CN109212969A - A kind of integral contragradience Sliding Mode Attitude control method considering quadrotor actuator failures - Google Patents

A kind of integral contragradience Sliding Mode Attitude control method considering quadrotor actuator failures Download PDF

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CN109212969A
CN109212969A CN201811078978.4A CN201811078978A CN109212969A CN 109212969 A CN109212969 A CN 109212969A CN 201811078978 A CN201811078978 A CN 201811078978A CN 109212969 A CN109212969 A CN 109212969A
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quadrotor
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attitude
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陈强
朱健宏
陶玫玲
胡轶
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Zhejiang University of Technology ZJUT
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    • G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
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Abstract

A kind of integral contragradience Sliding Mode Attitude control method considering quadrotor actuator failures, for the quadrotor attitude control system of actuator failures, utilize Backstepping design method, in conjunction with sliding formwork control, system is set to have very strong robustness to interference and unmodeled item, integral term is added in the design process of controller, advantageously reduce buffeting and guarantees the rapidity of system, designs a kind of integral contragradience Sliding Mode Attitude control method for considering quadrotor actuator failures.This method can effectively reduce the influence of quadrotor actuator failures, and can track the default desired value of attitude angle, realize the stability contorting to quadrotor posture.

Description

A kind of integral contragradience Sliding Mode Attitude control considering quadrotor actuator failures Method
Technical field
The present invention relates to a kind of integral contragradience Sliding Mode Attitude control methods for considering quadrotor actuator failures, make Quadrotor posture can track the default desired value of upper attitude angle in actuator failures.Scape technology
Quadrotor is because of its compact and flexible structure, VTOL, the characteristics of freely hovering, cheap cost, with And brilliant performance is widely used in every field, such as photography of taking photo by plane, disaster inspection.In addition to outside air power, parameter are taken the photograph Dynamic, unmodeled item of system etc. influences, and in the actual flight course of quadrotor, stablizes since quadrotor controls The probability that the influence of property and itself technique, motor and the lasting high speed rotation of propeller make it break down greatly improves, and one Denier breaks down in flight course, quadrotor drive lacking characteristic in nature and its high non-linearity and close coupling Property can be such that flight stability drastically reduces, and result even in out of control, cause serious consequence.Quadrotor most common failure includes Sensor fault and actuator failures, wherein actuator failures occurrence frequency is higher, and the influence to quadrotor is bigger, It is more difficult to solve.Therefore, it is necessary to design a kind of suitable quadrotor control strategy come the case where coping with actuator failures.Mesh It is preceding for gesture stability problem of the quadrotor in actuator failures there are many plant control method, as PID control, Sliding formwork control, self adaptive control and Reverse Step Control etc..
Reverse Step Control is suitable for can state linearisation or the uncertain nonlinear system with tight parameter feedback.Therefore for Drive lacking, non-linear and close coupling quadrotor attitude control system are conducive to using Backstepping design controller The quick response of aircraft and enhancing are not known to system and the robustness of external disturbance.And sliding formwork control and control target ginseng Uncertainty that is unrelated, therefore can effectively overcoming system is counted and disturbed, also has very strong robustness to interference and unmodeled item. Sliding formwork control there are also fast response time, without system on-line identification and physics realization it is simple the advantages that, to nonlinear Control system System has good control effect.Integral term is added in the design process of controller, is conducive to reduce and buffet, and further add The response time of fast quadrotor posture control.The combination of Reverse Step Control and sliding formwork control is conducive to quadrotor and flies Row device overcomes the uncertainty of system while quick response, enhances the robustness to unmodeled item, external disturbance etc..
Summary of the invention
In the case where overcoming existing quadrotor actuator to break down, quadrotor cannot keep peace The deficiency of stable flight entirely, the present invention provide a kind of integral contragradience Sliding Mode Attitude for considering quadrotor actuator failures Control method can effectively reduce the influence of quadrotor actuator failures, and track the default desired value of upper attitude angle, real Now to the stability contorting of quadrotor posture.
In order to solve the above-mentioned technical problem the technical solution proposed is as follows:
A kind of integral contragradience Sliding Mode Attitude control method considering quadrotor actuator failures, including following step It is rapid:
Step 1, establish the attitude dynamics model of quadrotor system, initialization system mode, the sampling time with And control parameter, process are as follows:
1.1 in the case where considering quadrotor actuator failures, is based on Euler's formula, from loading angle analysis four The dynamic model expression formula of rotor craft, attitude kinematics expression formula and its rotation process are as follows:
Wherein, η indicates attitude angle matrix,Indicating the first derivative of η, J indicates quadrotor moment of inertia matrix, Ω indicates quadrotor attitude angular velocity matrix,Indicate the first derivative of Ω, the multiplication cross of × representing matrix, D indicates four Rotor craft actuator efficiency matrix, u indicate quadrotor controller input matrix；
