CN108776485B - Self-adaptive control method of four-rotor aircraft based on arc tangent enhanced fast power approach law and fast terminal sliding mode surface - Google Patents
Self-adaptive control method of four-rotor aircraft based on arc tangent enhanced fast power approach law and fast terminal sliding mode surface Download PDFInfo
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Abstract
A self-adaptive control method of a four-rotor aircraft based on an arc tangent enhanced fast power approach law and a fast terminal sliding mode surface comprises the following steps: step 1, determining a transfer matrix from a body coordinate system based on a four-rotor aircraft to an inertial coordinate system based on the earth; step 2, analyzing a four-rotor aircraft dynamic model according to a Newton Euler formula; and 3, calculating a tracking error, and designing a controller according to the fast terminal sliding mode surface and the first derivative thereof. The method combines the enhanced fast power approximation law sliding mode control based on the arc tangent and the fast terminal sliding mode control, can increase the approximation speed when the sliding mode surface is far away, can reduce buffeting, improves the rapidity of the system, realizes fast and stable control, can realize limited time control of the tracking error, and solves the problem that the tracking error tends to 0 only when the time tends to infinity in the traditional sliding mode surface. Meanwhile, the interference boundary is estimated through self-adaptation, and the stability of the system is improved.
Description
Technical Field
The invention relates to a self-adaptive control method of a four-rotor aircraft based on an arc tangent enhanced fast power approach law and a fast terminal sliding mode surface.
Background
The four-rotor aircraft has attracted wide attention of domestic and foreign scholars and scientific research institutions due to the characteristics of simple structure, strong maneuverability and unique flight mode, and is rapidly one of the hotspots of international research at present. Compared with a fixed-wing aircraft, the rotary-wing aircraft can vertically lift, has low requirement on the environment, does not need a runway, reduces the cost and has great commercial value. The development of aircrafts makes many dangerous high-altitude operations easy and safe, so as to cause deterrence to other countries in the military aspect and greatly increase the working efficiency in the civil aspect. The four-rotor aircraft has strong flexibility, can realize rapid transition of motion and hovering at any time, and can be competent for more challenging flight tasks with less damage risk. In the field of scientific research, because a four-rotor aircraft has the dynamic characteristics of nonlinearity, under-actuation and strong coupling, researchers often use the four-rotor aircraft as an experimental carrier for theoretical research and method verification. An aircraft flight control system is built by relying on a small four-rotor aircraft to carry out high-performance motion control research on the aircraft, and the method is a hot research field of the current academic world.
The approach law sliding mode control has the characteristics that discontinuous control can be realized, the sliding mode is programmable and is not related to system parameters and disturbance. The approach law sliding mode not only can reasonably design the speed of reaching the sliding mode surface, reduce the time of the approach stage, improve the robustness of the system, but also can effectively weaken the buffeting problem in the sliding mode control. Currently, in the field of four-rotor control, approach law sliding mode control is less used. The enhanced approach law further accelerates the approach speed of the system to the sliding mode surface and simultaneously enables the buffeting to be smaller on the basis of the traditional approach law. Because the four-rotor aircraft can encounter external environment interference in flight, interference and compensation are carried out on the interference boundary through self-adaptation, and the stability of the system is improved.
Disclosure of Invention
In order to solve the problems that the traditional sliding mode surface can not realize limited time control, further accelerate the approaching speed of an approaching law and reduce buffeting, the method adopts the rapid terminal sliding mode control and the fast power approaching law based on the arc tangent enhancement mode, avoids the singularity problem through the switching control idea, accelerates the approaching speed of a system to the sliding mode surface, reduces buffeting and realizes the limited time control. Meanwhile, interference and compensation are carried out on the interference boundary through self-adaption, and the stability of the system is improved.
