CN109144087A - A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem - Google Patents

A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem Download PDF

Info

Publication number
CN109144087A
CN109144087A CN201811119494.XA CN201811119494A CN109144087A CN 109144087 A CN109144087 A CN 109144087A CN 201811119494 A CN201811119494 A CN 201811119494A CN 109144087 A CN109144087 A CN 109144087A
Authority
CN
China
Prior art keywords
set time
follows
rigid aircraft
sat
actuator
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Withdrawn
Application number
CN201811119494.XA
Other languages
Chinese (zh)
Inventor
陈强
谢树宗
孙明轩
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.)
Zhejiang University of Technology ZJUT
Original Assignee
Zhejiang University of Technology ZJUT
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 Zhejiang University of Technology ZJUT filed Critical Zhejiang University of Technology ZJUT
Priority to CN201811119494.XA priority Critical patent/CN109144087A/en
Publication of CN109144087A publication Critical patent/CN109144087A/en
Priority to CN201910878603.4A priority patent/CN110515389B/en
Withdrawn legal-status Critical Current

Links

Classifications

    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • 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

Abstract

A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem, for with external disturbance, rotary inertia is uncertain, the rigid aircraft attitude stabilization problem of actuator saturation and failure, using sliding-mode control, in conjunction with adaptive technique, nonsingular set time adaptive controller is devised；The design of nonsingular set time sliding-mode surface not only guarantees the set time convergence of system mode, but also solves singular value problem；It is not always known in addition, adaptive updates rule is used to estimating system, therefore always uncertain upper bound information is not necessarily to be known in advance.For the present invention in external interference, rotary inertia is uncertain, under the factor of actuator saturation and failure, realizes the control of the set time uniform ultimate bounded of system mode.

