,θ,ψ,ω y4 ,ω z4 ,η i ,ε i ,ω y ',ω z '] as the state variables can be established.

[0050] (6) Establish the observer of the rigid-flexible coupled rotary aircraft.

[0051] Based on the state equation of the control object of the rigid-flexible coupled rotary aircraft in step (5), the observer of the rigid-flexible coupled rotary aircraft is determined by using the general observer method, and the state variables [α * ,β * ,θ,ψ,ω y4 ,ω z4 ,η i ,ε i ] that need to be estimated are obtained.

[0052] (7) Establish the sliding mode control law of the rigid-flexible coupled rotary aircraft.

[0053] The sliding mode control law design mainly includes two steps: first, construct a suitable sliding surface, so that the system limited to the sliding surface has the desired dynamic characteristics. Then, design the control law so that the sliding surface is attractive to the system state when there are uncertainties and disturbances, and the trajectory of the system can reach the sliding surface in a finite time

[0054] The nonsingular terminal sliding surface S is designed as:

[0055] S = e1 + k0e2 p / q

[0056] wherein: k0, p, q are control parameters, k0 > 0, 1 < p / q < 2, e1, e2 are attitude angle error vector and attitude angular velocity error vector, p, q are positive odd numbers;

[0057] The control law based on the double power approaching law is determined as follows:

[0058]

[0059] wherein: u is the control quantity, b is the control parameter, k1, k2, a1, a2 are control parameters, k1 > 0, k2 > 0, a1 > 1, 0 < a2 < 1, sgn(S) is the sign function, f is the aircraft attitude dynamics, ω c is the desired attitude angular velocity vector.

[0060] (8) Establishing the rigid-flexible coupling rotating aircraft control law based on the sliding mode and the observer.

[0061] Based on the observer of the rigid-flexible coupling rotating aircraft in step (6) and the sliding mode control law of the rigid-flexible coupling rotating aircraft in step (7), the rigid-flexible coupling rotating aircraft control law based on the sliding mode and the observer is established, so as to realize the stable control of the rotating aircraft, as shown in Figure 2 When the sliding mode and the observer are integrated, the estimated elastic deformation is taken as the state variable to enter the control loop, or it is taken as a disturbance, and after the estimation by the observer, the disturbance estimation enters the control loop.

[0062] The contents not described in detail in the specification of the present application are the known technology of the person skilled in the art.

[0063] Although the present application has been disclosed with the above preferred embodiments, it is not intended to limit the present application, and any person skilled in the art can make possible changes and modifications to the technical solutions of the present application by using the disclosed methods and technical contents without departing from the spirit and scope of the present application. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present application, which does not depart from the technical solutions of the present application, belongs to the protection scope of the technical solutions of the present application.

Claims

1. A control method for a rigid-elastic coupled rotating aircraft based on sliding mode and an observer, characterized in that, include Given a rigid-elastic coupled rotating aircraft model; Establish a rigid body dynamics model of a rigid-elastic coupled rotating aircraft in the quasi-aircraft body coordinate system; Establish an elastic deformation dynamics model of a rigid-elastic coupled rotating aircraft in the aircraft body coordinate system; Establish a conversion model between the measured angular velocity and the rigid body angular velocity of a rigid-elastic coupled rotating aircraft; Establish the state equations for the control object of a rigid-elastic coupled rotating aircraft. Establish an observer for a rigid-elastic coupled rotating flight vehicle; Establish a sliding mode control law for a rigid-elastic coupled rotating aircraft; Establish a control law for a rigid-elastic coupled rotating aircraft based on sliding mode and observer to achieve stable control of the rotating aircraft; The rigid body dynamics model of the rigid-elastic coupled rotating aircraft in the quasi-aircraft body coordinate system is as follows: In the formula, α * It is the quasi-angle of attack, β * It is the quasi-sideslip angle, m is the mass of the rotating aircraft, J x J y J z These represent the moments of inertia of the rotating aircraft along the Ox4, Oy4, and Oz4 axes in the quasi-aircraft body coordinate system, respectively; V is the velocity of the rotating aircraft; and ω is the velocity of the rotating aircraft. x4 ,ω y4 ,ω z4 These represent the rotational angular velocities of the rotating aircraft along the Ox4, Oy4, and Oz4 axes in the quasi-aircraft body coordinate system, respectively; θ is the mass pitch angle of the rotating aircraft; γ is the roll angle of the rotating aircraft; ψ is the yaw angle of the rotating aircraft; and M... yc M zc These are the control torques of the rotating aircraft along the Oy4 and Oz4 axes in the quasi-aircraft body coordinate system, M. x4 M y4 M z4 These are the uncontrolled moments of the rotating aircraft along the Ox4, Oy4, and Oz4 axes in the quasi-aircraft body coordinate system, F. y4 ,F z4 These are the external forces acting on the Oy4 and Oz4 axes of the rotating aircraft in the quasi-aircraft body coordinate system; The elastic deformation dynamics model of a rigid-elastic coupled rotating aircraft in the aircraft body coordinate system is as follows: In the formula, η i ,ε i These are the generalized coordinates describing the deformation of the rotating aircraft in the y and z directions, respectively, ω i It is the natural frequency of the i-th mode shape of the aircraft, φ i (x) is the mode shape function, M i f is the generalized mass of the i-th mode shape of the aircraft. y (x),f z (x) represents the distributed forces acting on the infinitesimal element of the rotating aircraft shaft segment in the y and z directions, respectively, where x is the axial coordinate of the infinitesimal element, and μ i It is the damping ratio of the i-th mode of vibration of the aircraft, ω x1 ,ω y1 ,ω z1 These are the rotational angular velocities of the rotating aircraft along the Ox1, Oy1, and Oz1 axes in the aircraft's body coordinate system; ω x1 ,ω y1 ,ω z1 With ω x4 ,ω y4 ,ω z4 They can be transformed into each other using a transformation matrix; For the conversion model between the measured angular velocity and the rigid body angular velocity of a rigid-flexible coupled rotating aircraft, in the flight control system of the rigid-flexible coupled rotating aircraft, the attitude angular velocity directly measured by the sensor consists of the rigid body motion attitude angular velocity and the additional angular velocity caused by the elastic deformation of the aircraft: In the formula, ω y ′,ω z ' and ' represent the attitude angular velocities in the y and z directions, respectively, directly measured by the sensor, and ω... y ,ω z These are the rigid body motion attitude angular velocities in the y and z directions, respectively, and x... ω It refers to the sensor installation location; ω y ,ω z With ω y4 ,ω z4 Transform them using transformation matrices; The state equations for establishing the control object of a rigid-elastic coupled rotating aircraft include: The roll path of the rotating aircraft is not controlled; only the pitch and yaw paths are considered. [α] * ,β * ,θ,ψ,ω y4 ,ω z4 ,η i ,ε i ,ω y ′,ω z '] is used as a state variable, where ω y ′,ω z ' is the state variable directly measured by the sensor, and the rest are state variables that need to be estimated; based on the rigid body dynamics model of the rigid-elastic coupled rotating aircraft in the quasi-aircraft body coordinate system, the elastic deformation dynamics model of the rigid-elastic coupled rotating aircraft in the aircraft body coordinate system, and the conversion model between the measured angular velocity and the rigid body angular velocity of the rigid-elastic coupled rotating aircraft, a model is established with [α * ,β * ,θ,ψ,ω y4 ,ω z4 ,η i ,ε i ,ω y ′,ω z The state equation of the rigid-elastic coupled rotating aircraft control object as a state variable.

2. The control method for a rigid-elastic coupled rotating aircraft according to claim 1, characterized in that, The rigid-elastic coupling model of a rigid-elastic coupled rotating aircraft refers to the fact that rigid motion and elastic deformation are decoupled in terms of the degrees of freedom of motion. The coupling between rigidity and elasticity lies in the calculation of force and torque, and the elastic vibration signal of the aircraft enters the control loop through the measurement unit.

3. The control method for a rigid-elastic coupled rotating aircraft according to claim 1, characterized in that, The observers used to determine rigid-elastic coupled rotating vehicles include: Based on the state equations of the rigid-elastic coupled rotating aircraft control object, a general observer method is used to determine the observers of the rigid-elastic coupled rotating aircraft, and obtain the state variables [α] that need to be estimated. * ,β * ,θ,ψ,ω y4 ,ω z4 ,η i ,ε i ].

4. The control method for a rigid-elastic coupled rotating aircraft according to claim 1, characterized in that, Determining the sliding mode control law for a rigid-elastic coupled rotating aircraft includes: Construct a non-singular terminal sliding surface such that the system constrained to move on the sliding surface has the desired dynamic characteristics; determine a control law such that the sliding surface is still attractive to the system state in the presence of uncertainties and disturbances, and the system trajectory can reach the sliding surface in a finite time.

5. The control method for a rigid-elastic coupled rotating aircraft according to claim 4, characterized in that, The non-singular terminal sliding surface S is: S=e1+k0e2 p / q In the formula: k0, p, q are control parameters, k0 > 0, 1 < p / q < 2, e1 and e2 are attitude angle error vector and attitude angular velocity error vector, and p and q are positive odd numbers; The control law based on the double power-reaching law is determined as follows: In the formula: u is the control variable, b is the control parameter, k1, k2, a1, a2 are control parameters, k1 > 0, k2 > 0, a1 > 1, 0 < a2 < 1, sgn(S) is the sign function, f is the aircraft attitude dynamics, ω c It is the desired attitude angular velocity vector.

6. The control method for a rigid-elastic coupled rotating aircraft according to claim 1, characterized in that, When establishing a control law for a rigid-elastic coupled rotating aircraft based on sliding mode and observer, the estimated elastic deformation is used as a state variable and then enters the control loop when the sliding mode and observer are combined, or it is used as a disturbance, estimated by the observer, and then enters the control loop as a disturbance estimate.

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