The technical scheme proposed for solving the technical problems is as follows:
a self-adaptive control method of a four-rotor aircraft based on an arc tangent enhanced fast power approach law and a fast terminal sliding mode surface comprises the following steps:
where psi, theta, phi are the yaw, pitch, roll angles of the aircraft, respectively, representing the angle of rotation of the aircraft about each axis of the inertial frame in sequence, TψTransition matrix, T, representing psiθA transition matrix, T, representing thetaφA transition matrix representing phi;
2.1, the translation process comprises the following steps:
wherein x, y and z respectively represent the position of the four rotors under an inertial coordinate system, m represents the mass of the aircraft, g represents the gravity acceleration, mg represents the gravity borne by the four rotors, and the resultant force U generated by the four rotorsr；
2.2, the rotation process comprises the following steps:
wherein tau isx、τy、τzRespectively representing the axial moment components, I, in the coordinate system of the machine bodyxx、Iyy、IzzRespectively representing the rotational inertia component of each axis on the coordinate system of the machine body, x represents cross product, wp、wq、wrRespectively representing the attitude angular velocity components of each axis on the coordinate system of the body,
respectively represent machine coordinate systemThe attitude angular acceleration component of each axis;
considering that the aircraft is in a low-speed flight or hovering state, consider
2.3, connecting the vertical type (1), (2) and (4), and obtaining the dynamic model of the aircraft as shown in the formula (5)
Wherein
Ux、Uy、UzThe input quantities of the three position controllers are respectively;
according to the formula (5), decoupling calculation is carried out on the position and posture relation, and the result is as follows:
wherein phidIs the desired signal value of phi, thetadDesired signal value of theta, psidFor desired signal values of ψ, the arcsin function is an arcsine function and the arctan function is an arctangent function;
equation (5) can also be written in matrix form as follows:
wherein X1＝[x,y,z,φ,θ,ψ]T,
B(X)＝diag(1,1,1,b1,b2,b3),U＝[Ux,Uy,Uz,τx,τy,τz]T，
3.1, defining the tracking error and its first and second differentials:
e＝X1-Xd (8)
wherein, Xd＝[xd,yd,zd,φd,θd,ψd]T，xd,yd,zd,φd,θd,ψdThe conductive desired signals are x, y, z, phi, theta, psi, respectively,
i＝1，2,3,4,5,6，Di，c0i，c1i，c2i，ei，
respectively corresponding ith element;
3.2, designing a quick terminal sliding mode surface:
wherein, sigα(x)＝|x|α·sign(x)，α1＞α2＞1，λ1＞0，λ2＞0；
Derivation of equation (11) yields:
order to
Formula (12) is simplified to formula (13)
But due to the presence of alpha (e)
When α (e) is 0 and β (e) is not equal to 0, the negative power term of (a) causes a singularity problem;
consider the method of handover control:
wherein q isi(e),αi(e),βi(e) Q (e), alpha (e), beta (e) respectively,
combining formula (13) and formula (14) to obtain:
conjunctive formula (7), formula (10) and formula (15) yields:
3.3 design enhanced approach law
Wherein
N-1(X) is the inverse of N (X), k1＞0，k2＞0，0＜β1Less than 1, delta is more than 0 and less than 1, gamma is more than 0, mu is more than 1, and p is a positive integer;
3.4, combined vertical (16) and formula (17), to obtain a controller
Wherein B is-1(X) is the inverse of B (X),
respectively corresponding ith element;
the adaptive law is designed as follows:
when the system moves away from the sliding mode, | s | is large, N(s) approaches δ,
the approach speed of the system is accelerated; when the system approaches the sliding mode, | s | approaches 0, N(s) approaches μ,
the buffeting of the system is reduced.
The technical conception of the invention is as follows: aiming at a four-rotor aircraft system, a self-adaptive control method of the four-rotor aircraft based on an arc tangent enhanced type rapid power approach law and a rapid terminal sliding mode surface is designed by combining rapid power approach law sliding mode control and rapid terminal sliding mode control. The quick terminal sliding mode surface can realize the limited time control of the tracking error, and solves the problems that the time tends to be infinite and the error tends to be 0 in the traditional sliding mode surface. Based on the arctangent enhanced approach law, the approach speed can be increased when the sliding mode surface is far away, buffeting can be reduced, the rapidness and the robustness of the system are improved, and rapid and stable control is realized. Meanwhile, interference and compensation are carried out on the interference boundary through self-adaption, and the stability of the system is improved.
The invention has the beneficial effects that: compared with the traditional fast power approximation law sliding mode control, the method can increase the approximation speed when the system is far away from the sliding mode, reduce buffeting and shorten the arrival time of the sliding mode, thereby enabling the system to realize stable convergence more quickly. In addition, the invention utilizes the quick terminal sliding mode, solves the problems that the time tends to be infinite and the error tends to be 0 in the traditional sliding mode surface, and realizes the limited time control. Meanwhile, interference and compensation are carried out on the interference boundary through self-adaption, and the stability of the system is improved.
