CN113110554A - Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method - Google Patents

Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method Download PDF

Info

Publication number
CN113110554A
CN113110554A CN202110483380.9A CN202110483380A CN113110554A CN 113110554 A CN113110554 A CN 113110554A CN 202110483380 A CN202110483380 A CN 202110483380A CN 113110554 A CN113110554 A CN 113110554A
Authority
CN
China
Prior art keywords
channel
aerial vehicle
unmanned aerial
derivative
sliding mode
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
CN202110483380.9A
Other languages
Chinese (zh)
Other versions
CN113110554B (en
Inventor
赵振华
肖亮
梅劲松
曹东
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Nanjing University of Aeronautics and Astronautics
Original Assignee
Nanjing University of Aeronautics and Astronautics
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Nanjing University of Aeronautics and Astronautics filed Critical Nanjing University of Aeronautics and Astronautics
Priority to CN202110483380.9A priority Critical patent/CN113110554B/en
Publication of CN113110554A publication Critical patent/CN113110554A/en
Application granted granted Critical
Publication of CN113110554B publication Critical patent/CN113110554B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft

Landscapes

  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Abstract

The invention discloses a composite continuous rapid terminal sliding mode attitude control method for a four-rotor unmanned aerial vehicle, which comprises the steps of firstly establishing a disturbed dynamics model of an attitude loop of the four-rotor unmanned aerial vehicle, and then establishing a rolling, pitching and yawing three-channel decoupling equation of attitude angle tracking error of the four-rotor unmanned aerial vehicle; then designing an interference observer aiming at three channels of rolling, pitching and yawing based on an extended state observer algorithm to realize the estimation of three-channel lumped interference; and finally, constructing a composite continuous rapid terminal sliding mode controller of the attitude loop based on a rapid terminal sliding mode algorithm and the lumped interference estimation information. The invention avoids the phenomenon of buffeting; the interference resistance of the control system is obviously improved through the estimation and feedforward compensation of the interference. Compared with the traditional sliding mode control method, the method has the advantages of higher convergence speed and stronger anti-interference performance, and effectively inhibits the influence of multi-source interference on the attitude control performance of the quad-rotor unmanned aerial vehicle.

