CN109917651A - A kind of Spacecraft Attitude Control that symmetrical time-varying output is limited - Google Patents
A kind of Spacecraft Attitude Control that symmetrical time-varying output is limited Download PDFInfo
- Publication number
- CN109917651A CN109917651A CN201910191632.3A CN201910191632A CN109917651A CN 109917651 A CN109917651 A CN 109917651A CN 201910191632 A CN201910191632 A CN 201910191632A CN 109917651 A CN109917651 A CN 109917651A
- Authority
- CN
- China
- Prior art keywords
- formula
- indicate
- quadrotor
- follows
- derivative
- 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.)
- Withdrawn
Links
- 238000013461 design Methods 0.000 claims abstract description 34
- 238000000034 method Methods 0.000 claims abstract description 23
- 230000001133 acceleration Effects 0.000 claims description 10
- 238000005096 rolling process Methods 0.000 claims description 6
- 238000005070 sampling Methods 0.000 claims description 6
- NAWXUBYGYWOOIX-SFHVURJKSA-N (2s)-2-[[4-[2-(2,4-diaminoquinazolin-6-yl)ethyl]benzoyl]amino]-4-methylidenepentanedioic acid Chemical compound C1=CC2=NC(N)=NC(N)=C2C=C1CCC1=CC=C(C(=O)N[C@@H](CC(=C)C(O)=O)C(O)=O)C=C1 NAWXUBYGYWOOIX-SFHVURJKSA-N 0.000 claims description 4
- 230000005484 gravity Effects 0.000 claims description 3
- 238000012546 transfer Methods 0.000 claims description 3
- 238000013519 translation Methods 0.000 claims description 3
- 241000208340 Araliaceae Species 0.000 claims 1
- 235000005035 Panax pseudoginseng ssp. pseudoginseng Nutrition 0.000 claims 1
- 235000003140 Panax quinquefolius Nutrition 0.000 claims 1
- 235000008434 ginseng Nutrition 0.000 claims 1
- 239000011159 matrix material Substances 0.000 claims 1
- 150000001875 compounds Chemical class 0.000 abstract description 4
- 238000010586 diagram Methods 0.000 description 5
- 230000000694 effects Effects 0.000 description 3
- 238000013507 mapping Methods 0.000 description 3
- 230000001052 transient effect Effects 0.000 description 3
- 235000013399 edible fruits Nutrition 0.000 description 1
- 238000009415 formwork Methods 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 238000011160 research Methods 0.000 description 1
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
- G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
- G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
- G05B13/04—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
- G05B13/042—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators in which a parameter or coefficient is automatically adjusted to optimise the performance
-
- 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
- G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
- G05D1/0825—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using mathematical models
-
- 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)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Evolutionary Computation (AREA)
- Artificial Intelligence (AREA)
- Computer Vision & Pattern Recognition (AREA)
- Health & Medical Sciences (AREA)
- Medical Informatics (AREA)
- Software Systems (AREA)
- Algebra (AREA)
- Mathematical Analysis (AREA)
- Mathematical Optimization (AREA)
- Mathematical Physics (AREA)
- Pure & Applied Mathematics (AREA)
- Feedback Control In General (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
A kind of Spacecraft Attitude Control that symmetrical time-varying output is limited, for the dynamic system of quadrotor, a kind of compound constraint liapunov function of symmetrical time-varying logarithm tangent is selected, a kind of quadrotor output constrained control method based on the compound constraint liapunov function of symmetrical time-varying logarithm tangent is designed.The design of the symmetrical compound constraint liapunov function of time-varying logarithm tangent is while can also to reduce arrival time to guarantee that the output of system can limit and avoid excessive overshoot in a certain range.So as to improve the dynamic response performance of quadrotor system.The present invention provides a kind of Spacecraft Attitude Control that symmetrical time-varying output is limited, and system is made to have preferable dynamic response process.
Description
Technical field
The present invention relates to a kind of Spacecraft Attitude Controls that symmetrical time-varying output is limited, make quadrotor system
There is preferable dynamic response process.
Background technique
The one kind of quadrotor as rotary aircraft, it is small in size with its, mobility is good, design is simple, system
The advantages that low in cost is made, the extensive concern of domestic and international university, research institution, company has been attracted.However, since quadrotor is flown
Device is small in size and light-weight, in-flight vulnerable to external disturbance, how to realize the High Performance Motion Control to quadrotor
Have become a hot issue.For the control problem of quadrotor, there are many control methods, such as PID control,
Active Disturbance Rejection Control, sliding formwork control, Reverse Step Control etc..
Wherein Reverse Step Control has been widely used for nonlinear system, and advantage includes fast response time, easy to implement, right
The uncertain robustness etc. with external disturbance of system.Traditional Reverse Step Control only considers the stability of quadrotor
Can, there is no pay close attention to its transient response performance too much.Therefore, traditional backstepping control method makes quadrotor system
Application in a practical situation has very big obstruction.To solve this problem, the Reverse Step Control based on constraint liapunov function
Method is suggested, and this method can effectively improve the mapping of quadrotor system in a practical situation.
Summary of the invention
In order to improve quadrotor system transients performance, the present invention provides what a kind of symmetrical time-varying output was limited to fly
Row device attitude control method, reduces overshoot and overshoot time, and quadrotor system is made to have a good dynamic
Response performance.
In order to solve the above-mentioned technical problem the technical solution proposed is as follows:
A kind of Spacecraft Attitude Control that symmetrical time-varying output is limited, comprising the following steps:
Step 1, the dynamic model for establishing quadrotor system sets initial value, sampling time and the control of system
Parameter processed, process are as follows:
1.1 determine from the body coordinate system based on quadrotor system to the transfer square of the inertial coordinate based on the earth
Battle array T:
Wherein, φ, θ, ψ are roll angle, pitch angle, the yaw angle of quadrotor respectively, indicate aircraft successively around
The angle of each reference axis rotation of inertial coodinate system；
Dynamic model during the translation of 1.2 quadrotors is as follows:
Wherein, x, y, z respectively indicate three positions of the quadrotor under inertial coodinate system, UfIndicate that quadrotor flies
The input torque of row device, m are the quality of quadrotor, and g indicates acceleration of gravity；
Formula (1) is substituted into formula (2) to obtain:
Dynamic model in 1.3 quadrotor rotation processes are as follows:
Wherein, τx, τy, τzRespectively represent the moment components of each axis on body coordinate system, Ixx, Iyy, IzzRespectively indicate machine
The component of the rotary inertia of each axis under body coordinate system, × indicate multiplication cross, ωpIndicate rolling angular speed, ωqIndicate pitch angle
Speed, ωrIndicate yaw rate,Indicate rolling angular acceleration,Indicate pitching angular acceleration,Indicate that yaw angle adds
Speed；
In view of aircraft is in low-speed operations or floating state, it is believed that Therefore formula (4) is rewritten are as follows:
Joint type (3) and formula (5), obtain the kinetic model of quadrotor are as follows:
Wherein, ux=cos φ sin θ cos ψ+sin φ sin ψ, uy=cos φ sin θ sin ψ-sin φ cos ψ；
1.4, according to formula (6), define φ, the desired value of θ are as follows:
Wherein, φdFor the expected signal value of φ, θdFor θ expected signal value, arcsin is arcsin function；
Step 2, in each sampling instant, calculating position tracking error and its first derivative；Posture angle tracking is calculated to miss
Difference and its first derivative；Design position and posture angle controller, process are as follows:
2.1 define z tracking error and its first derivative:
Wherein, zdIndicate the desired signal of z；
2.2 design constraint liapunov functionsAnd solve its first derivative:
Wherein, Kb1For time-varying parameter, meet Kb1> | e1|,α1For virtual controlling amount, expression formula are as follows:
Wherein, k11For normal number；
Formula (10) are substituted into formula (9), are obtained:
Wherein,
2.3 design liapunov function V12Are as follows:
The first derivative of solution formula (12), obtains:
Wherein
Formula (14) and formula (6) are substituted into formula (13), obtained:
2.4 design Uf:
Wherein, k12For normal number；
2.5 define x, and y tracking error is respectively e2, e3, then have:
Wherein, xd, ydRespectively indicate x, the desired signal of y；
2.6 design constraint liapunov functions Its first derivative is solved respectively, is obtained:
Wherein, Kb2For time-varying parameter, meet Kb2> | e2|；Kb3For time-varying parameter, meet Kb3> | e3|；α2, α3For virtual controlling amount, expression formula are as follows:
Wherein, k21, k31For normal number；
Formula (19) are substituted into formula (18), are obtained:
Wherein,
2.7 design liapunov function V22, V32
The first derivative of solution formula (21), obtains:
Wherein
By formula (23), (6) substitute into formula (22), respectively:
2.8 by formula (24), and (25) separately design ux, uy:
Wherein, k22, k32For normal number；
2.9 define posture angle tracking error and its first derivative:
Wherein, j=4,5,6, x4=φ, x5=θ, x6=ψ, x4dIndicate the desired value of φ, x5dIndicate the desired value of θ, x6d
Indicate the desired value of ψ, e4Indicate the tracking error of φ, e5Indicate the tracking error of θ, e6Indicate the tracking error of ψ；
2.10 design constraint liapunov functionAnd solve its first derivative:
Wherein, kjFor time-varying parameter, meet Kbj> | ej|；αjFor the virtual controlling amount of attitude angle, table
Up to formula are as follows:
Wherein, kj1For normal number；
Formula (29) are substituted into formula (28), are obtained:
Wherein,
2.11 design constraint liapunov functions:
The first derivative of solution formula (31), obtains:
Wherein
Formula (33) and formula (6) are substituted into formula (32), respectively:
2.12 by formula (34), and (35), (36) separately design τx, τy, τz:
Wherein, k42, k52, k62For normal number；
Step 3, the stability of quadrotor system is verified, process is as follows:
Formula (16) are substituted into formula (15) by 3.1, are obtained:
Formula (26) are substituted into formula (24), (25) by 3.2, are obtained:
3.3 wushu (37) substitute into formula (34), (35), (36), obtain:
3.4 by (38), and (39), (40) know that quadrotor system is stable.
The present invention provides a kind of Spacecraft Attitude Control that symmetrical time-varying output is limited, improves the transient state of system
Can, reduce overshoot and arrival time.
Technical concept of the invention are as follows: for the dynamic system of quadrotor, design a kind of symmetrical time-varying output
Limited Spacecraft Attitude Control.The design of the symmetrical compound constraint liapunov function of time-varying logarithm tangent be in order to
The output of guarantee system can limit and avoid excessive overshoot in a certain range, while can also reduce arrival time.To
Improve the dynamic response performance of quadrotor system.
Advantage of the present invention are as follows: reduce overshoot, reduce arrival time, improve mapping.
Detailed description of the invention
Fig. 1 is position tracking effect diagram of the invention.
Fig. 2 is attitude angle tracking effect schematic diagram of the invention.
Fig. 3 is that positioner of the invention inputs schematic diagram.
Fig. 4 is that posture angle controller of the invention inputs schematic diagram.
Fig. 5 is control flow schematic diagram of the invention.
Specific embodiment
The present invention will be further described with reference to the accompanying drawing.
- Fig. 5 referring to Fig.1, a kind of Spacecraft Attitude Control that symmetrical time-varying output is limited, comprising the following steps:
Step 1, the dynamic model for establishing quadrotor system sets initial value, sampling time and the control of system
Parameter processed, process are as follows:
1.1 determine from the body coordinate system based on quadrotor system to the transfer square of the inertial coordinate based on the earth
Battle array T:
Wherein, φ, θ, ψ are roll angle, pitch angle, the yaw angle of quadrotor respectively, indicate aircraft successively around
The angle of each reference axis rotation of inertial coodinate system；
Dynamic model during the translation of 1.2 quadrotors is as follows:
Wherein, x, y, z respectively indicate three positions of the quadrotor under inertial coodinate system, UfIndicate that quadrotor flies
The input torque of row device, m are the quality of quadrotor, and g indicates acceleration of gravity；
Formula (1) is substituted into formula (2) to obtain:
Dynamic model in 1.3 quadrotor rotation processes are as follows:
Wherein, τx, τy, τzRespectively represent the moment components of each axis on body coordinate system, Ixx, Iyy, IzzRespectively indicate machine
The component of the rotary inertia of each axis under body coordinate system, × indicate multiplication cross, ωpIndicate rolling angular speed, ωqIndicate pitch angle
Speed, ωrIndicate yaw rate,Indicate rolling angular acceleration,Indicate pitching angular acceleration,Indicate that yaw angle adds
Speed；
In view of aircraft is in low-speed operations or floating state, attitude angle variation is smaller, it is believed thatTherefore formula (4) is rewritten are as follows:
Joint type (3) and formula (5), obtain the kinetic model of quadrotor are as follows:
Wherein, ux=cos φ sin θ cos ψ+sin φ sin ψ, uy=cos φ sin θ sin ψ-sin φ cos ψ；
1.4, according to formula (6), define φ, the desired value of θ are as follows:
Wherein, φdFor the expected signal value of φ, θdFor θ expected signal value, arcsin is arcsin function；
Step 2, in each sampling instant, calculating position tracking error and its first derivative；Posture angle tracking is calculated to miss
Difference and its first derivative；Design position and posture angle controller, process are as follows:
2.1 define z tracking error and its first derivative:
Wherein, zdIndicate the desired signal of z；
2.2 design constraint liapunov functionsAnd solve its first derivative:
Wherein, Kb1For time-varying parameter, meet Kb1> | e1|,α1For virtual controlling amount, expression formula are as follows:
Wherein, k11For normal number；
Formula (10) are substituted into formula (9), are obtained:
Wherein,
2.3 design liapunov function V12Are as follows:
The first derivative of solution formula (12), obtains:
Wherein
Formula (14) and formula (6) are substituted into formula (13), obtained:
2.4 design Uf:
Wherein, k12For normal number；
2.5 define x, and y tracking error is respectively e2, e3, then have:
Wherein, xd, ydRespectively indicate x, the desired signal of y；
2.6 design constraint liapunov functions Its first derivative is solved respectively, is obtained:
Wherein, Kb2For time-varying parameter, meet Kb2> | e2|；Kb3For time-varying parameter, meet Kb3> | e3|；α2, α3For virtual controlling amount, expression formula are as follows:
Wherein, k21, k31For normal number；
Formula (19) are substituted into formula (18), are obtained:
Wherein,
2.7 design liapunov function V22, V32
The first derivative of solution formula (21), obtains:
Wherein
By formula (23), (6) substitute into formula (22), respectively:
2.8 by formula (24), and (25) separately design ux, uy:
Wherein, k22, k32For normal number；
2.9 define posture angle tracking error and its first derivative:
Wherein, j=4,5,6, x4=φ, x5=θ, x6=ψ, x4dIndicate the desired value of φ, x5dIndicate the desired value of θ, x6d
Indicate the desired value of ψ, e4Indicate the tracking error of φ, e5Indicate the tracking error of θ, e6Indicate the tracking error of ψ；
2.10 design constraint liapunov functionAnd solve its first derivative:
Wherein, kjFor time-varying parameter, meet Kbj> | ej|；αjFor the virtual controlling amount of attitude angle, table
Up to formula are as follows:
Wherein, kj1For normal number；
Formula (29) are substituted into formula (28), are obtained:
Wherein,
2.11 design constraint liapunov functions:
The first derivative of solution formula (31), obtains:
Wherein
Formula (33) and formula (6) are substituted into formula (32), respectively:
2.12 by formula (34), and (35), (36) separately design τx, τy, τz:
Wherein, k42, k52, k62For normal number；
Step 3, the stability of quadrotor system is verified, process is as follows:
Formula (16) are substituted into formula (15) by 3.1, are obtained:
Formula (26) are substituted into formula (24), (25) by 3.2, are obtained:
3.3 wushu (37) substitute into formula (34), (35), (36), obtain:
3.4 by (38), and (39), (40) know that quadrotor system is stable.
In order to verify the feasibility of proposed method, the emulation knot that The present invention gives the control methods on MATLAB platform
Fruit:
Parameter is given below: m=1.1kg, g=9.81N/kg in formula (2)；In formula (4), Ixx=1.22kgm2, Iyy=
1.22kg·m2, Izz=2.2kgm2；Z in formula (8), formula (17) and formula (27)d=0.5, xd=0.5, yd=0.5, ψd=0.5；
K in formula (10), formula (19) and formula (29)11=2, k21=2, k31=2；k41=2, k51=2, k61=2；Formula (16), formula (26) and
K in formula (37)12=2, k22=2, k32=2, k42=2, k52=2, k62=2；Formula (9), formula (18) and formula (28) kb1=2.5+
0.1sint, kb2=2.5+0.1sint, kb3=2.5+0.1sint, kb4=2+0.1sint, kb5=2+0.1sint, kb6=2+
0.1sint。
From Fig. 1 and Fig. 2 it is found that system has good transient response, arrival time is 5.81 seconds, overshoot 0.
In conclusion the Spacecraft Attitude Control that symmetrical time-varying output is limited can effectively improve quadrotor
The mapping of system.
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 Spacecraft Attitude Control that symmetrical time-varying output is limited, which comprises the following steps:
Step 1, the dynamic model for establishing quadrotor system sets initial value, sampling time and the control ginseng of system
Number, process are as follows:
1.1 determine from the body coordinate system based on quadrotor system to the transfer matrix T of the inertial coordinate based on the earth:
Wherein, φ, θ, ψ are roll angle, pitch angle, the yaw angle of quadrotor respectively, indicate aircraft successively around inertia
The angle of each reference axis rotation of coordinate system；
Dynamic model during the translation of 1.2 quadrotors is as follows:
Wherein, x, y, z respectively indicate three positions of the quadrotor under inertial coodinate system, UfIndicate quadrotor
Input torque, m be quadrotor quality, g indicate acceleration of gravity；
Formula (1) is substituted into formula (2) to obtain:
Dynamic model in 1.3 quadrotor rotation processes are as follows:
Wherein, τx, τy, τzRespectively represent the moment components of each axis on body coordinate system, Ixx, Iyy, IzzRespectively indicate body seat
The component of the rotary inertia of each axis under mark system, × indicate multiplication cross, ωpIndicate rolling angular speed, ωqIndicate rate of pitch,
ωrIndicate yaw rate,Indicate rolling angular acceleration,Indicate pitching angular acceleration,Indicate yaw angular acceleration；
In view of aircraft is in low-speed operations or floating state, it is believed that Therefore formula (4) is rewritten are as follows:
Joint type (3) and formula (5), obtain the kinetic model of quadrotor are as follows:
Wherein, ux=cos φ sin θ cos ψ+sin φ sin ψ, uy=cos φ sin θ sin ψ-sin φ cos ψ；
1.4, according to formula (6), define φ, the desired value of θ are as follows:
Wherein, φdFor the expected signal value of φ, θdFor θ expected signal value, arcsin is arcsin function；
Step 2, in each sampling instant, calculating position tracking error and its first derivative；Calculate posture angle tracking error and
Its first derivative；Design position and posture angle controller, process are as follows:
2.1 define z tracking error and its first derivative:
Wherein, zdIndicate the desired signal of z；
2.2 design constraint liapunov functionsAnd solve its first derivative:
Wherein, Kb1For time-varying parameter, meet Kb1> | e1|,α1For virtual controlling amount, expression formula are as follows:
Wherein, k11For normal number；
Formula (10) are substituted into formula (9), are obtained:
Wherein,
2.3 design liapunov function V12Are as follows:
The first derivative of solution formula (12), obtains:
Wherein
Formula (14) and formula (6) are substituted into formula (13), obtained:
2.4 design Uf:
Wherein, k12For normal number；
2.5 define x, and y tracking error is respectively e2, e3, then have:
Wherein, xd, ydRespectively indicate x, the desired signal of y；
2.6 design constraint liapunov functions Its first derivative is solved respectively, is obtained:
Wherein, Kb2For time-varying parameter, meet Kb2> | e2|；Kb3For time-varying parameter, meet Kb3> | e3|；α2, α3For virtual controlling amount, expression formula are as follows:
Wherein, k21, k31For normal number；
Formula (19) are substituted into formula (18), are obtained:
Wherein,
2.7 design liapunov function V22, V32
The first derivative of solution formula (21), obtains:
Wherein
By formula (23), (6) substitute into formula (22), respectively:
2.8 by formula (24), and (25) separately design ux, uy:
Wherein, k22, k32For normal number；
2.9 define posture angle tracking error and its first derivative:
Wherein, j=4,5,6, x4=φ, x5=θ, x6=ψ, x4dIndicate the desired value of φ, x5dIndicate the desired value of θ, x6dIndicate ψ
Desired value, e4Indicate the tracking error of φ, e5Indicate the tracking error of θ, e6Indicate the tracking error of ψ；
2.10 design constraint liapunov functionAnd solve its first derivative:
Wherein, kjFor time-varying parameter, meet Kbj> | ej|；αjFor the virtual controlling amount of attitude angle, expression formula
Are as follows:
Wherein, kj1For normal number；
Formula (29) are substituted into formula (28), are obtained:
Wherein,
2.11 design constraint liapunov functions:
The first derivative of solution formula (31), obtains:
Wherein
Formula (33) and formula (6) are substituted into formula (32), respectively:
2.12 by formula (34), and (35), (36) separately design τx, τy, τz:
Wherein, k42, k52, k62For normal number；
Step 3, the stability of quadrotor system is verified, process is as follows:
Formula (16) are substituted into formula (15) by 3.1, are obtained:
Formula (26) are substituted into formula (24), (25) by 3.2, are obtained:
3.3 wushu (37) substitute into formula (34), (35), (36), obtain:
3.4 by (38), and (39), (40) know that quadrotor system is stable.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810212226.6A CN108549217A (en) | 2018-03-15 | 2018-03-15 | Quadrotor based on the compound constraint liapunov function of symmetrical time-varying logarithm tangent exports constrained control method |
| CN2018102122266 | 2018-03-15 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN109917651A true CN109917651A (en) | 2019-06-21 |
Family
ID=63516276
Family Applications (2)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201810212226.6A Withdrawn CN108549217A (en) | 2018-03-15 | 2018-03-15 | Quadrotor based on the compound constraint liapunov function of symmetrical time-varying logarithm tangent exports constrained control method |
| CN201910191632.3A Withdrawn CN109917651A (en) | 2018-03-15 | 2019-03-14 | A kind of Spacecraft Attitude Control that symmetrical time-varying output is limited |
Family Applications Before (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201810212226.6A Withdrawn CN108549217A (en) | 2018-03-15 | 2018-03-15 | Quadrotor based on the compound constraint liapunov function of symmetrical time-varying logarithm tangent exports constrained control method |
Country Status (1)
| Country | Link |
|---|---|
| CN (2) | CN108549217A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112327897A (en) * | 2020-11-04 | 2021-02-05 | 江苏师范大学 | Four-rotor unmanned aerial vehicle attitude control method with input dead zone |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106094855A (en) * | 2016-07-27 | 2016-11-09 | 浙江工业大学 | Terminal cooperative control method for quad-rotor unmanned aerial vehicle |
| CN106681154A (en) * | 2017-03-01 | 2017-05-17 | 重庆大学 | Self-adaptation control method for electric vehicle for uncertain barycenter and unknown input saturation |
| CN107368088A (en) * | 2017-07-11 | 2017-11-21 | 浙江工业大学 | Four-rotor aircraft nonlinear sliding mode pose control method based on error exponential function |
| CN107368089A (en) * | 2017-07-11 | 2017-11-21 | 浙江工业大学 | Nonlinear sliding mode pose control method of quadrotor aircraft based on double exponential function |
| CN107450584A (en) * | 2017-08-29 | 2017-12-08 | 浙江工业大学 | Aircraft self-adaptive attitude control method based on fixed time sliding mode |
| CN107561931A (en) * | 2017-07-11 | 2018-01-09 | 浙江工业大学 | Nonlinear sliding mode pose control method of quadrotor aircraft based on single exponential function |
| 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 |
-
2018
- 2018-03-15 CN CN201810212226.6A patent/CN108549217A/en not_active Withdrawn
-
2019
- 2019-03-14 CN CN201910191632.3A patent/CN109917651A/en not_active Withdrawn
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106094855A (en) * | 2016-07-27 | 2016-11-09 | 浙江工业大学 | Terminal cooperative control method for quad-rotor unmanned aerial vehicle |
| CN106681154A (en) * | 2017-03-01 | 2017-05-17 | 重庆大学 | Self-adaptation control method for electric vehicle for uncertain barycenter and unknown input saturation |
| CN107368088A (en) * | 2017-07-11 | 2017-11-21 | 浙江工业大学 | Four-rotor aircraft nonlinear sliding mode pose control method based on error exponential function |
| CN107368089A (en) * | 2017-07-11 | 2017-11-21 | 浙江工业大学 | Nonlinear sliding mode pose control method of quadrotor aircraft based on double exponential function |
| CN107561931A (en) * | 2017-07-11 | 2018-01-09 | 浙江工业大学 | Nonlinear sliding mode pose control method of quadrotor aircraft based on single exponential function |
| CN107450584A (en) * | 2017-08-29 | 2017-12-08 | 浙江工业大学 | Aircraft self-adaptive attitude control method based on fixed time sliding mode |
| 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 |
Non-Patent Citations (3)
| Title |
|---|
| JACEK KABZIŃSK,等: "Comparison of Various Barrier Lyapynov Functions for Adaptive Control of Nonlinear Systems", 《2016 21ST INTERNATIONAL CONFERENCE ON METHODS AND MODELS IN AUTOMATION AND ROBOTICS (MMAR)》 * |
| 陈乐剑，等: "基于BarrierLyapunov函数的模糊自适应迭代学习控制", 《第27届中国控制与决策会议论文集》 * |
| 陈强，等: "基于全阶滑模的四旋翼无人机有限时间控制", 《第七届中国航空学会青年科技论坛文集》 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112327897A (en) * | 2020-11-04 | 2021-02-05 | 江苏师范大学 | Four-rotor unmanned aerial vehicle attitude control method with input dead zone |
| CN112327897B (en) * | 2020-11-04 | 2022-07-29 | 江苏师范大学 | Quad-rotor unmanned aerial vehicle attitude control method with input dead zone |
Also Published As
| Publication number | Publication date |
|---|---|
| CN108549217A (en) | 2018-09-18 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN108037662B (en) | A kind of limited backstepping control method of quadrotor output based on Integral Sliding Mode obstacle liapunov function | |
| CN107831671B (en) | A kind of limited backstepping control method of quadrotor output based on asymmetric time-varying obstacle liapunov function | |
| CN107831670B (en) | It is a kind of based on it is asymmetric when constant obstacle liapunov function quadrotor export limited backstepping control method | |
| CN108267961A (en) | Quadrotor total state constrained control method based on symmetrical time-varying tangential type constraint liapunov function | |
| CN109917650A (en) | A kind of Spacecraft Attitude Control of asymmetric varying constraint | |
| CN109932902A (en) | A kind of quadrotor output constrained control method | |
| CN108388119A (en) | Based on it is symmetrical when constant tangential type constrain the quadrotor total state constrained control method of liapunov function | |
| CN109917651A (en) | A kind of Spacecraft Attitude Control that symmetrical time-varying output is limited | |
| CN108052115B (en) | It is a kind of based on it is asymmetric when constant obstacle liapunov function quadrotor total state be limited backstepping control method | |
| CN109870912A (en) | It is a kind of using it is asymmetric when constraint independent of time function Spacecraft Attitude Control | |
| CN109870913A (en) | It is a kind of with when varying index tangent constraint aircraft total state constrained control method | |
| CN109613829A (en) | A kind of quadrotor total state constrained control method | |
| CN108333950A (en) | Quadrotor based on the compound constraint liapunov function of symmetrical time-varying tangent cosine exports constrained control method | |
| CN108536162A (en) | Based on it is symmetrical when the not compound constraint liapunov function of varying index tangent quadrotor total state constrained control method | |
| CN108427279A (en) | Based on it is symmetrical when the compound constraint liapunov function of varying index tangent quadrotor export constrained control method | |
| CN108549218A (en) | Based on it is symmetrical when the constant compound constraint liapunov function of tangent cosine quadrotor export constrained control method | |
| CN108303892A (en) | The quadrotor that liapunov function is constrained based on asymmetric time-varying tangential type exports constrained control method | |
| CN109870911A (en) | A kind of quadrotor tangent cosine output constrained control method | |
| CN109917649A (en) | A kind of aircraft arc tangent output constraint control method | |
| CN109828470A (en) | A kind of quadrotor index tangent output constraint control method | |
| CN109917652A (en) | A kind of aircraft logarithm tangent total state constrained control method | |
| CN108594647A (en) | Quadrotor total state constrained control method based on the compound constraint liapunov function of symmetrical time-varying logarithm tangent | |
| CN108563116A (en) | Based on it is asymmetric when constant cut type anyway constrain the quadrotor of liapunov function and export constrained control method | |
| CN108563115A (en) | Based on it is symmetrical when the constant compound constraint liapunov function of logarithm tangent quadrotor total state constrained control method | |
| CN108427274A (en) | Quadrotor based on the compound constraint liapunov function of symmetrical time-varying logarithm secant exports constrained control method |
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 | ||
| WW01 | Invention patent application withdrawn after publication |
Application publication date: 20190621 |
|
| WW01 | Invention patent application withdrawn after publication |