Description

A kind of adaptive set time posture of rigid aircraft considering actuator constraints problem Calm method
Technical field
The present invention relates to a kind of calm sides of adaptive set time posture of rigid aircraft for considering actuator constraints problem Method is especially in the presence of external disturbance, and moment of inertia matrix is uncertain, and the rigid aircraft posture of actuator saturation and failure is calm Method.
Background technique
Rigid aircraft attitude control system reliably plays important angle in movement in the health of rigid aircraft Color.In complicated space environment, rigid aircraft attitude control system will receive various external disturbances and rigid aircraft In long-term continuous task, failures such as existing aging and failure etc. influence.In order to effectively maintain the performance of system, need to make it There is stronger robustness to external disturbance and actuator failures；In addition, rigid aircraft there is also moment of inertia matrix not It determines, therefore controlling saturation is also the frequent problem of aircraft.In conclusion rigid aircraft in the task of execution, needs Want a kind of makes system stable convergence, high-precision fault tolerant control method in a short time.
Sliding formwork control is considered as an effective robust control side in terms of solving systematic uncertainty and external disturbance Method.Sliding-mode control has algorithm simple, fast response time, excellent to extraneous noise jamming and Parameter Perturbation strong robustness etc. Point.TSM control is a kind of improvement project of traditional sliding formwork control that stability in finite time may be implemented.However, existing Finite time technology estimation convergence time need to know the initial information of system, this is difficult to know for designer.Closely Nian Lai, set time technology are widely used, set time control method and existing finite-time control method phase Than, have without knowing the initial information of system, also can conservative estimation system convergence time superiority.
Self adaptive control refers to the dynamic that controller can correct itself control parameter to adapt to system itself with external disturbance Characteristic makes system be optimal control to obtain satisfied dynamic property.This method is not only suitable for linear system, is also applied for Nonlinear system is controlled mainly for the uncertainty of system.The research object of self adaptive control is that have to a certain degree The uncertain and easy system by external environmental interference.For these reasons, many self-adaptation control methods are used to control Spacecraft system.
Therefore, set time sliding formwork control technology and self-adaptation control method are effectively combined, reduce external disturbance and The set time control of rigid aircraft posture is realized in influence of the system parameter uncertainty to system control performance.
Summary of the invention
In order to overcome the problems, such as unknown nonlinear existing for existing rigid aircraft attitude control system, the present invention provides one Kind considers the calm method of the adaptive set time posture of rigid aircraft of actuator constraints problem, there is external do in system It disturbs, rotary inertia is uncertain, in the case where actuator saturation and failure, realizes the set time uniform ultimate bounded of system mode Control.
In order to solve the above-mentioned technical problem technical solution proposed by the present invention is as follows:
A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem, including it is following Step:
Step 1, the kinematics and dynamics modeling of rigid aircraft system is established, system mode and control ginseng are initialized Number, process are as follows:
The kinematical equation of 1.1 rigid aircraft systems are as follows:
Wherein qv=[q1,q2,q3]TAnd q4The respectively vector section and scalar component and satisfaction of unit quaternionq1,q2,q3Respectively it is mapped in rectangular coordinate system in space x, y, the value in z-axis；It is q respectivelyvAnd q4's Derivative；For qvTransposition；Ω∈R3It is the angular speed of rigid aircraft；I3It is R3×3Unit matrix；It indicates are as follows:
The kinetics equation of 1.2 rigid aircraft systems are as follows:
Wherein J ∈ R3×3It is the rotator inertia matrix of rigid aircraft；It is the angular acceleration of rigid aircraft；u∈ R3With d ∈ R3It is control moment and external disturbance；D=diag (D1,D2,D3)∈R3×3It is 3 × 3 symmetrical diagonal actuator efficiency Matrix meets 0 < Di(t)≤1, i=1,2,3；Sat (u)=[sat (u1),sat(u2),sat(u3)]TIt is generated for actuator Practical control moment, sat (ui) it is the actuator with saturated characteristic, it is expressed as sat (ui)=sgn (ui)min{umi,|ui|, umiFor the control moment that maximum provides, sgn (ui) it is sign function, min { umi,|ui| be both minimum value；In order to more square Just expression control constraints, sat (u) are expressed as sat (u)=Θ (u) u, Θ (u)=diag (Θ1(u),Θ2(u),Θ3(u)) ∈R3×3For 3 × 3 symmetrical diagonal matrix, Θi(u) it indicates are as follows:
Meet 0 < ξ≤min (DiΘi(u))≤1, min (DiΘiIt (u)) is minimum value, i=1,2,3, ξ be unknown normal Number；Ω×It indicates are as follows:
1.3 rotator inertia matrix Js meet J=J0+ Δ J, wherein J0With Δ J respectively indicate J nominal section and uncertain portion Point, then formula (4) is write as again:
Further obtain:
1.4 pairs of formulas (1) carry out differential, obtain:
Wherein ΩTFor the transposition of Ω；For qvSecond dervative；For J0It is inverse；It indicates are as follows:
Respectively q1,q2,q3Derivative；
Step 2, for external disturbance is had, rotary inertia is uncertain, the rigid aircraft system of actuator saturation and failure System, the sliding-mode surface of design, process are as follows:
Select nonsingular set time sliding-mode surface are as follows:
Wherein, With sgn (qi) it is sign function, λ1> 0, λ2> 0, a2> 1, For qiLead Number, i=1,2,3；
Step 3, nonsingular adaptive set time controller is designed, process is as follows:
3.1 design set time controllers are as follows:
Wherein S=[S1,S2,S3]T,Γ=diag (Γ1,Γ2,Γ3)∈ R3×3For 3 × 3 symmetrical diagonal matrix；K1=diag (k11,k12,k13)∈R3×3For 3 × 3 symmetrical diagonal matrix, K2=diag (k21,k22,k23)∈R3×3It is symmetrical diagonal for 3 × 3 Matrix, K3=diag (k31,k32,k33)∈R3×3For 3 × 3 symmetrical diagonal matrix, k11,k12,k13,k21,k22,k23,k31,k32, k33For normal number,Respectively c1,c2,c3, the estimation of ξ；ηs>=1, c1,c2,c3For unknown normal number；
The more new law of 3.2 design auto-adaptive parameters:
Wherein η1,η2,η3,η4,ε1,ε2,ε3,ε4For normal number；RespectivelyDerivative；ForTwo norms,ForTwo norms, | | Ω | | be Ω two norms；
Step 4, set time stability proves that process is as follows:
4.1 prove that all signals of rigid aircraft system are all uniform ultimate boundeds, and design liapunov function is such as Lower form:
WhereinSTIt is the transposition of S；
Derivation is carried out to formula (18), is obtained:
Wherein k3min=min { k31,k32,k33, min { } indicates minimum Value；For the derivative of S；δ1,δ2,δ3,δ4For normal number；
Then determine that all signals of rigid aircraft system are all uniform ultimate boundeds；
4.2 prove set time convergence, and design liapunov function is following form:
Derivation is carried out to formula (20), is obtained:
Wherein γ2It is greater than zero upper dividing value for one；
Based on the above analysis, rigid aircraft system mode is in set time uniform ultimate bounded.
For the present invention in external interference, rotary inertia is uncertain, under the factor of actuator saturation and failure, with adaptive solid Posture of fixing time is calmed method, realizes system stability contorting, guarantees that system mode realizes set time uniform ultimate bounded.This hair Bright technical concept are as follows: for external disturbance is contained, rotary inertia is uncertain, the rigid aircraft system of actuator saturation and failure System, devises nonsingular adaptive set time controller in conjunction with self adaptive control using sliding-mode control.It is nonsingular The design of set time sliding-mode surface not only guarantees the set time convergence of system mode, but also solves singular value problem.In addition, It is restrained based on designed adaptive updates, it is not necessary that total uncertain upper bound information is known in advance.The present invention exists extraneous in system Under interference and the uncertain situation of rotary inertia, the control method of the set time uniform ultimate bounded of system mode is realized.
The invention has the benefit that designed set time sliding-mode surface effective solution singular value problem；It is being There are external interferences for system, and rotary inertia is uncertain, in the case where actuator saturation and failure, realize the set time of system mode Uniform ultimate bounded, and convergence time is unrelated with the original state of system.
Detailed description of the invention
Fig. 1 is rigid aircraft attitude quaternion schematic diagram of the invention；
Fig. 2 is rigid aircraft angular speed schematic diagram of the invention；
Fig. 3 is rigid aircraft sliding-mode surface schematic diagram of the invention；
Fig. 4 is rigid aircraft control moment schematic diagram of the invention；
Fig. 5 is rigid aircraft parameter Estimation schematic diagram of the invention, wherein (a) is parameter c1,c2,c3Estimation；(b) For the estimation of parameter ξ；
Fig. 6 is control flow schematic diagram of the invention.
Specific embodiment
The present invention will be further described with reference to the accompanying drawing.
- Fig. 6 referring to Fig.1, a kind of adaptive set time posture of rigid aircraft considering actuator constraints problem are calm Method the described method comprises the following steps:
Step 1, the kinematics and dynamics modeling of rigid aircraft system is established, system mode and control ginseng are initialized Number, process are as follows:
The kinematical equation of 1.1 rigid aircraft systems are as follows:
Wherein qv=[q1,q2,q3]TAnd q4The respectively vector section and scalar component and satisfaction of unit quaternionq1,q2,q3Respectively it is mapped in rectangular coordinate system in space x, y, the value in z-axis；It is q respectivelyvAnd q4's Derivative；For qvTransposition；Ω∈R3It is the angular speed of rigid aircraft；I3It is R3×3Unit matrix；It indicates are as follows:
The kinetics equation of 1.2 rigid aircraft systems are as follows:
Wherein J ∈ R3×3It is the rotator inertia matrix of rigid aircraft；It is the angular acceleration of rigid aircraft；u∈ R3With d ∈ R3It is control moment and external disturbance；D=diag (D1,D2,D3)∈R3×3It is 3 × 3 symmetrical diagonal actuator efficiency Matrix meets 0 < Di(t)≤1, i=1,2,3；Sat (u)=[sat (u1),sat(u2),sat(u3)]TIt is generated for actuator Practical control moment, sat (ui) it is the actuator with saturated characteristic, it is expressed as sat (ui)=sgn (ui)min{umi,|ui|, umiFor the control moment that maximum provides, sgn (ui) it is sign function, min { umi,|ui| be both minimum value；In order to more square Just expression control constraints, sat (u) are expressed as sat (u)=Θ (u) u, Θ (u)=diag (Θ1(u),Θ2(u),Θ3(u)) ∈R3×3For 3 × 3 symmetrical diagonal matrix, Θi(u) it indicates are as follows:
Meet 0 < ξ≤min (DiΘi(u))≤1, min (DiΘiIt (u)) is minimum value, i=1,2,3, ξ be unknown normal Number；Ω×It indicates are as follows:
1.3 rotator inertia matrix Js meet J=J0+ Δ J, wherein J0With Δ J respectively indicate J nominal section and uncertain portion Point, then formula (4) is write as again:
Further obtain:
1.4 pairs of formulas (1) carry out differential, obtain:
Wherein ΩTFor the transposition of Ω；For qvSecond dervative；For J0It is inverse；It indicates are as follows:
Respectively q1,q2,q3Derivative；
Step 2, for external disturbance is had, rotary inertia is uncertain, the rigid aircraft system of actuator saturation and failure System, the sliding-mode surface of design, process are as follows:
Select nonsingular set time sliding-mode surface are as follows:
Wherein, With sgn (qi) it is sign function, λ1> 0, λ2> 0, a2> 1, For qiLead Number, i=1,2,3；
Step 3, nonsingular adaptive set time controller is designed, process is as follows:
3.1 design set time controllers are as follows:
Wherein S=[S1,S2,S3]T,Γ=diag (Γ1,Γ2,Γ3)∈ R3×3For 3 × 3 symmetrical diagonal matrix；K1=diag (k11,k12,k13)∈R3×3For 3 × 3 symmetrical diagonal matrix, K2=diag (k21,k22,k23)∈R3×3It is symmetrical diagonal for 3 × 3 Matrix, K3=diag (k31,k32,k33)∈R3×3For 3 × 3 symmetrical diagonal matrix, k11,k12,k13,k21,k22,k23,k31,k32, k33For normal number,Respectively c1,c2,c3, the estimation of ξ；ηs>=1, c1,c2,c3For unknown normal number；
The more new law of 3.2 design auto-adaptive parameters:
Wherein η1,η2,η3,η4,ε1,ε2,ε3,ε4For normal number；RespectivelyDerivative；For Two norms,ForTwo norms, | | Ω | | be Ω two norms；
Step 4, set time stability proves that process is as follows:
4.1 prove that all signals of rigid aircraft system are all uniform ultimate boundeds, and design liapunov function is such as Lower form:
WhereinSTIt is the transposition of S；
Derivation is carried out to formula (18), is obtained:
Wherein k3min=min { k31,k32,k33, min { } indicates minimum Value；For the derivative of S；δ1,δ2,δ3,δ4For normal number；
Then determine that all signals of rigid aircraft system are all uniform ultimate boundeds；
4.2 prove set time convergence, and design liapunov function is following form:
Derivation is carried out to formula (20), is obtained:
Wherein γ2It is greater than zero upper dividing value for one；
Based on the above analysis, rigid aircraft system mode is in set time uniform ultimate bounded.
For the validity for verifying proposed method, this method carries out simulating, verifying for rigid aircraft system.System is initial It is as follows to change parameter setting:
The initial value of system: q (0)=[0.3, -0.2, -0.3,0.8832]T, Ω (0)=[1,0, -1]TRadian per second；Turn The nominal section J of dynamic inertial matrix0=[40,1.2,0.9；1.2,17,1.4；0.9,1.4,15] kilogram * square metres, the moment of inertia Uncertain portion's Δ J=diag [sin (0.1t), 2sin (0.2t), 3sin (0.3t)] of battle array；External disturbance d (t)=[0.2sin (0.1t),0.3sin(0.2t),0.5sin(0.2t)]T* meters of ox；The parameter of sliding-mode surface is as follows: λ1=1, λ2=1, a1=1.5, a2 =1.5；The parameter of controller is as follows:K1=K2=K3=I3；More new law parameter is as follows: ηi=0.1, εi= 0.001, i=1,2,3,4, ηs=30,Maximum control moment umi =25 Ns * meters, the selection of actuator efficiency value are as follows:
Wherein i=1,2,3.
The attitude quaternion of rigid aircraft and the response schematic diagram difference of angular speed are as depicted in figs. 1 and 2, it can be seen that Attitude quaternion and angular speed can converge in zero domain of equalization point at 7 seconds or so；The sliding-mode surface of rigid aircraft is rung Answer schematic diagram as shown in Figure 3, it can be seen that sliding-mode surface can converge in zero domain of equalization point at 5 seconds or so；Rigidity flight The control moment of device is as shown in Figure 4, it can be seen that control moment clipping is in 25 Ns * meters；Parameter Estimation responds schematic diagram difference As shown in Fig. 5 (a) and Fig. 5 (b).
Therefore, the present invention designs set time sliding-mode surface effective solution singular value problem；Exist in system extraneous Interference, rotary inertia is uncertain, in the case where actuator saturation and failure, realizes that the set time of system mode unanimously finally has Boundary, and convergence time is unrelated with the original state 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 method 1. adaptive set time posture of rigid aircraft for considering actuator constraints problem is calmed, feature exist In: it the described method comprises the following steps:
    Step 1, the kinematics and dynamics modeling of rigid aircraft system is established, system mode and control parameter are initialized, Process is as follows:
    The kinematical equation of 1.1 rigid aircraft systems are as follows:
    Wherein qv=[q1,q2,q3]TAnd q4The respectively vector section and scalar component and satisfaction of unit quaternionq1,q2,q3Respectively it is mapped in rectangular coordinate system in space x, y, the value in z-axis；It is q respectivelyvAnd q4's Derivative；For qvTransposition；Ω∈R3It is the angular speed of rigid aircraft；I3It is R3×3Unit matrix；It indicates are as follows:
    The kinetics equation of 1.2 rigid aircraft systems are as follows:
    Wherein J ∈ R3×3It is the rotator inertia matrix of rigid aircraft；It is the angular acceleration of rigid aircraft；u∈R3And d ∈R3It is control moment and external disturbance；D=diag (D1,D2,D3)∈R3×3It is 3 × 3 symmetrical diagonal actuator efficiency matrix, Meet 0 < Di(t)≤1, i=1,2,3；Sat (u)=[sat (u1),sat(u2),sat(u3)]TThe practical control generated for actuator Torque processed, sat (ui) it is the actuator with saturated characteristic, it is expressed as sat (ui)=sgn (ui)min{umi,|ui|, umiFor most The control moment provided greatly, sgn (ui) it is sign function, min { umi,|ui| be both minimum value；For more easily table Show that control constraints, sat (u) are expressed as sat (u)=Θ (u) u, Θ (u)=diag (Θ1(u),Θ2(u),Θ3(u))∈R3×3For 3 × 3 symmetrical diagonal matrix, Θi(u) it indicates are as follows:
    Meet 0 < ξ≤min (DiΘi(u))≤1, min (DiΘiIt (u)) is minimum value, i=1,2,3, ξ be unknown normal number；Ω× It indicates are as follows:
    1.3 rotator inertia matrix Js meet J=J0+ Δ J, wherein J0With Δ J respectively indicate J nominal section and uncertain part, Then formula (4) is write as again:
    Further obtain:
    1.4 pairs of formulas (1) carry out differential, obtain:
    Wherein ΩTFor the transposition of Ω；For qvSecond dervative；For J0It is inverse；It indicates are as follows:
    Respectively q1,q2,q3Derivative；
    Step 2, for external disturbance is had, rotary inertia is uncertain, the rigid aircraft system of actuator saturation and failure, if Sliding-mode surface needed for meter, process are as follows:
    Select nonsingular set time sliding-mode surface are as follows:
    Wherein,
    With sgn (qi) it is sign function, λ1> 0, λ2> 0, a2> 1, For qiDerivative, i=1,2,3；
    Step 3, nonsingular adaptive set time controller is designed, process is as follows:
    3.1 design set time controllers are as follows:
    Wherein S=[S1,S2,S3]T,Γ=diag (Γ1,Γ2,Γ3)∈R3×3 For 3 × 3 symmetrical diagonal matrix；0 < r1< 1, r2> 1, K1=diag (k11, k12,k13)∈R3×3For 3 × 3 symmetrical diagonal matrix, K2=diag (k21,k22,k23)∈R3×3For 3 × 3 symmetrically to angular moment Battle array, K3=diag (k31,k32,k33)∈R3×3For 3 × 3 symmetrical diagonal matrix, k11,k12,k13,k21,k22,k23,k31,k32,k33 For normal number,Respectively c1,c2,c3, the estimation of ξ；ηs>=1, c1,c2,c3For unknown normal number；
    The more new law of 3.2 design auto-adaptive parameters:
    Wherein η1,η2,η3,η4,ε1,ε2,ε3,ε4For normal number；RespectivelyDerivative；ForTwo Norm,ForTwo norms, | | Ω | | be Ω two norms；
    Step 4, set time stability proves that process is as follows:
    4.1 prove that all signals of rigid aircraft system are all uniform ultimate boundeds, and design liapunov function is following shape Formula:
    WhereinSTIt is the transposition of S；
    Derivation is carried out to formula (18), is obtained:
    Wherein k3min=min { k31,k32,k33, min { } indicates minimum Value；For the derivative of S；δ1,δ2,δ3,δ4For normal number；
    Then determine that all signals of rigid aircraft system are all uniform ultimate boundeds；
    4.2 prove set time convergence, and design liapunov function is following form:
    Derivation is carried out to formula (20), is obtained:
    Wherein γ2It is greater than zero upper dividing value for one；
    Based on the above analysis, rigid aircraft system mode is in set time uniform ultimate bounded.
CN201811119494.XA 2018-09-25 2018-09-25 A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem Withdrawn CN109144087A (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
CN201811119494.XA CN109144087A (en) 2018-09-25 2018-09-25 A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem
CN201910878603.4A CN110515389B (en) 2018-09-25 2019-09-18 Rigid aircraft self-adaptive fixed-time attitude stabilization method considering actuator limitation problem

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201811119494.XA CN109144087A (en) 2018-09-25 2018-09-25 A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem

Publications (1)

Publication Number Publication Date
CN109144087A true CN109144087A (en) 2019-01-04

Family

ID=64812275

Family Applications (2)

Application Number Title Priority Date Filing Date
CN201811119494.XA Withdrawn CN109144087A (en) 2018-09-25 2018-09-25 A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem
CN201910878603.4A Active CN110515389B (en) 2018-09-25 2019-09-18 Rigid aircraft self-adaptive fixed-time attitude stabilization method considering actuator limitation problem

Family Applications After (1)

Application Number Title Priority Date Filing Date
CN201910878603.4A Active CN110515389B (en) 2018-09-25 2019-09-18 Rigid aircraft self-adaptive fixed-time attitude stabilization method considering actuator limitation problem

Country Status (1)

Country Link
CN (2) CN109144087A (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110244747A (en) * 2019-08-02 2019-09-17 大连海事大学 Heterogeneous fleet fault-tolerant control method based on actuator fault and saturation
CN112180960A (en) * 2020-09-29 2021-01-05 西北工业大学 Unmanned aerial vehicle fault-tolerant flight method and flight system for actuator faults

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
KR101434310B1 (en) * 2013-05-22 2014-08-26 현대로템 주식회사 Stabilization enhancement method of drive unit mounted on moving equipment and system, that moving equipment
CN105843240B (en) * 2016-04-08 2019-01-22 北京航空航天大学 A kind of spacecraft attitude Integral Sliding Mode fault tolerant control method considering actuator failures
CN106406086B (en) * 2016-05-26 2019-05-07 北京航空航天大学 A kind of flexible spacecraft interference compensation method based on sliding formwork interference observer
CN109564406B (en) * 2016-08-03 2021-06-15 孟强 Self-adaptive terminal sliding mode control method
CN106774373B (en) * 2017-01-12 2019-11-05 哈尔滨工业大学 A kind of quadrotor drone finite time Attitude tracking control method
CN106886149B (en) * 2017-02-23 2019-10-22 哈尔滨工业大学 A kind of spacecraft robust finite time saturation Attitude tracking control method
CN107450584B (en) * 2017-08-29 2020-06-30 浙江工业大学 Aircraft self-adaptive attitude control method based on fixed time sliding mode
CN107703952B (en) * 2017-08-29 2020-10-30 浙江工业大学 Nonsingular fixed time self-adaptive attitude control method for rigid aircraft

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110244747A (en) * 2019-08-02 2019-09-17 大连海事大学 Heterogeneous fleet fault-tolerant control method based on actuator fault and saturation
CN112180960A (en) * 2020-09-29 2021-01-05 西北工业大学 Unmanned aerial vehicle fault-tolerant flight method and flight system for actuator faults
CN112180960B (en) * 2020-09-29 2021-09-14 西北工业大学 Unmanned aerial vehicle fault-tolerant flight method and flight system for actuator faults

Also Published As

Publication number Publication date
CN110515389A (en) 2019-11-29
CN110515389B (en) 2022-06-17

Similar Documents

Publication Publication Date Title
CN109143866A (en) A kind of adaptive set time Attitude tracking control method of rigid aircraft considering actuator constraints problem
CN107577144B (en) A kind of aircraft finite time Adaptive Attitude control method based on enhanced index Reaching Law
CN110471438A (en) A kind of set time Adaptive Attitude Tracking control method of rigid aircraft
CN106774273B (en) For the algorithm based on sliding mode prediction fault tolerant control method of time_varying delay control system actuator failures
Mao et al. Design and implementation of continuous finite-time sliding mode control for 2-DOF inertially stabilized platform subject to multiple disturbances
CN103728882B (en) The self-adaptation inverting non-singular terminal sliding-mode control of gyroscope
CN107908114B (en) Robust nonlinear control method and robust controller system for aircraft
He et al. Robust tracking control of helicopters using backstepping with disturbance observers
CN107703952A (en) A kind of nonsingular set time Adaptive Attitude control method of rigid aircraft
CN109164823A (en) A kind of nonsingular set time Attitude tracking control method of rigid-body spacecraft considering actuator constraints problem
CN109240321A (en) A kind of nonsingular set time neural network control method of rigid aircraft
CN109143846A (en) A kind of rigid aircraft adaptive neural network tracking and controlling method considering actuator constraints problem
CN109062240A (en) A kind of rigid aircraft set time Adaptive Attitude Tracking control method based on neural network estimation
CN105676852B (en) Small-sized depopulated helicopter is unpowered to learn model structure Adaptive Attitude control method
CN109188910A (en) A kind of fault-tolerant tracking and controlling method of the adaptive neural network of rigid aircraft
CN109375639A (en) A kind of rigid aircraft posture restraint tracking and controlling method based on asymmetric modified obstacle liapunov function
CN113110048B (en) Nonlinear system output feedback adaptive control system and method adopting HOSM observer
CN110515389A (en) A kind of calm method of the adaptive set time posture of rigid aircraft considering actuator constraints problem
CN109164816B (en) Non-cooperative target attitude tracking control method for controller fault and saturation characteristics
Bu et al. Robust tracking control of hypersonic flight vehicles: a continuous model-free control approach
CN110488854A (en) A kind of rigid aircraft set time Attitude tracking control method based on neural network estimation
CN108958043A (en) A kind of set time Adaptive Attitude fault tolerant control method of rigid aircraft
CN110488855A (en) A kind of adaptive set time posture fault tolerant control method of rigid aircraft based on neural network estimation
Serhal et al. Application of the extended high gain observer to underactuated mechanical systems
Han et al. Time delay control with sliding mode observer for a class of nonlinear systems: Performance and stability

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: 20190104

WW01 Invention patent application withdrawn after publication