Description

Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method
Technical Field
The invention belongs to the technical field of flight control.
Background
The quad-rotor unmanned aerial vehicle has simple structure, compact size and higher flexibility, and can be widely applied to various fields such as civil use, military use and the like. The quad-rotor unmanned aerial vehicle realizes tracking of a specific track by adjusting the attitude so as to execute a flight task, so that high-precision attitude tracking control is the key of design of a quad-rotor unmanned aerial vehicle flight control system. The developments of four rotor unmanned aerial vehicle attitude systems are the system of an essence nonlinearity, strong coupling, multiple input multiple output to can receive the influence of multisource interference such as external environment interference, inside pneumatic parameter perturbation, unmodeled developments among the flight process, these factors bring huge challenge for four rotor unmanned aerial vehicle attitude control system design.
The sliding mode control algorithm is widely applied to the design of the attitude control system of the quad-rotor unmanned aerial vehicle due to the simple design and strong robustness. When an attitude controller is designed by the conventional sliding mode control algorithm, multi-source interference is usually inhibited by self robustness, so that the influence of the multi-source interference cannot be quickly inhibited; in addition, because the influence of interference needs to be suppressed by means of robustness of the existing method, a sign function is generally adopted as a switching term in the existing method, and then a serious buffeting problem is caused. Therefore, it is necessary to provide an attitude system control method with strong anti-interference capability and fast convergence performance under the condition of ensuring the continuity of the system control quantity.
Disclosure of Invention
The purpose of the invention is as follows: in order to solve the problems in the prior art, the invention provides a composite continuous rapid terminal sliding mode attitude control method for a quad-rotor unmanned aerial vehicle.
The technical scheme is as follows: the invention provides a composite continuous rapid terminal sliding mode attitude control method for a quad-rotor unmanned aerial vehicle, which specifically comprises the following steps:
s1, establishing a disturbed dynamics model of the attitude loop system of the quad-rotor unmanned aerial vehicle;
s2 moments tau acting on the x, y and z axes of the fuselage coordinate systemx、τyAnd τzThree-channel decoupling is carried out to obtain three-channel virtual torque, a three-channel decoupling equation of attitude angle tracking error dynamic of the quad-rotor unmanned aerial vehicle is established based on the three-channel virtual torque, and the three channels compriseA roll channel, a pitch channel and a yaw channel;
s3, designing a roll channel, a pitch channel and a yaw channel expansion state observer according to the three-channel decoupling equation in the S2;
s4, designing a four-rotor unmanned aerial vehicle attitude loop fast terminal sliding mode surface;
and S5, performing inverse solution on the three-channel virtual moment in S2 to obtain the actual control quantity of the attitude system of the four-rotor unmanned aerial vehicle, and designing a three-channel composite rapid terminal sliding mode virtual controller to realize the attitude control of the four-rotor unmanned aerial vehicle by combining interference estimation information observed by a three-channel extended state observer in S3 and a rapid terminal sliding mode surface of an attitude loop of the four-rotor unmanned aerial vehicle in S4 according to a three-channel decoupling equation of the attitude angle tracking error dynamics of the unmanned aerial vehicle in S2.
Further, in S1, the disturbed dynamics model of the quad-rotor drone attitude loop system is:
Figure BDA0003050014040000021
wherein the content of the first and second substances,
Figure BDA0003050014040000022
is the first derivative of theta and is,
Figure BDA0003050014040000023
phi denotes the roll angle of the quad-rotor drone, theta denotes the pitch angle of the quad-rotor drone, psi denotes the yaw angle of the quad-rotor drone,
Figure BDA0003050014040000024
sφis sin phi, cφIs cos phi, cθIs cos θ, tθIs a function of the number of tan theta,
Figure BDA0003050014040000025
wherein wxRepresenting angular velocity of rotation, w, of quad-rotor drone about the x-axis of the rectangular coordinate system of the fuselageyRepresenting the y-axis of a quad-rotor drone around a rectangular coordinate system of the fuselageAngular velocity of rotation, wzRepresenting the rotation angular velocity of the quad-rotor unmanned aerial vehicle around the z axis of the rectangular coordinate system of the fuselage;
Figure BDA0003050014040000026
is the first derivative of the omega and,
Figure BDA0003050014040000027
Jx，Jyand JzRespectively represents the rotational inertia of the quadrotor unmanned plane around the x axis, the y axis and the z axis of the rectangular coordinate system of the fuselage,
Figure BDA0003050014040000028
τx、τyand τzRespectively representing the moments of an x axis, a y axis and a z axis of a rectangular coordinate system of the fuselage;
Figure BDA0003050014040000029
Dx，Dyand DzThe lumped disturbances in the x-axis direction, the y-axis direction and the z-axis direction of the rectangular coordinate system of the fuselage are respectively represented.
Further, the three-channel decoupling equation in S2 is:
Figure BDA0003050014040000031
Figure BDA0003050014040000032
Figure BDA0003050014040000033
wherein the content of the first and second substances,
Figure BDA0003050014040000034
tracking error e for roll angle commandφSecond derivative of,
Figure BDA0003050014040000035
Error e is tracked for pitch angle commandθSecond derivative of,
Figure BDA0003050014040000036
Tracking error e for yaw angle commandψThe second derivative of (a); e.g. of the typeφ＝φ-φd、eθ＝θ-θd，eψ＝ψ-ψd，φd、θdAnd psidRespectively, expected instructions of a roll angle, a pitch angle and a yaw angle; tau isφ、τθ、τψRepresenting the virtual moments, f, of the roll, pitch and yaw channels, respectivelyA φ、fA θ、fA ψRespectively represents cross-coupling nonlinear terms in the decoupling equations of three channels of rolling, pitching and yawing,
Figure BDA0003050014040000037
representing lumped disturbances, τ, in roll, pitch and yaw triple-channel decoupling equations, respectivelyφ、τθ、τψ，fA φ、fA θ、fA ψAnd
Figure BDA0003050014040000038
the expression of (a) is:
Figure BDA0003050014040000039
wherein
Figure BDA00030500140400000310
Is the first derivative of W and is,
Figure BDA00030500140400000311
is thetadSecond derivative of (theta)d＝[φd θd ψd]TFor the desired attitude angle, T represents the matrix transpose, J-1Is the inverse matrix of J.
Further, in S3, specifically, the method includes:
establishing a rolling channel extended state observer for estimation
Figure BDA00030500140400000312
Figure BDA00030500140400000313
Wherein the content of the first and second substances,
Figure BDA00030500140400000314
in order to roll the channel lumped interference estimates,
Figure BDA00030500140400000315
and
Figure BDA00030500140400000316
to roll the channel to expand the internal dynamics of the state observer,
Figure BDA00030500140400000317