Formula (1) is substituted into formula (2):
Wherein,Indicate the second dervative of η；
Expansion (3):
Wherein, Jx、Jy、JzRespectively under body coordinate system, the component of x, y, z axis rotary inertia, φ, θ, ψ, which are respectively indicated, to be turned over Roll angle, pitch angle and yaw angle,The first derivative of roll angle, pitch angle and yaw angle is respectively indicated,Respectively indicate the second dervative of roll angle, pitch angle and yaw angle, ux、uy、uzRespectively indicate quadrotor The controller of x, y, z axis inputs, δ1、δ2、δ3The working efficiency for respectively indicating quadrotor x, y, z axis actuator, works as δi= When 1, actuator fault-free, as 0 < δiWhen < 1, actuator section failure, but remain to work on, wherein i=1,2,3；
Step 2, the error of quadrotor attitude angle is calculated, design sliding-mode surface restrains attitude angle, and process is as follows:
2.1 define posture angle tracking error e:
E=ηd-η (5)
Wherein, ηd=[φd,θd,ψd]TIndicate expectation attitude angle, φd、θd、ψdRespectively indicate desired roll angle, expectation pitching Angle and desired yaw angle；
Define sliding-mode surface s:
S=e+KI∫edt (6)
Wherein, KI∈R3×3For the diagonal integral coefficient matrix of positive definite；
The first derivative of formula (6) are as follows:
2.2 design liapunov function V1Are as follows:
The first derivative of formula (8) are as follows:
Wherein,Indicate ηdFirst derivative, virtual controlling restrain α1Expression formula are as follows:
Wherein, k11∈R3×3For positive definite diagonal matrix；
The first derivative of formula (10) are as follows:
Enable the expression formula of h are as follows:
The first derivative of formula (12) are as follows:
Formula (12) are substituted into formula (9):
Formula (10) are substituted into formula (12):
Formula (15) are transplanted:
Formula (10) are substituted into formula (14):
Step 3, it is based on quadrotor attitude dynamics model, according to the sliding-mode surface designed, is considering quadrotor In the case where aircraft actuator failures, design integrates contragradience Sliding Mode Attitude controller, and process is as follows:
3.1 consider formula (3), and integral contragradience Sliding Mode Attitude controller u is designed to:
Wherein,Indicate ηdSecond dervative, k22∈R3×3For positive definite diagonal matrix, ε is a positive real number；
Saturation function sat (h) is introduced to reduce to buffet, is defined as:
Wherein, | | h | | indicate the norm of h, Δ is a positive real number；
It is obtained by formula (19):
Then formula (20) is positive definite；
3.2 design liapunov function V2Are as follows:
The first derivative of formula (21) are as follows:
Formula (17) are substituted into formula (22):
Formula (13) are substituted into formula (23):
Formula (11) are substituted into formula (24):
Formula (16) are substituted into formula (25):
Formula (3) are substituted into formula (26):
Formula (18), formula (19) are substituted into formula (27):
Therefore, s, h can converge to zero；
Since s can converge to zero, consider formula (6), it is known that posture angle tracking error e can converge to zero, show system It is stable.
The present invention is based on it is a kind of consider quadrotor actuator failures integral contragradience Sliding Mode Attitude control method, In the case that quadrotor actuator breaks down, influence of the failure to quadrotor can be effectively reduced, and can The default desired value of attitude angle in tracking.
Technical concept of the invention are as follows: for the quadrotor attitude control system of actuator failures, utilize contragradience Design method makes system have very strong robustness to interference and unmodeled item, in the design process of controller in conjunction with sliding formwork control Middle addition integral term advantageously reduces buffeting and guarantees the rapidity of system, designs a kind of consideration quadrotor execution The integral contragradience Sliding Mode Attitude control method of device failure.This method can effectively reduce the shadow of quadrotor actuator failures It rings, and the default desired value of attitude angle can be tracked, realize the stability contorting to quadrotor posture.
The influence generated when the invention has the benefit that reducing quadrotor actuator failures, and can track The default desired value of upper attitude angle.
Detailed description of the invention
Attitude angle tracking effect schematic diagram when Fig. 1 is quadrotor actuator fault-free of the present invention.
Sliding-mode surface s schematic diagram when Fig. 2 is quadrotor actuator fault-free of the present invention.
Controller when Fig. 3 is quadrotor actuator fault-free of the present invention inputs u schematic diagram.
Fig. 4 is that attitude angle tracking effect of the quadrotor actuator of the present invention when partial fault occurs for 0.25s shows It is intended to.
Fig. 5 is sliding-mode surface s schematic diagram of the quadrotor actuator of the present invention when partial fault occurs for 0.25s.
Fig. 6 is that controller of the quadrotor actuator of the present invention when partial fault occurs for 0.25s inputs u signal Figure.
Fig. 7 is control flow schematic diagram of the invention.
Specific embodiment
The present invention will be further described with reference to the accompanying drawing.
Referring to Fig.1~Fig. 7, a kind of integral contragradience Sliding Mode Attitude controlling party considering quadrotor actuator failures Method, the control method the following steps are included:
Step 1, establish the attitude dynamics model of quadrotor system, initialization system mode, the sampling time with And control parameter, process are as follows:
1.1 in the case where considering quadrotor actuator failures, is based on Euler's formula, from loading angle analysis four The dynamic model expression formula of rotor craft, attitude kinematics expression formula and its rotation process are as follows:
Wherein, η indicates attitude angle matrix,Indicating the first derivative of η, J indicates quadrotor moment of inertia matrix, Ω indicates quadrotor attitude angular velocity matrix,Indicate the first derivative of Ω, the multiplication cross of × representing matrix, D indicates four Rotor craft actuator efficiency matrix, u indicate quadrotor controller input matrix；
Formula (1) is substituted into formula (2):
Wherein,Indicate the second dervative of η；
Expansion (3):
Wherein, Jx、Jy、JzRespectively under body coordinate system, the component of x, y, z axis rotary inertia, φ, θ, ψ, which are respectively indicated, to be turned over Roll angle, pitch angle and yaw angle,The first derivative of roll angle, pitch angle and yaw angle is respectively indicated,Respectively indicate the second dervative of roll angle, pitch angle and yaw angle, ux、uy、uzRespectively indicate quadrotor The controller of x, y, z axis inputs, δ1、δ2、δ3The working efficiency for respectively indicating quadrotor x, y, z axis actuator, works as δi= When 1, actuator fault-free, as 0 < δiWhen < 1, actuator section failure, but remain to work on, wherein i=1,2,3；
Step 2, the error of quadrotor attitude angle is calculated, design sliding-mode surface restrains attitude angle, and process is as follows:
2.1 define posture angle tracking error e:
E=ηd-η (5)
Wherein, ηd=[φd,θd,ψd]TIndicate expectation attitude angle, φd、θd、ψdRespectively indicate desired roll angle, expectation pitching Angle and desired yaw angle；
Define sliding-mode surface s:
S=e+KI∫edt (6)
Wherein, KI∈R3×3For the diagonal integral coefficient matrix of positive definite；
The first derivative of formula (6) are as follows:
2.2 design liapunov function V1Are as follows:
The first derivative of formula (8) are as follows:
Wherein,Indicate ηdFirst derivative, virtual controlling restrain α1Expression formula are as follows:
Wherein, k11∈R3×3For positive definite diagonal matrix；
The first derivative of formula (10) are as follows:
Enable the expression formula of h are as follows:
The first derivative of formula (12) are as follows:
Formula (12) are substituted into formula (9):
Formula (10) are substituted into formula (12):
Formula (15) are transplanted:
Formula (10) are substituted into formula (14):
Step 3, it is based on quadrotor attitude dynamics model, according to the sliding-mode surface designed, is considering quadrotor In the case where aircraft actuator failures, design integrates contragradience Sliding Mode Attitude controller, and process is as follows:
3.1 consider formula (3), and integral contragradience Sliding Mode Attitude controller u is designed to:
Wherein,Indicate ηdSecond dervative, k22∈R3×3For positive definite diagonal matrix, ε is a positive real number；
Saturation function sat (h) is introduced to reduce to buffet, is defined as:
Wherein, | | h | | indicate the norm of h, Δ is a positive real number；
It is obtained by formula (19):
Then formula (20) is positive definite；
3.2 design liapunov function V2Are as follows:
The first derivative of formula (21) are as follows:
Formula (17) are substituted into formula (22):
Formula (13) are substituted into formula (23):
Formula (11) are substituted into formula (24):
Formula (16) are substituted into formula (25):
Formula (3) are substituted into formula (26):
Formula (18), formula (19) are substituted into formula (27):
Therefore, s, h can converge to zero；
Since s can converge to zero, consider formula (6), it is known that posture angle tracking error e can converge to zero, show system It is stable.
For the validity of the mentioned method of the verifying present invention, The present invention gives quadrotors in actuator fault-free feelings The Contrast on effect of condition and actuator partial fault situation:
More effectively to compare, all parameters of system are all consistent, it may be assumed that quadrotor system is in zero moment For stationary state: η (0)=[0,0,0]TRad, Ω (0)=[0,0,0]TRad, attitude angle desired value are as follows: ηd=[5,5,5]TRad, quadrotor moment of inertia matrix are as follows: J=diag (0.00235,0.00235,0.0526) kgm2, quadrotor fly Row device actuator efficiency are as follows:Other control parameters of system are as follows: KI= Diag (3,3,3), k11=diag (1.5,1.5,1.5), k22=diag (5,5,5), ε=10, Δ=1.
Posture angle tracking schematic diagram, sliding-mode surface s schematic diagram when Fig. 1-Fig. 3 respectively indicates quadrotor normal flight U schematic diagram is inputted with controller, in the case where quadrotor actuator does not break down, the convergence time of attitude angle is 1.2s, sliding-mode surface s convergence time are 0.4s, and controller inputs u and reaches stable state in 0.75s.Fig. 4-6 respectively indicates quadrotor flight Posture angle tracking schematic diagram, sliding-mode surface s schematic diagram and controller when device actuator failures input u schematic diagram, quadrotor flight The actuator of device x, y, z axis section failure in 0.25s loses 60%, 50%, 90% performance respectively.It is considerable by Fig. 6 The actuator for observing x, y, z axis generates spike in 0.25s, but remains to restrain after the adjusting of controller.And attitude angle Change unobvious, respectively 1.2s and 0.4s with the convergence time of sliding-mode surface s, therefore this control method can reduce actuator event While barrier influences, the good dynamic property of quadrotor attitude control system is kept.
In conclusion considering that the integral contragradience sliding-mode control of quadrotor actuator failures can be in quadrotor In the case that aircraft actuator breaks down, influence of the failure to quadrotor is effectively reduced, and appearance can be tracked The default desired value at state angle realizes the stability contorting to quadrotor posture.
Described above is the excellent effect of optimization that one embodiment that the present invention provides is shown, it is clear that the present invention is not only It is limited to above-described embodiment, without departing from essence spirit of the present invention and without departing from the premise of range involved by substantive content of the present invention Under it can be made it is various deformation be implemented.