Drawings
Fig. 1 is a schematic diagram of a position tracking effect of a four-rotor aircraft, wherein a dotted line represents "1" type enhanced fast power approach law adaptive control under a linear sliding mode surface, and a dotted line represents inverse tangent "mu" type enhanced fast power approach law adaptive control under a fast terminal sliding mode surface.
Fig. 2 is a schematic diagram of the attitude tracking effect of a four-rotor aircraft, wherein a dotted line represents the enhanced type 1 fast power approach law adaptive control of a linear sliding mode surface, and a dotted line represents the enhanced type fast power approach law adaptive control based on an arctangent type mu of the fast terminal sliding mode surface.
Fig. 3 is a schematic input diagram of a position controller for enhanced type 1 fast power-law adaptive control under a linear sliding mode surface of a four-rotor aircraft.
Fig. 4 is an input schematic diagram of a position controller for self-adaptive control of a sliding mode surface of a fast terminal of a four-rotor aircraft based on an arctangent mu-type enhanced fast power approach law.
Fig. 5 is an input schematic diagram of an attitude controller for "1" -type enhanced fast power-law adaptive control under a linear sliding-mode surface of a four-rotor aircraft.
Fig. 6 is an input schematic diagram of an attitude controller for self-adaptive control of a sliding mode surface of a fast terminal of a four-rotor aircraft based on an arctangent mu-type enhanced fast power approach law.
Fig. 7 is a schematic diagram of local amplification of input of an attitude controller for "1" -type enhanced fast power-law adaptive control under a linear sliding-mode surface of a four-rotor aircraft.
Fig. 8 is a schematic diagram of local amplification of input of an attitude controller of a sliding mode surface of a fast terminal of a four-rotor aircraft based on arctangent 'mu' type enhanced fast power approach law adaptive control.
Fig. 9 is an estimation of the bounds of the position disturbance of the sliding mode surface of the fast terminal of the four-rotor aircraft based on the arctangent 'mu' type enhanced fast power approach law adaptive control.
Fig. 10 is an estimation of the boundary of attitude disturbance of a four-rotor aircraft fast terminal sliding mode surface based on arctangent 'mu' type enhanced fast power approach law adaptive control.
FIG. 11 is a control flow diagram of the present invention.
Detailed Description
The invention is further described below with reference to the accompanying drawings.
Referring to fig. 1-11, a self-adaptive control method of a four-rotor aircraft based on an arc tangent enhanced fast power approach law and a fast terminal sliding mode surface includes the following steps:
where psi, theta, phi are the yaw, pitch, roll angles of the aircraft, respectively, representing the angle of rotation of the aircraft about each axis of the inertial frame in sequence, TψTransition matrix, T, representing psiθA transition matrix, T, representing thetaφA transition matrix representing phi;
2.1, the translation process comprises the following steps:
wherein x, y and z respectively represent the position of the four rotors under an inertial coordinate system, m represents the mass of the aircraft, g represents the gravity acceleration, mg represents the gravity borne by the four rotors, and the resultant force U generated by the four rotorsr；
2.2, the rotation process comprises the following steps:
wherein tau isx、τy、τzRespectively representing the axial moment components, I, in the coordinate system of the machine bodyxx、Iyy、IzzRespectively representing the rotational inertia component of each axis on the coordinate system of the machine body, x represents cross product, wp、wq、wrRespectively representing the attitude angular velocity components of each axis on the coordinate system of the body,
respectively representing the attitude angular acceleration components of all axes on the coordinate system of the machine body;
considering that the aircraft is in a low-speed flight or hovering state, consider
2.3, connecting the vertical type (1), (2) and (4), and obtaining the dynamic model of the aircraft as shown in the formula (5)
Wherein
Ux、Uy、UzThe input quantities of the three position controllers are respectively;
according to the formula (5), decoupling calculation is carried out on the position and posture relation, and the result is as follows:
wherein phidIs the desired signal value of phi, thetadDesired signal value of theta, psidFor desired signal values of ψ, the arcsin function is an arcsine function and the arctan function is an arctangent function;
equation (5) can also be written in matrix form as follows:
wherein X1＝[x,y,z,φ,θ,ψ]T,
B(X)＝diag(1,1,1,b1,b2,b3),U＝[Ux,Uy,Uz,τx,τy,τz]T，
3.1, defining the tracking error and its first and second differentials:
e＝X1-Xd (8)
wherein, Xd＝[xd,yd,zd,φd,θd,ψd]T，xd,yd,zd,φd,θd,ψdThe conductive desired signals are x, y, z, phi, theta, psi, respectively,
i＝1，2,3,4,5,6，Di，c0i，c1i，c2i，ei，
respectively corresponding ith element;
3.2, designing a quick terminal sliding mode surface:
wherein, sigα(x)＝|x|α·sign(x)，α1＞α2＞1，λ1＞0，λ2＞0；
Derivation of equation (11) yields:
order to
Formula (12) is simplified to formula (13)
But due to the presence of alpha (e)
When α (e) is 0 and β (e) is not equal to 0, the negative power term of (a) causes a singularity problem;
consider the method of handover control:
wherein q isi(e),αi(e),βi(e) Q (e), alpha (e), beta (e) respectively,
combining formula (13) and formula (14) to obtain:
conjunctive formula (7), formula (10) and formula (15) yields:
3.3 design enhanced approach law
Wherein
N-1(X) is the inverse of N (X), k1＞0，k2＞0，0＜β1Less than 1, delta is more than 0 and less than 1, gamma is more than 0, mu is more than 1, and p is a positive integer;
3.4, combined vertical (16) and formula (17), to obtain a controller
Wherein B is-1(X) is the inverse of B (X),
respectively corresponding ith element;
the adaptive law is designed as follows:
when the system moves away from the sliding mode, | s | is large, N(s) approaches δ,
the approach speed of the system is accelerated; when the system approaches the sliding mode, | s | approaches 0, N(s) approaches μ,
the buffeting of the system is reduced.
In order to verify the effectiveness of the method, the invention provides a comparison between a sliding mode control method of a fast terminal sliding mode surface based on an arc tangent mu-type enhanced fast power approach law and a sliding mode control method of a linear sliding mode surface 1-type enhanced fast power approach law:
wherein the 1-type enhanced fast power approach law is
For more efficient comparison, all parameters of the system are consistent, i.e. xd＝yd＝zd＝2、ψd0.5, fast terminal sliding mode surface parameters: lambda [ alpha ]1＝0.5、λ2＝2、α1＝2、α21.1, epsilon 0.3, linear slip-form face: lambda [ alpha ]10.5, "μ" type enhanced proximity law parameter: k is a radical of1＝1、k2＝10、δ＝0.1、p＝1、γ＝1、μ＝10、β10.7, enhanced approximation rule parameter of "1": k is a radical of1＝1、k2＝10、δ＝0.1、p＝1、γ＝1、β10.7, adaptive initial value setting
p0i＝p1i＝p2i＝0.1、ε0i＝ε1i＝ε2i0.001, 1,2,3,4,5,6, interference parameter: dx＝dy＝dz＝0.2sin(0.2t)、
Parameters of the four-rotor aircraft: 1.1 and Ixx＝1.22、Iyy＝1.22、Izz2.2, g 9.81, sampling parameters: t is ts＝0.007，N＝5000。
As can be seen from fig. 1 and 2, the adaptive control of the quadrotor aircraft based on the arctangent enhanced fast power approach law and the fast terminal sliding mode surface can reach the expected position faster; with reference to fig. 3-8, the adaptive control of the quadrotor aircraft based on the arctangent enhanced fast power approach law and the fast terminal sliding mode surface has smaller buffeting. Fig. 9 and 10 can see the effectiveness of the estimation of the adaptive epipolar.
In conclusion, the self-adaptive control of the four-rotor aircraft based on the arctangent enhanced fast power approach law and the fast terminal sliding mode surface can reduce the buffeting, reduce the tracking time, improve the tracking performance and enable the system to enter stable convergence more quickly.
While the foregoing has described a preferred embodiment of the invention, it will be appreciated that the invention is not limited to the embodiment described, but is capable of numerous modifications without departing from the basic spirit and scope of the invention as set out in the appended claims.