is composed of
Figure BDA00030500140400000318
The first derivative of (a) is,
Figure BDA00030500140400000319
is composed of
Figure BDA00030500140400000320
The first derivative of (a) is,
Figure BDA00030500140400000321
is composed of
Figure BDA00030500140400000322
The first derivative of (a);
Figure BDA00030500140400000323
and
Figure BDA00030500140400000324
extending state observer gain for a roll channel
Figure BDA00030500140400000325
And
Figure BDA00030500140400000326
values are all normal numbers and satisfy characteristic polynomial
Figure BDA00030500140400000327
s is a variable;
establishing a pitching channel extended state observer, estimating
Figure BDA00030500140400000328
Figure BDA0003050014040000041
Wherein the content of the first and second substances,
Figure BDA0003050014040000042
to aggregate the interference estimate for the pitch channel,
Figure BDA0003050014040000043
the internal dynamics of the state observer are expanded for the pitch channel,
Figure BDA0003050014040000044
is composed of
Figure BDA0003050014040000045
The first derivative of (a) is,
Figure BDA0003050014040000046
is composed of
Figure BDA0003050014040000047
The first derivative of (a) is,
Figure BDA0003050014040000048
is composed of
Figure BDA0003050014040000049
The first derivative of (a) is,
Figure BDA00030500140400000410
and
Figure BDA00030500140400000411
the state observer gain is extended for the pitch channel,
Figure BDA00030500140400000412
and
Figure BDA00030500140400000413
values are all normal numbers and satisfy characteristic polynomial
Figure BDA00030500140400000414
Establishing a yaw channel extended state observer, estimating
Figure BDA00030500140400000415
Figure BDA00030500140400000416
Wherein the content of the first and second substances,
Figure BDA00030500140400000417
to aggregate the disturbance estimate for the yaw channel,
Figure BDA00030500140400000418
the state observer internal dynamics are extended for the yaw channel,
Figure BDA00030500140400000419
is composed of
Figure BDA00030500140400000420
The first derivative of (a) is,
Figure BDA00030500140400000421
is composed of
Figure BDA00030500140400000422
The first derivative of (a) is,
Figure BDA00030500140400000423
is composed of
Figure BDA00030500140400000424
The first derivative of (a) is,
Figure BDA00030500140400000425
and
Figure BDA00030500140400000426
the state observer gain is extended for the yaw path,
Figure BDA00030500140400000427
and
Figure BDA00030500140400000428
are all normal numbers and satisfy characteristic polynomials
Figure BDA00030500140400000429
Further, the four-rotor unmanned aerial vehicle attitude loop fast terminal sliding mode surface is:
rolling channel rapid terminal sliding mode surface sigmaφComprises the following steps:
Figure BDA00030500140400000430
wherein alpha is1，β1，h1，g1，p1And q is1Are all positive odd numbers, and alpha1＞0,β1＞0,
Figure BDA00030500140400000431
Figure BDA00030500140400000432
Is eφThe first derivative of (a);
pitching channel rapid terminal sliding mode surface sigmaθComprises the following steps:
Figure BDA0003050014040000051
wherein alpha is2，β2，h2,g2,p2And q is2Are all positive odd numbers, and alpha2＞0,β2＞0，
Figure BDA0003050014040000052
Figure BDA0003050014040000053
Is eθThe first derivative of (a);
yaw channel rapid terminal sliding mode surface sigmaψComprises the following steps:
Figure BDA0003050014040000054
wherein alpha is3，β3，h3，g3，p3And q is3Are all positive odd numbers, and alpha3＞0,β3＞0,
Figure BDA0003050014040000055
Figure BDA0003050014040000056
Is eψThe first derivative of (a).
Further, the three-channel composite rapid terminal sliding mode attitude controller in the S5 includes a roll channel virtual moment controller, a pitch channel virtual moment controller and a yaw channel virtual moment controller;
the rolling channel virtual torque controller comprises:
Figure BDA0003050014040000057
wherein
Figure BDA0003050014040000058
Is a positive real number, m1,n1Is positive odd number, and
Figure BDA0003050014040000059
the pitching channel virtual moment controller comprises:
Figure BDA00030500140400000510
wherein
Figure BDA00030500140400000511
Is a positive real number, m2,n2Is positive odd number, and
Figure BDA00030500140400000512
the yaw channel virtual moment controller is as follows:
Figure BDA00030500140400000513
wherein
Figure BDA00030500140400000514
Is a positive real number, m3,n3Is positive odd number, and
Figure BDA00030500140400000515
has the advantages that:
(1) according to the method, a three-channel decoupling equation of rolling, pitching and yawing of the attitude loop of the disturbed quad-rotor unmanned aerial vehicle is established, a three-channel extended state observer is designed according to the equation to realize estimation of lumped interference of each channel, a composite sliding mode controller is designed based on estimation information to realize compensation of the lumped interference, and the anti-interference performance of a closed-loop system is obviously improved.
(2) Compared with the traditional sliding mode control method, the method has higher convergence rate.
(3) The sliding mode controller is designed by replacing a symbolic function with an attractor, so that discontinuous terms in control quantity are completely eliminated, system buffeting is avoided, and steady-state fluctuation of tracking errors is reduced.
Drawings
FIG. 1 is a block diagram of a composite continuous fast terminal sliding mode attitude control strategy proposed by the present invention;
FIG. 2 is a response curve of a roll angle tracking error of an attitude system of a disturbed quad-rotor unmanned aerial vehicle and a response curve of a control quantity of the channel, which are obtained by adopting the method of the invention; wherein (a) is a rolling angle tracking error response curve and (b) is a control quantity response curve of the channel;
FIG. 3 is a response curve of a pitch angle tracking error of an attitude system of a disturbed quad-rotor unmanned aerial vehicle and a response curve of a control quantity of the channel, which are obtained by the method of the present invention; wherein (a) is a pitch angle tracking error response curve diagram, and (b) is a control quantity response curve diagram of the channel;
FIG. 4 is a response curve of a yaw angle tracking error of an attitude system of a disturbed quad-rotor unmanned aerial vehicle and a response curve of a control quantity of the channel, which are obtained by adopting the method of the invention; wherein (a) is a yaw angle tracking error response curve diagram, and (b) is a channel control quantity response curve diagram.
Detailed Description
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate an embodiment of the invention and, together with the description, serve to explain the invention and not to limit the invention.
The invention is concretely realized as follows: firstly, a disturbed attitude loop model of the quad-rotor unmanned aerial vehicle is built by using a Simulink toolbox in simulation software MATLAB R2015b, and then simulation and experiment are carried out; fig. 1 is a structural block diagram of a sliding mode attitude control strategy of a composite continuous and rapid terminal provided by the invention. The invention provides a composite continuous rapid terminal sliding mode attitude control method for a quad-rotor unmanned aerial vehicle, which comprises the following specific steps of:
s1, establishing a disturbed dynamics model of the quad-rotor unmanned aerial vehicle attitude loop system.
S2 moments tau acting on the x, y and z axes of the fuselage coordinate systemx、τyAnd τzAnd performing three-channel decoupling to obtain three-channel virtual torque, and establishing a three-channel decoupling equation of attitude angle tracking error dynamics of the quad-rotor unmanned aerial vehicle based on the three-channel virtual torque.
S4, designing a four-rotor unmanned aerial vehicle attitude loop quick terminal sliding mode surface.
And S5, reversely solving the three-channel virtual moment in the S2 to obtain the actual control quantity of the four-rotor unmanned plane attitude system. A sliding mode attitude controller of a compound continuous and rapid terminal of a quad-rotor unmanned aerial vehicle is designed.
The specific operation steps of step S1 include:
the four-rotor unmanned aerial vehicle attitude loop disturbed dynamics model is described as shown in formula 1:
Figure BDA0003050014040000071
Figure BDA0003050014040000072
wherein phi represents the roll angle of the quad-rotor unmanned aerial vehicle, theta represents the pitch angle of the quad-rotor unmanned aerial vehicle, psi represents the yaw angle of the quad-rotor unmanned aerial vehicle,
Figure BDA0003050014040000073
is the first derivative of phi and is,
Figure BDA0003050014040000074
is the first derivative of the theta and is,
Figure BDA0003050014040000075
is the first derivative of ψ; w is ax，wyAnd wzEach represents fourRotational angular velocities of the rotorcraft about the x, y, and z axes;
Figure BDA0003050014040000076
is wxThe first derivative of (a) is,
Figure BDA0003050014040000077
is wyThe first derivative of (a) is,
Figure BDA0003050014040000078
is wzThe first derivative of (a); j. the design is a squarex，JyAnd JzRespectively representing the rotational inertia of the quad-rotor unmanned aerial vehicle around the x, y and z axes; tau isx，τyAnd τzRepresenting moments acting on the x, y and z axes, respectively; dx,Dy,DzRepresenting lumped disturbances in the x, y, z axes. (ii) a
In this example Jx＝5.445×10-3,Jy＝5.445×10-3,Jz＝1.089×10-2。
For the convenience of subsequent analysis, the following definitions are introduced:
Figure BDA0003050014040000079
wherein s isφIs sin phi, cφIs cos phi, cθIs cos θ, tθIs a function of the number of tan theta,
the dynamics of the quad-rotor drone attitude loop system can then be rewritten as follows:
Figure BDA0003050014040000081
Figure BDA0003050014040000082
is the first derivative of theta and is,
Figure BDA0003050014040000083
first derivative of Ω
The second-order dynamic state of the attitude angle can be obtained according to the dynamic state 3 of the attitude system
Figure BDA0003050014040000084
Comprises the following steps:
Figure BDA0003050014040000085
w is
Figure BDA0003050014040000086
The first derivative of (a).
The specific operation steps of step S2 include:
and a three-channel decoupling equation of attitude angle tracking error dynamics of the quad-rotor unmanned aerial vehicle is established, and the attitude control problem of the quad-rotor unmanned aerial vehicle is converted into the stabilization problem of the three-channel attitude instruction tracking error. The design method comprises the following specific steps:
defining an attitude tracking error eΘ：
Figure BDA0003050014040000087
Wherein, thetad＝[φd θd ψd]TFor the desired attitude angle, T is transposed, φd、θdAnd psidRespectively, expected instructions of a roll angle, a pitch angle and a yaw angle;
attitude angle tracking error dynamics can be obtained:
Figure BDA0003050014040000088
Figure BDA0003050014040000089
is eΘThe first derivative of (a) is,
Figure BDA00030500140400000810
is eΘThe second derivative of (a) is,
Figure BDA00030500140400000811
is thetadThe first derivative of (a) is,
Figure BDA00030500140400000812
is thetadThe second derivative of (a);
at this time, the attitude angle tracking error system can be dynamically written as:
Figure BDA00030500140400000813
wherein DAThe lumped interference in the attitude tracking error system is expressed as follows:
Figure BDA00030500140400000814
wherein the content of the first and second substances,
Figure BDA0003050014040000091
is thetadThe second derivative of (a).
For attitude angle tracking error dynamic equation 5, let:
Figure BDA0003050014040000092
the tracking error dynamics (5) can be decoupled as:
Figure BDA0003050014040000093
wherein the content of the first and second substances,
Figure BDA0003050014040000094
tracking error e for roll angle commandφSecond derivative of,
Figure BDA0003050014040000095
Error e is tracked for pitch angle commandθSecond derivative of,
Figure BDA0003050014040000096
Tracking error e for yaw angle commandψThe second derivative of (a); e.g. of the typeφ＝φ-φd、eθ＝θ-θdAnd eψ＝ψ-ψd；τφ、τθ、τψRespectively representing roll, pitch and yaw three-channel virtual moments, fA φ、fA θ、fA ψRespectively represents cross-coupling nonlinear terms in a three-channel decoupling dynamic model (formula 6) of rolling, pitching and yawing,
Figure BDA0003050014040000097
lumped disturbances in the roll, pitch and yaw three-channel decoupled dynamics model (equation 6) are represented separately.
The specific operation steps of step S3 include:
designing a rolling, pitching and yawing three-channel extended state observer for a decoupled attitude tracking error system (formula 6, namely a three-channel decoupling equation for enabling attitude angle tracking errors of the quad-rotor unmanned aerial vehicle to be dynamic) so as to realize estimation of three-channel lumped interference, wherein the design method specifically comprises the following steps:
designing a rolling channel extended state observer according to a formula 6, and estimating
Figure BDA0003050014040000098
Figure BDA0003050014040000099
Wherein the content of the first and second substances,
Figure BDA00030500140400000910
in order to roll the channel lumped interference estimates,
Figure BDA00030500140400000911
in order to extend the internal dynamics of the state observer,
Figure BDA00030500140400000912
is composed of
Figure BDA00030500140400000913
The first derivative of (a) is,
Figure BDA00030500140400000914
is composed of
Figure BDA00030500140400000915
The first derivative of (a) is,
Figure BDA00030500140400000916
is composed of
Figure BDA00030500140400000917
The first derivative of (a);
Figure BDA00030500140400000918
and
Figure BDA00030500140400000919
in order to gain the observer,
Figure BDA00030500140400000920
and
Figure BDA00030500140400000921
values are all normal numbers and satisfy characteristic polynomial
Figure BDA00030500140400000922
s is a variable
Establishing a pitching channel extended state observer, estimating
Figure BDA0003050014040000101
Figure BDA0003050014040000102
Wherein
Figure BDA0003050014040000103
To aggregate the interference estimate for the pitch channel,
Figure BDA0003050014040000104
in order to extend the internal dynamics of the state observer,
Figure BDA0003050014040000105
is composed of
Figure BDA0003050014040000106
The first derivative of (a) is,
Figure BDA0003050014040000107
is composed of
Figure BDA0003050014040000108
The first derivative of (a) is,
Figure BDA0003050014040000109
is composed of
Figure BDA00030500140400001010
The first derivative of (a) is,
Figure BDA00030500140400001011
and
Figure BDA00030500140400001012
in order to gain the observer,
Figure BDA00030500140400001013
and
Figure BDA00030500140400001014