Claims (1)

1. a kind of integral contragradience Sliding Mode Attitude control method for considering quadrotor actuator failures, it is characterised in that: institute State control method the following steps are included:
Step 1, the attitude dynamics model of quadrotor system, initialization system mode, sampling time and control are established Parameter processed, process are as follows:
1.1 in the case where considering quadrotor actuator failures, is based on Euler's formula, analyzes quadrotor from loading angle The dynamic model expression formula of aircraft, attitude kinematics expression formula and its rotation process are as follows:
Wherein, η indicates attitude angle matrix,Indicate the first derivative of η, J indicates quadrotor moment of inertia matrix, Ω table Show quadrotor attitude angular velocity matrix,Indicate the first derivative of Ω, the multiplication cross of × representing matrix, D indicates quadrotor Aircraft actuator efficiency matrix, u indicate quadrotor controller input matrix；Formula (1) is substituted into formula (2):
Wherein,Indicate the second dervative of η；
Expansion (3):
Wherein, Jx、Jy、JzRespectively under body coordinate system, the component of x, y, z axis rotary inertia, φ, θ, ψ respectively indicate rolling Angle, pitch angle and yaw angle,The first derivative of roll angle, pitch angle and yaw angle is respectively indicated, Respectively indicate the second dervative of roll angle, pitch angle and yaw angle, ux、uy、uzRespectively indicate quadrotor x, y, z axis Controller input, δ1、δ2、δ3The working efficiency for respectively indicating quadrotor x, y, z axis actuator, works as δiWhen=1, hold Row device fault-free, as 0 < δiWhen < 1, actuator section failure, but remain to work on, wherein i=1,2,3；
Step 2, the error of quadrotor attitude angle is calculated, design sliding-mode surface restrains attitude angle, and process is as follows:
2.1 define posture angle tracking error e:
E=ηd-η (5)
Wherein, ηd=[φd,θd,ψd]TIndicate expectation attitude angle, φd、θd、ψdRespectively indicate desired roll angle, expectation pitch angle and It is expected that yaw angle；
Define sliding-mode surface s:
S=e+KI∫edt (6)
Wherein, KI∈R3×3For the diagonal integral coefficient matrix of positive definite；
The first derivative of formula (6) are as follows:
2.2 design liapunov function V1Are as follows:
The first derivative of formula (8) are as follows:
Wherein,Indicate ηdFirst derivative, virtual controlling restrain α1Expression formula are as follows:
Wherein, k11∈R3×3For positive definite diagonal matrix；
The first derivative of formula (10) are as follows:
Enable the expression formula of h are as follows:
The first derivative of formula (12) are as follows:
Formula (12) are substituted into formula (9):
Formula (10) are substituted into formula (12):
Formula (15) are transplanted:
Formula (10) are substituted into formula (14):
Step 3, it is based on quadrotor attitude dynamics model, according to the sliding-mode surface designed, is considering quadrotor flight In the case where device actuator failures, design integrates contragradience Sliding Mode Attitude controller, and process is as follows:
3.1 consider formula (3), and integral contragradience Sliding Mode Attitude controller u is designed to:
Wherein,Indicate ηdSecond dervative, k22∈R3×3For positive definite diagonal matrix, ε is a positive real number；
Saturation function sat (h) is introduced to reduce to buffet, is defined as:
Wherein, | | h | | indicate the norm of h, Δ is a positive real number；
It is obtained by formula (19):
Then formula (20) is positive definite；
3.2 design liapunov function V2Are as follows:
The first derivative of formula (21) are as follows:
Formula (17) are substituted into formula (22):
Formula (13) are substituted into formula (23):
Formula (11) are substituted into formula (24):
Formula (16) are substituted into formula (25):
Formula (3) are substituted into formula (26):
Formula (18), formula (19) are substituted into formula (27):
Therefore, s, h can converge to zero；
Since s can converge to zero, consider formula (6), it is known that posture angle tracking error e can converge to zero, show that system is steady Fixed.
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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112947513A (en) * 2021-01-27 2021-06-11 西北工业大学 Four-rotor unmanned aerial vehicle attitude control method based on fault-tolerant and anti-saturation mechanism

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104102128A (en) * 2013-04-09 2014-10-15 中国人民解放军第二炮兵工程大学 Anti-interference attitude control method suitable for miniaturized unmanned aircraft
US20160209850A1 (en) * 2014-12-09 2016-07-21 Embry-Riddle Aeronautical University, Inc. System and method for robust nonlinear regulation control of unmanned aerial vehicles syntetic jet actuators
CN107688295A (en) * 2017-08-29 2018-02-13 浙江工业大学 A kind of quadrotor finite time self-adaptation control method based on fast terminal sliding formwork
CN107703952A (en) * 2017-08-29 2018-02-16 浙江工业大学 A kind of nonsingular set time Adaptive Attitude control method of rigid aircraft
CN108037662A (en) * 2017-12-06 2018-05-15 浙江工业大学 A kind of limited backstepping control method of quadrotor output based on Integral Sliding Mode obstacle liapunov function

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104102128A (en) * 2013-04-09 2014-10-15 中国人民解放军第二炮兵工程大学 Anti-interference attitude control method suitable for miniaturized unmanned aircraft
US20160209850A1 (en) * 2014-12-09 2016-07-21 Embry-Riddle Aeronautical University, Inc. System and method for robust nonlinear regulation control of unmanned aerial vehicles syntetic jet actuators
CN107688295A (en) * 2017-08-29 2018-02-13 浙江工业大学 A kind of quadrotor finite time self-adaptation control method based on fast terminal sliding formwork
CN107703952A (en) * 2017-08-29 2018-02-16 浙江工业大学 A kind of nonsingular set time Adaptive Attitude control method of rigid aircraft
CN108037662A (en) * 2017-12-06 2018-05-15 浙江工业大学 A kind of limited backstepping control method of quadrotor output based on Integral Sliding Mode obstacle liapunov function

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
XIAOGANG JIN 等: "Intelligent fuzzy controller of a quadrotor", 《PROCEEDINGS 2010 IEEE INTERNATIONAL CONFERENCE ON INTELLIGENT SYSTEMS AND KNOWLEDGE ENGINEERING (ISKE 2010) 》 *
陈强 等: "基于全阶滑模的四旋翼无人机有限时间控制", 《探索 创新 交流（第7集）-第七届中国航空学会青年科技论坛文集（下册）》 *

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112947513A (en) * 2021-01-27 2021-06-11 西北工业大学 Four-rotor unmanned aerial vehicle attitude control method based on fault-tolerant and anti-saturation mechanism
CN112947513B (en) * 2021-01-27 2022-10-21 西北工业大学 Four-rotor unmanned aerial vehicle attitude control method based on fault-tolerant and anti-saturation mechanism

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