Claims (1)
1. A self-adaptive control method of a four-rotor aircraft based on an arc tangent enhanced fast power approach law and a fast terminal sliding mode surface is characterized by comprising the following steps:
step 1, determining a transfer matrix from a body coordinate system based on a four-rotor aircraft to an inertial coordinate system based on the earth;
wherein psi, theta and phi are respectively the yaw angle, pitch angle and roll angle of the aircraft, and represent the angle of the aircraft sequentially rotating around each axis of the inertial coordinate system, and TψTransition matrix, T, representing psiθA transition matrix, T, representing thetaφA transition matrix representing phi;
step 2, analyzing a four-rotor aircraft dynamic model according to a Newton Euler formula, wherein the process is as follows:
2.1, the translation process comprises the following steps:
wherein x, y and z respectively represent the position of the four rotors under an inertial coordinate system, m represents the mass of the aircraft, g represents the acceleration of gravity, mg represents the gravity borne by the four rotors, and the sum of the four rotors is generatedForce Ur；
2.2, the rotation process comprises the following steps:
wherein tau isx、τy、τzRespectively representing the axial moment components, I, in the coordinate system of the machine bodyxx、Iyy、IzzRespectively representing the rotational inertia component of each axis on the coordinate system of the machine body, x represents cross product, wp、wq、wrRespectively representing the attitude angular velocity components of each axis on the coordinate system of the body,
respectively representing the attitude angular acceleration components of all axes on the coordinate system of the machine body;
considering that the aircraft is in a low-speed flight or hovering state, consider
2.3, connecting the vertical type (1), (2) and (4), and obtaining the dynamic model of the aircraft as shown in the formula (5)
Wherein
Ux、Uy、UzThe input quantities of the three position controllers are respectively;
according to the formula (5), decoupling calculation is carried out on the position and posture relation, and the result is as follows:
wherein phidIs the desired signal value of phi, thetadDesired signal value of theta, psidFor desired signal values of ψ, the arcsin function is an arcsine function and the arctan function is an arctangent function;
considering further the case where interference exists, equation (5) can be written in a matrix form as follows:
wherein X1＝[x,y,z,φ,θ,ψ]T,
B(X)＝diag(1,1,1,b1,b2,b3),U＝[Ux,Uy,Uz,τx,τy,τz]T，
Step 3, calculating a tracking error, and designing a controller according to the fast terminal sliding mode surface and the first derivative thereof, wherein the process is as follows:
3.1, defining the tracking error and its first and second differentials:
e＝X1-Xd (8)
wherein, Xd＝[xd,yd,zd,φd,θd,ψd]T，xd,yd,zd,φd,θd,ψdThe conductive desired signals are x, y, z, phi, theta, psi, respectively,
Di，c0i，c1i，c2i，ei，
respectively corresponding ith element;
3.2, designing a quick terminal sliding mode surface:
wherein, sigα(x)＝|x|α·sign(x)，α1＞α2＞1，λ1＞0，λ2＞0；
Derivation of equation (11) yields:
order to
Formula (12) is simplified to formula (13)
But because of
In existence of
When α (e) is 0 and β (e) is not equal to 0, the negative power term of (a) causes a singularity problem;
consider the method of handover control:
wherein q isi(e),αi(e),βi(e) Q (e), α (e), β (e), i ═ 1,2,3,4,5, 6;
combining formula (13) and formula (14) to obtain:
conjunctive formula (7), formula (10) and formula (15) yields:
3.3 design enhanced approach law
Wherein
N-1(X) is the inverse of N (X), k1＞0，k2＞0，0＜β1Less than 1, delta is more than 0 and less than 1, gamma is more than 0, mu is more than 1, and p is a positive integer;
3.4, combined vertical (16) and formula (17), to obtain a controller
Wherein B is-1(X) is the inverse of B (X),
respectively corresponding ith element;
the adaptive law is designed as follows:
step 4, enhanced property specification, the process is as follows:
when the system moves away from the sliding mode, | s | is large, N(s) approaches δ,
the approach speed of the system is accelerated; when the system approaches the sliding mode, | s | approaches 0, N(s) approaches μ,
the buffeting of the system is reduced.
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Application publication date: 20181109 Assignee: Tongxiang Tujian Intelligent Technology Co.,Ltd. Assignor: JIANG University OF TECHNOLOGY Contract record no.: X2023980037545 Denomination of invention: Adaptive Control Method for Quadcopter Aircraft Based on Arctangent Enhanced Fast Power Reaching Law and Fast Terminal Sliding Surface Granted publication date: 20210803 License type: Common License Record date: 20230705 |