values are all normal numbers and satisfy characteristic polynomial
Figure BDA00030500140400001015
Establishing a yaw channel extended state observer, estimating
Figure BDA00030500140400001016
Figure BDA00030500140400001017
Wherein the content of the first and second substances,
Figure BDA00030500140400001018
to aggregate the disturbance estimate for the yaw channel,
Figure BDA00030500140400001019
the internal dynamics of the state observer is expanded,
Figure BDA00030500140400001020
is composed of
Figure BDA00030500140400001021
The first derivative of (a) is,
Figure BDA00030500140400001022
is composed of
Figure BDA00030500140400001023
The first derivative of (a) is,
Figure BDA00030500140400001024
is composed of
Figure BDA00030500140400001025
The first derivative of (a) is,
Figure BDA00030500140400001026
and
Figure BDA00030500140400001027
in order to gain the observer,
Figure BDA00030500140400001028
and
Figure BDA00030500140400001029
are all normal numbers and satisfy characteristic polynomials
Figure BDA00030500140400001030
In this example
Figure BDA00030500140400001031
Figure BDA00030500140400001032
The specific operation steps of step S4 include:
designing a sliding form surface of a rolling channel at a quick terminal as follows:
Figure BDA00030500140400001033
wherein alpha is1＞0,β1＞0,h1,g1,p1,q1Are all positive odd numbers, and
Figure BDA00030500140400001034
Figure BDA00030500140400001035
is eφThe first derivative of (a);
designing a sliding mode surface of a pitching channel quick terminal as follows:
Figure BDA0003050014040000111
wherein alpha is2＞0,β2＞0,h2,g2,p2,q2Are all positive odd numbers, and
Figure BDA0003050014040000112
Figure BDA0003050014040000113
is eθThe first derivative of (a);
designing a yaw channel rapid terminal sliding mode surface as follows:
Figure BDA0003050014040000114
wherein alpha is3＞0,β3＞0,h3,g3,p3,q3Are all positive odd numbers, and
Figure BDA0003050014040000115
Figure BDA0003050014040000116
is eψThe first derivative of (a).
In this example α1＝1,β1＝5,h1＝3,g1＝7,p1＝5,q1＝3；α2＝1,β2＝5,h2＝3,g2＝7,p2＝5,q2＝3；α3＝1,β3＝5,h3＝3,g3＝7,p3＝5,q3＝3。
The specific operation steps of step S5 include:
interference estimation information for equation 6 in combination with a three-channel extended state observer (equations 7, 9, 11)
Figure BDA0003050014040000117
And
Figure BDA0003050014040000118
designing a rapid terminal sliding mode attitude controller, adopting an attractor to replace switching gain for ensuring the continuity of control quantity, and specifically designing the attitude controller by the following steps:
according to equation 6, the roll channel controller is designed as follows:
Figure BDA0003050014040000119
wherein
Figure BDA00030500140400001110
Is a positive real number, m1,n1Is positive odd number, and
Figure BDA00030500140400001111
the pitching channel virtual moment controller comprises:
Figure BDA00030500140400001112
wherein
Figure BDA0003050014040000121
Is a positive real number, m2,n2Is positive odd number, and
Figure BDA0003050014040000122
the yaw channel virtual moment controller is as follows:
Figure BDA0003050014040000123
wherein
Figure BDA0003050014040000124
Is a positive real number, m3,n3Is positive odd number, and
Figure BDA0003050014040000125
and performing inverse solution on the three-channel virtual moment in the S2 to obtain the actual control quantity tau of the four-rotor unmanned aerial vehicle attitude systemx、τyAnd τz：
Figure BDA0003050014040000126
Wherein W-1Is the inverse of the matrix W.
In this example
Figure BDA0003050014040000127
m1＝1,n1＝3；
Figure BDA0003050014040000128
m2＝1,n2＝3；
Figure BDA0003050014040000129
Figure BDA00030500140400001210
m3＝1,n3＝3。
In order to verify the anti-interference performance, the tracking error quick convergence and the buffeting removing effect of the method, the control algorithm provided by the invention is simulated and verified by using an MATLAB simulation environment under the condition that the influence of various external interferences on the quad-rotor unmanned aerial vehicle is considered. In the simulation process, three initial values of attitude angles and three initial values of axial angular speeds are respectively set as:
Figure BDA00030500140400001211
ψ(0)＝0,ωx(0)＝0,ωy(0)＝0,ωz(0)＝0
to make the control task more challenging, the attitude angle command is set to a time-varying form (in radians):
Figure BDA00030500140400001212
wherein t is time
The external interference in the simulation process is set as follows:
Figure BDA0003050014040000131
the invention provides a composite continuous rapid terminal sliding mode attitude control method for a four-rotor unmanned aerial vehicle, which realizes the asymptotic tracking of three-channel attitude angle reference instructions of rolling, pitching and yawing of the four-rotor unmanned aerial vehicle by using continuous control quantity. In order to verify the superior effect of the Composite Continuous Fast Terminal Sliding Mode (CCFNTSM) algorithm, the prior different Sliding Mode control methods are adopted to carry out MATLAB comparison simulation aiming at the four-rotor unmanned aerial vehicle attitude system under the same simulation environment, and 1) the Continuous Fast Terminal Sliding Mode algorithm (FNTSM) is adopted; 2) a Composite Fast probabilistic graphical Sliding Mode (CFNTSM) algorithm; 3) composite Continuous Nonsingular Terminal Sliding Mode (CCNTSM).
Fig. 2, fig. 3 and fig. 4 show three-channel attitude angle tracking error response curves and control quantity curves of the four-rotor unmanned aerial vehicle in rolling, pitching and yawing under different control methods, and the four methods are compared in the following aspects of control quantity continuity, anti-interference performance and tracking error convergence rate of the control system respectively. 1) Control quantity continuity, namely, the control quantity continuity can be ensured by the CCFNTSM of the control scheme, the control quantity continuity can be ensured by both CCNTSM and FNTSM, and the control quantity continuity can not be ensured by CFNTSM, as can be seen by comparing control quantity response curves in figures 2-4; 2) the anti-interference performance is as follows: by comparing the attitude angle tracking error response curves in fig. 2-4 and their enlarged views at 6.4s-7.2s, it can be seen that: the control scheme has the advantages that the CCFNTSM and the CCNTSM have the best anti-interference performance, the FNTSM has the worst anti-interference performance, and the CFNTSM has the anti-interference performance between the two. 3) Tracking error convergence rate by comparing enlarged images of attitude angle tracking error response curves in time periods of 0.4s-0.8s in fig. 2-4, it can be seen that the control method CCFNTSM provided by the invention has a faster convergence rate compared with CCNTSM. The control effects of the above four methods are summarized in table 1:
TABLE 1
Control method FNTSM CFNTSM CCFNTSM CCNTSM
Continuity of control quantity (Continuous) Is discontinuous (Continuous) (Continuous)
Anti-interference performance Difference (D) In general High strength High strength
Convergence rate of tracking error Fastest speed Fastest speed Fast-acting toy Slow
In conclusion, the method provided by the invention can ensure that the disturbed quad-rotor unmanned aerial vehicle attitude system can realize the rapid and high-precision tracking of the attitude instruction under the condition of continuous control quantity.

Claims (6)

1. A four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method is characterized by comprising the following steps: the method specifically comprises the following steps:
s1, establishing a disturbed dynamics model of the attitude loop system of the quad-rotor unmanned aerial vehicle;
s2 moments tau acting on the x, y and z axes of the fuselage coordinate systemx、τyAnd τzPerforming three-channel decoupling to obtain three-channel virtual torque, and establishing a three-channel decoupling equation of attitude angle tracking error dynamics of the quad-rotor unmanned aerial vehicle based on the three-channel virtual torque, wherein the three channels comprise a rolling channel, a pitching channel and a yawing channel;
s3, designing a roll channel, a pitch channel and a yaw channel expansion state observer according to the three-channel decoupling equation in the S2;
s4, designing a four-rotor unmanned aerial vehicle attitude loop fast terminal sliding mode surface;
and S5, performing inverse solution on the three-channel virtual moment in S2 to obtain the actual control quantity of the attitude system of the four-rotor unmanned aerial vehicle, and designing a three-channel composite rapid terminal sliding mode virtual controller to realize the attitude control of the four-rotor unmanned aerial vehicle by combining interference estimation information observed by a three-channel extended state observer in S3 and a rapid terminal sliding mode surface of an attitude loop of the four-rotor unmanned aerial vehicle in S4 according to a three-channel decoupling equation of the attitude angle tracking error dynamics of the unmanned aerial vehicle in S2.
2. The compound continuous fast terminal sliding mode attitude control method of the quad-rotor unmanned aerial vehicle according to claim 1, characterized in that: in S1, the disturbed dynamics model of the attitude loop system of the quad-rotor unmanned aerial vehicle is as follows:
Figure FDA0003050014030000011
wherein the content of the first and second substances,
Figure FDA0003050014030000012
is the first derivative of theta and is,
Figure FDA0003050014030000013
phi denotes the roll angle of the quad-rotor drone, theta denotes the pitch angle of the quad-rotor drone, psi denotes the yaw angle of the quad-rotor drone,
Figure FDA0003050014030000014
sφis sin phi, cφIs cos phi, cθIs cos θ, tθIs a function of the number of tan theta,
Figure FDA0003050014030000015
wherein wxRepresenting angular velocity of rotation, w, of quad-rotor drone about the x-axis of the rectangular coordinate system of the fuselageyRepresenting angular velocity of rotation, w, of quad-rotor drone about the y-axis of the rectangular coordinate system of the fuselagezRepresenting the rotation angular velocity of the quad-rotor unmanned aerial vehicle around the z axis of the rectangular coordinate system of the fuselage;
Figure FDA0003050014030000016
is the first derivative of the omega and,
Figure FDA0003050014030000021
Jx，Jyand JzRespectively represents the rotational inertia of the quadrotor unmanned plane around the x axis, the y axis and the z axis of the rectangular coordinate system of the fuselage,
Figure FDA0003050014030000022
τx、τyand τzRespectively representing the moments of an x axis, a y axis and a z axis of a rectangular coordinate system of the fuselage;
Figure FDA0003050014030000023
Dx，Dyand DzThe lumped disturbances in the x-axis direction, the y-axis direction and the z-axis direction of the rectangular coordinate system of the fuselage are respectively represented.
3. The compound continuous fast terminal sliding mode attitude control method of the quad-rotor unmanned aerial vehicle according to claim 2, characterized in that: the three-channel decoupling equation in the S2 is as follows:
Figure FDA0003050014030000024
Figure FDA0003050014030000025
Figure FDA0003050014030000026
wherein the content of the first and second substances,
Figure FDA0003050014030000027
tracking error e for roll angle commandφSecond derivative of,
Figure FDA0003050014030000028
Error e is tracked for pitch angle commandθSecond derivative of,
Figure FDA0003050014030000029
Tracking error e for yaw angle commandψThe second derivative of (a); e.g. of the typeφ＝φ-φd、eθ＝θ-θd，eψ＝ψ-ψd，φd、θdAnd psidRespectively, expected instructions of a roll angle, a pitch angle and a yaw angle; tau isφ、τθ、τψRepresenting the virtual moments, f, of the roll, pitch and yaw channels, respectivelyA φ、fA θ、fA ψRespectively represents cross-coupling nonlinear terms in the decoupling equations of three channels of rolling, pitching and yawing,
Figure FDA00030500140300000210
representing lumped disturbances, τ, in roll, pitch and yaw triple-channel decoupling equations, respectivelyφ、τθ、τψ，fA φ、fA θ、fA ψAnd
Figure FDA00030500140300000211
the expression of (a) is:
Figure FDA00030500140300000212
wherein
Figure FDA00030500140300000213
Is the first derivative of W and is,
Figure FDA00030500140300000214
is thetadSecond derivative of (theta)d＝[φdθdψd]TFor the desired attitude angle, T represents the matrix transpose, J-1Is the inverse matrix of J.
4. The compound continuous fast terminal sliding mode attitude control method of the quad-rotor unmanned aerial vehicle according to claim 3, characterized in that: the step S3 specifically includes:
establishing a rolling channel extended state observer for estimation
Figure FDA0003050014030000031
Figure FDA0003050014030000032
Wherein the content of the first and second substances,
Figure FDA0003050014030000033
integrating stems for cascading channelsThe value of the disturbance estimate is,
Figure FDA0003050014030000034
and
Figure FDA0003050014030000035
to roll the channel to expand the internal dynamics of the state observer,
Figure FDA0003050014030000036
is composed of
Figure FDA0003050014030000037
The first derivative of (a) is,
Figure FDA0003050014030000038
is composed of
Figure FDA0003050014030000039
The first derivative of (a) is,
Figure FDA00030500140300000310
is composed of
Figure FDA00030500140300000311
The first derivative of (a);
Figure FDA00030500140300000312
and
Figure FDA00030500140300000313
to roll the channel to expand the state observer gain,
Figure FDA00030500140300000314
and
Figure FDA00030500140300000315
values are all normal numbers and satisfy characteristic polynomial
Figure FDA00030500140300000316
s is a variable;
establishing a pitching channel extended state observer, estimating
Figure FDA00030500140300000317
Figure FDA00030500140300000318
Wherein the content of the first and second substances,
Figure FDA00030500140300000319
to aggregate the interference estimate for the pitch channel,
Figure FDA00030500140300000320
the internal dynamics of the state observer are expanded for the pitch channel,
Figure FDA00030500140300000321
is composed of
Figure FDA00030500140300000322
The first derivative of (a) is,
Figure FDA00030500140300000323
is composed of
Figure FDA00030500140300000324
The first derivative of (a) is,
Figure FDA00030500140300000325
is composed of
Figure FDA00030500140300000326
The first derivative of (a) is,
Figure FDA00030500140300000327
and
Figure FDA00030500140300000328
the state observer gain is extended for the pitch channel,
Figure FDA00030500140300000329
and
Figure FDA00030500140300000330
values are all normal numbers and satisfy characteristic polynomial
Figure FDA00030500140300000331
Establishing a yaw channel extended state observer, estimating
Figure FDA00030500140300000332
Figure FDA00030500140300000333
Wherein the content of the first and second substances,
Figure FDA00030500140300000334
to aggregate the disturbance estimate for the yaw channel,
Figure FDA00030500140300000335
the state observer internal dynamics are extended for the yaw channel,
Figure FDA0003050014030000041
is composed of
Figure FDA0003050014030000042
The first derivative of (a) is,
Figure FDA0003050014030000043
is composed of
Figure FDA0003050014030000044
The first derivative of (a) is,
Figure FDA0003050014030000045
is composed of
Figure FDA0003050014030000046
The first derivative of (a) is,
Figure FDA0003050014030000047
and
Figure FDA0003050014030000048
the state observer gain is extended for the yaw path,
Figure FDA0003050014030000049
and
Figure FDA00030500140300000410
are all normal numbers and satisfy characteristic polynomials
Figure FDA00030500140300000411
5. The compound continuous fast terminal sliding mode attitude control method of the quad-rotor unmanned aerial vehicle according to claim 4, characterized in that: four rotor unmanned aerial vehicle gesture return circuits quick terminal slipform face does:
rolling channel rapid terminal sliding mode surface sigmaφComprises the following steps:
Figure FDA00030500140300000412
wherein alpha is1，β1，h1，g1，p1And q is1Are all positive odd numbers, and alpha1＞0,β1＞0,
Figure FDA00030500140300000413
Figure FDA00030500140300000414
Is eφThe first derivative of (a);
pitching channel rapid terminal sliding mode surface sigmaθComprises the following steps:
Figure FDA00030500140300000415
wherein alpha is2，β2，h2,g2,p2And q is2Are all positive odd numbers, and alpha2＞0,β2＞0，
Figure FDA00030500140300000416
Figure FDA00030500140300000417
Is eθThe first derivative of (a);
yaw channel rapid terminal sliding mode surface sigmaψComprises the following steps:
Figure FDA00030500140300000418
wherein alpha is3，β3，h3，g3，p3And q is3Are all positive odd numbers, and alpha3＞0,β3＞0,
Figure FDA00030500140300000419
Figure FDA00030500140300000420
Is eψThe first derivative of (a).
6. The compound continuous fast terminal sliding mode attitude control method of the quad-rotor unmanned aerial vehicle according to claim 5, characterized in that: the three-channel composite rapid terminal sliding mode attitude controller in the S5 comprises a rolling channel virtual moment controller, a pitching channel virtual moment controller and a yawing channel virtual moment controller;
the rolling channel virtual torque controller comprises:
Figure FDA0003050014030000051
wherein
Figure FDA0003050014030000052
Is a positive real number, m1,n1Is positive odd number, and
Figure FDA0003050014030000053
the pitching channel virtual moment controller comprises:
Figure FDA0003050014030000054
wherein
Figure FDA0003050014030000055
Is a positive real number, m2,n2Is positive odd number, and
Figure FDA0003050014030000056
the yaw channel virtual moment controller is as follows:
Figure FDA0003050014030000057
wherein
Figure FDA0003050014030000058
Is a positive real number, m3,n3Is positiveIs odd, and
Figure FDA0003050014030000059
CN202110483380.9A 2021-04-30 2021-04-30 Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method Active CN113110554B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202110483380.9A CN113110554B (en) 2021-04-30 2021-04-30 Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202110483380.9A CN113110554B (en) 2021-04-30 2021-04-30 Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method

Publications (2)

Publication Number Publication Date
CN113110554A true CN113110554A (en) 2021-07-13
CN113110554B CN113110554B (en) 2022-03-11

Family

ID=76720786

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202110483380.9A Active CN113110554B (en) 2021-04-30 2021-04-30 Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method

Country Status (1)

Country Link
CN (1) CN113110554B (en)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114489105A (en) * 2022-01-25 2022-05-13 南京邮电大学 Novel unmanned aerial vehicle attitude system integral sliding mode control method based on disturbance observer
CN115390447A (en) * 2022-08-16 2022-11-25 哈尔滨逐宇航天科技有限责任公司 Aircraft preset performance control method suitable for large-amplitude attitude maneuver
CN116846271A (en) * 2023-07-22 2023-10-03 湖南工业大学 Permanent magnet synchronous motor model-free continuous and rapid terminal sliding mode fault-tolerant control method and system

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20130092785A1 (en) * 2008-07-11 2013-04-18 Davidson Technologies, Inc. System and method for guiding and controlling a missile using high order sliding mode control
CN106444799A (en) * 2016-07-15 2017-02-22 浙江工业大学 Quadrotor unmanned plane control method based on fuzzy expansion state observer and adaptive sliding formwork
CN106802660A (en) * 2017-03-09 2017-06-06 北京航天自动控制研究所 One kind is combined strong anti-interference attitude control method
CN108646556A (en) * 2018-05-08 2018-10-12 中国人民解放军战略支援部队航天工程大学 Input saturation spacecraft is without unwinding Attitude tracking control method
CN110221539A (en) * 2019-05-17 2019-09-10 江苏理工学院 Quadrotor non-singular terminal sliding-mode control based on linear expansion observer
CN110320925A (en) * 2019-04-18 2019-10-11 南京航空航天大学 Quadrotor method of controlling security based on high-order interference observer
CN110850887A (en) * 2019-10-10 2020-02-28 南京航空航天大学 Composite dynamic inverse anti-interference attitude control method for quad-rotor unmanned aerial vehicle
CN111650951A (en) * 2020-05-22 2020-09-11 南京航空航天大学 Full-loop composite dynamic inverse tracking control method for complex track of quad-rotor unmanned aerial vehicle
CN111722634A (en) * 2020-05-28 2020-09-29 南京邮电大学 Four-rotor aircraft sliding mode control method based on nonlinear disturbance observer

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20130092785A1 (en) * 2008-07-11 2013-04-18 Davidson Technologies, Inc. System and method for guiding and controlling a missile using high order sliding mode control
CN106444799A (en) * 2016-07-15 2017-02-22 浙江工业大学 Quadrotor unmanned plane control method based on fuzzy expansion state observer and adaptive sliding formwork
CN106802660A (en) * 2017-03-09 2017-06-06 北京航天自动控制研究所 One kind is combined strong anti-interference attitude control method
CN108646556A (en) * 2018-05-08 2018-10-12 中国人民解放军战略支援部队航天工程大学 Input saturation spacecraft is without unwinding Attitude tracking control method
CN110320925A (en) * 2019-04-18 2019-10-11 南京航空航天大学 Quadrotor method of controlling security based on high-order interference observer
CN110221539A (en) * 2019-05-17 2019-09-10 江苏理工学院 Quadrotor non-singular terminal sliding-mode control based on linear expansion observer
CN110850887A (en) * 2019-10-10 2020-02-28 南京航空航天大学 Composite dynamic inverse anti-interference attitude control method for quad-rotor unmanned aerial vehicle
CN111650951A (en) * 2020-05-22 2020-09-11 南京航空航天大学 Full-loop composite dynamic inverse tracking control method for complex track of quad-rotor unmanned aerial vehicle
CN111722634A (en) * 2020-05-28 2020-09-29 南京邮电大学 Four-rotor aircraft sliding mode control method based on nonlinear disturbance observer

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
王剑颖等: "高超声速飞行器连续终端滑模姿态控制方法", 《哈尔滨工程大学学报》 *

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114489105A (en) * 2022-01-25 2022-05-13 南京邮电大学 Novel unmanned aerial vehicle attitude system integral sliding mode control method based on disturbance observer
CN114489105B (en) * 2022-01-25 2024-01-16 南京邮电大学 Novel unmanned aerial vehicle attitude system integral sliding mode control method based on interference observer
CN115390447A (en) * 2022-08-16 2022-11-25 哈尔滨逐宇航天科技有限责任公司 Aircraft preset performance control method suitable for large-amplitude attitude maneuver
CN115390447B (en) * 2022-08-16 2024-05-14 哈尔滨逐宇航天科技有限责任公司 Aircraft preset performance control method adapting to great maneuvering of gestures
CN116846271A (en) * 2023-07-22 2023-10-03 湖南工业大学 Permanent magnet synchronous motor model-free continuous and rapid terminal sliding mode fault-tolerant control method and system
CN116846271B (en) * 2023-07-22 2024-04-19 湖南工业大学 Permanent magnet synchronous motor model-free rapid terminal sliding mode fault-tolerant control method and system

Also Published As

Publication number Publication date
CN113110554B (en) 2022-03-11

Similar Documents

Publication Publication Date Title
CN113110554B (en) Four-rotor unmanned aerial vehicle composite continuous rapid terminal sliding mode attitude control method
CN110850887B (en) Composite dynamic inverse anti-interference attitude control method for quad-rotor unmanned aerial vehicle
CN110262557B (en) Four-rotor formation finite time fault-tolerant control method based on sliding mode control technology
CN108681240B (en) Fault diagnosis method for unmanned aerial vehicle distributed formation based on unknown input observer
CN114578691A (en) Active anti-interference fault-tolerant attitude control method of flying wing unmanned aerial vehicle considering control plane fault
CN107976903B (en) Enhanced double-power-order approach law sliding mode control method of quad-rotor unmanned aerial vehicle system
CN111948944A (en) Four-rotor formation fault-tolerant control method based on adaptive neural network
CN107688295A (en) A kind of quadrotor finite time self-adaptation control method based on fast terminal sliding formwork
CN107957682B (en) Enhanced fast power-order approach law sliding mode control method of quad-rotor unmanned aerial vehicle system
Cheng et al. Robust finite-time cooperative formation control of UGV-UAV with model uncertainties and actuator faults
Wu et al. Quantized fixed-time fault-tolerant attitude control for hypersonic reentry vehicles
CN111026160B (en) Trajectory tracking control method for quad-rotor unmanned aerial vehicle
Wang et al. Sliding mode decoupling control of a generic hypersonic vehicle based on parametric commands
CN108803648A (en) Unmanned vehicle robust attitude control method, device and electronic equipment
CN109283932A (en) A kind of quadrotor attitude control method based on integral contragradience sliding formwork
CN115639830B (en) Air-ground intelligent agent cooperative formation control system and formation control method thereof
CN109460055B (en) Aircraft control capability determining method and device and electronic equipment
Tian et al. Nonsingular fixed-time attitude coordinated tracking control for multiple rigid spacecraft
CN111650951B (en) Full-loop composite dynamic inverse tracking control method for complex track of quad-rotor unmanned aerial vehicle
CN105115503B (en) A kind of servo loop decoupling method of three-axle steady platform system
Liu et al. Integrated guidance and control with partial state constraints and actuator faults
CN107942672B (en) Four-rotor aircraft output limited backstepping control method based on symmetric time invariant obstacle Lyapunov function
CN112327629A (en) Small unmanned helicopter self-adaptive fault-tolerant control method based on dynamic compensation
CN108107726B (en) Four-rotor aircraft output limited backstepping control method based on symmetric time-varying obstacle Lyapunov function
CN116430828A (en) Four-rotor fault disaster recovery degradation control method based on observer

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant