CN108227728A - A kind of spacecraft attitude dynamic control allocation method for considering the switching of mixing executing agency - Google Patents
A kind of spacecraft attitude dynamic control allocation method for considering the switching of mixing executing agency Download PDFInfo
- Publication number
- CN108227728A CN108227728A CN201810269654.2A CN201810269654A CN108227728A CN 108227728 A CN108227728 A CN 108227728A CN 201810269654 A CN201810269654 A CN 201810269654A CN 108227728 A CN108227728 A CN 108227728A
- Authority
- CN
- China
- Prior art keywords
- control
- sgcmg
- spacecraft
- executing agency
- matrix
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000000034 method Methods 0.000 title claims abstract description 51
- 238000005457 optimization Methods 0.000 claims abstract description 29
- 239000011159 matrix material Substances 0.000 claims description 49
- 238000004422 calculation algorithm Methods 0.000 claims description 17
- 230000008859 change Effects 0.000 claims description 17
- 238000004364 calculation method Methods 0.000 claims description 13
- 238000009434 installation Methods 0.000 claims description 7
- 230000014509 gene expression Effects 0.000 claims description 5
- 230000015572 biosynthetic process Effects 0.000 claims description 3
- 238000005516 engineering process Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 2
- 230000007246 mechanism Effects 0.000 description 2
- 239000000203 mixture Substances 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000006978 adaptation Effects 0.000 description 1
- 230000008901 benefit Effects 0.000 description 1
- 230000007812 deficiency Effects 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
- 230000006641 stabilisation Effects 0.000 description 1
- 238000011105 stabilization Methods 0.000 description 1
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
- 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
Landscapes
- Physics & Mathematics (AREA)
- Engineering & Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Pure & Applied Mathematics (AREA)
- Mathematical Optimization (AREA)
- Mathematical Physics (AREA)
- Mathematical Analysis (AREA)
- Algebra (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
A kind of spacecraft attitude dynamic control allocation method for considering the switching of mixing executing agency, includes the following steps：First, it establishes using the spacecraft attitude control system model of mixing executing agency being made of single-gimbal control momentum gyro and counteraction flyback；Then, the control law of attitude control system is designed, ensures the stability of system；Later, the optimization object function of design control distribution method, and design the handoff parameter that can realize executing agency's switching；Finally, method based on method of Lagrange multipliers and liapunov function devises dynamic control allocation method, this method can realize the steady switching of mixing executing agency, Stability and veracity when rapidity and posture when improving Spacecraft During Attitude Maneuver orient, the computational efficiency of raising that can be larger.
Description
Technical field
The present invention relates to technical field of spacecraft attitude control, are mainly used in spacecraft attitude and become from fast reserve campaign
To realize the steady switching of executing agency, and it can be considered that executing agency's constraint and single frame control force in posture orientation process
Square gyro singular problem, more particularly to a kind of spacecraft attitude dynamic control allocation method for considering the switching of mixing executing agency.
Background technology
With the development of science and technology, task of the spacecraft in space is also more and more diversified.Just whether executing agency
Often work will directly determine Spacecraft Attitude Control can normal operation, so executing agency usually carries out redundant configuration.At this moment
With regard to needing rationally effectively to be assigned to the control instruction of controller design in the executing agency of redundancy using control allocation algorithm.
According to whether current control distribution method can be divided by unoptimizable method and optimization using the thought of optimization.Optimization utilizes
Control assignment problem is converted into an object function to be optimized and is solved by mathematical knowledge.
In-orbit spacecraft is expected to realize the fast reserve of posture to cope with a variety of different execution tasks.It can adopt
Ensure to realize that the fast reserve of posture can ensure the accurate orientation of posture again with mixing executing agency, and mix and hold
Row mechanism can be made of counteraction flyback (RW) and single-gimbal control momentum gyro (SGCMG).There is larger matter for one
For the spacecraft of amount, if carrying out the fast reserve of posture, executing agency is needed to provide larger control moment.SGCMG
For RW, larger control moment can be provided, but SGCMG usually there is relatively low torque resolution ratio can cause to navigate
Accuracy when its device posture orients reduces；Counteraction flyback, which can provide, is accurately controlled torque, but usually reaction flies
It is smaller to take turns the torque being capable of providing, it is difficult to meet the requirement of spacecraft fast reserve.
Invention content
It is an object of the invention to overcome the deficiencies of the prior art and provide a kind of space flight for considering the switching of mixing executing agency
Device posture dynamic control allocation method, since executing agency in Spacecraft Attitude Control task, both having been needed to provide larger power
Square carries out spacecraft attitude fast reserve, and executing agency is needed to provide smaller but accurate control moment and realizes posture orientation,
The mixing executing agency formed using SGCMG and RW, and devise the optimization object function of control allocation algorithm, optimization aim letter
Devised in number realize executing agency switching handoff parameter so that spacecraft need carry out posture fast reserve when mainly by
SGCMG provides control moment, and mainly control moment is provided by RW when needing to carry out posture orientation.In addition, optimization object function
In also contemplate executing agency constraint and SGCMG singular problem.Finally, by using method of Lagrange multipliers and Li Yapu
Control allocation algorithm based on optimization is converted into dynamic control allocation algorithm, can effectively reduced by the method for promise husband's function
The calculation amount of system.
The present invention provide it is a kind of consider mixing executing agency switching spacecraft attitude dynamic control allocation method, including with
Lower step：
(1) it is established based on attitude motion of spacecraft with kinetic model using the spacecraft attitude control for mixing executing agency
System model processed；
(2) the spacecraft attitude control system model established based on step (1), designs the control of spacecraft attitude control system
System rule ensures the stability of entire spacecraft attitude control system；
(3) consider executing agency's constraint and the single frame of counteraction flyback RW and single-gimbal control momentum gyro SGCMG
The singular problem of control-moment gyro SGCMG, the optimization object function of design control allocation algorithm, and design can realize execution
The handoff parameter of mechanism switching；
(4) optimization object function established based on step (3), according to method of Lagrange multipliers and liapunov function
Method has obtained dynamic control allocation method.
Wherein, the attitude motion of spacecraft model established in step (1) is as follows：
Four element [q of posture0,qv T]T∈R4Represent posture of the spacecraft ontology relative to inertial coodinate system I；qv=[q1 q2
q3]TIt is the vector portion of four element of posture, andex、ey、ezRepresent boat
Components of the unit vector e in three reference axis of inertial coodinate system on Euler's axis of its device rotation, φ represent spacecraft around Europe
The angle that pulling shaft turns over；q0It is the scalar component of four element of posture, andVector portion qvWith scalar component q0It is full
Sufficient equationω=[ω1 ω2 ω3]T∈R3It is angular speed of the spacecraft ontology relative to inertial coodinate system I,
ω1、ω2、ω3Respectively roll angular speed, yaw rate and the rate of pitch of satellite；Wherein qv ×Represent it is a kind of about
Two vectorial vector product calculations are converted into matrix and vector by the skew symmetric matrix of four element vector portion of spacecraft attitude
Product calculation, with following form：
The kinetic model of spacecraft attitude is as follows：
Wherein, J represents the moment of inertia matrix of spacecraft, and is 3 × 3 positive definite symmetric matrices；h∈R3It is spacecraft
Total angular momentum, and represent on three axis of body coordinate system B, total angular momentum is by spacecraft ontology angular momentum Jω, control moment
Gyro angular momentum AwIwΩ and flywheel angular momentum ArwIrwΩrwThree parts form, so h=J ω+AwIwΩ+ArwIrwΩrw；Aw∈
R3×mAnd At∈R3×mIt is the installation matrix of SGCMG, m is the number of SGCMG, AwColumn vector be by m SGCMG frame axis direction
On unit vector form, AtColumn vector be to be made of the unit vector on m SGCMG frame lateral shaft, Arw∈R3×nIt is RW
Installation matrix, n is the number of RW；Matrix A is installedwAnd AtValue depend on angle γ=[γ for turning over of SGCMG frames1,
γ2…γm]T, wherein γi, i=1,2 ..., m represent the angle that the frame of i-th of SGCMG turns over, and Aw=Aw0[Cos
γ]d+At0[Sinγ]d, At=At0[Cosγ]d-Aw0[Sinγ]d, Aw0And At0Represent SGCMG installation matrix γ=[0,
0,…,0]T∈RmWhen value；[Cosγ]dIt represents using the element of vectorial Cos γ as the diagonal matrix of diagonal entry, [Sin
γ]dIt represents using the element of vectorial Sin γ as the diagonal matrix of diagonal entry；Wherein Sin (γ)=[sin (γ1),sin
(γ2),…,sin(γm)]T, Cos (γ)=[cos (γ1),cos(γ2),…,cos(γm)]T；
Iw∈Rm×mIt is diagonal matrix, diagonal entry is made of the rotary inertia of m identical SGCMG, Irw∈Rn×n
It is also diagonal matrix, diagonal entry is that the rotary inertia of RW communicated by n is formed；Ω=[Ω1,Ω2...Ωm]TIt represents
The angular speed of SGCMG rotors, Ωi, i=1,2 ..., m, the angular speed of i-th of SGCMG rotor of expression；Ωrw∈RnRepresent flywheel
The angular speed of rotor, Ωrw,i, i=1,2 ..., m, the angular speed of i-th of flywheel rotor of expression；ΩdIt represents with the element of vectorial Ω
Diagonal matrix for diagonal entry；It represents the change rate of γ, is the angle of rotation speed of SGCMG frames
Degree, Then represent the rotational angular velocity of the frame of i-th of SGCMG；urw∈RnIt is the control force of RW outputs
Square, urw,i, i=1,2 ..., n, then it represents that the control moment of i-th of RW output；ω×It is about a kind of oblique of spacecraft angular speed
Symmetrical matrix, can be converted into two vectorial vector product calculations the product calculation of matrix and vector, and form is as follows：
Wherein, in step (2), the control law of attitude control system is as follows：
τdesignIt is the torque of design of control law, wherein μ, ρ, σ, γkBe control law parameter and be greater than zero constant；K is
The control law parameter of one variation, change rate areK is taken in the value of zero momentThen it will be apparent that
And when k level off to zero when,Also level off to zero, thus k be one successively decrease but consistently greater than zero number；Tanh () is hyperbolic
Tangent function, for some n-dimensional vector x=[x1,x2,…,xn]T, Tanh (x)=[tanh (x1),tanh(x2),...,tanh
(xn)]T。
Wherein, the optimization object function of the control allocation algorithm based on optimization is as follows in step (3)：
J (u, γ)=J1+J2+J3
Wherein：
Defined in itIt is the output for needing the control acquired distribution and the control for passing to executing agency
Signal processed；η is for realizing the handoff parameter switched between SGCMG and RW Liang Zhong executing agencies；Because the control signal of SGCMG
It is it is expected rotating speed, and the control signal of RW is it is expected torque, so with a by the control signal of Liang Zhong executing agenciesAnd urw's
Unit unitizes, and a=(AtIw)2；ImIt is the unit matrix of a m rank, InIt is the unit matrix of a n rank；ucmg,maxWith-
ucmg,maxRepresent the upper and lower bound of SGCMG frame rotating speeds；urw,maxWith-urw,maxRepresent the upper and lower bound of RW output torques；
Det () represents to take the determinant of square formation；Optimization object function J is made of three parts, J1Represent executing agency's consumption energy
Estimation；J2For the inequality constraints of executing agency SGCMG and RW are considered in optimization object function,WithIt is greater than
Zero constant, whenOr urw,iLevel off to corresponding bound when, J2It will level off to infinity, minimize to object function J,
The obtained output of control distributionExecuting agency's constraint will be met；J3For considering the strange of SGCMG mounting configurations
The opposite sex,Be one be more than zero constant, ε " be one be more than zero and smaller constant, when SGCMG levels off to it is unusual when,
det(At) it will level off to zero, andTo level off to it is just infinite, if ε "=0, J3It will level off to positive nothing
Thoroughly, it minimizes to J, J3To be a limited value, so the unusual state of SGCMG will be avoided by；
Executing agency's handoff parameter is：
Wherein K, P are values of the η respectively when spacecraft attitude fast reserve and posture orient, be greater than zero it is normal
Number, usual K are much larger than P；R represents steepness of the η in change procedure, is greater than zero constant；θ represents current spacecraft attitude
State, and be defined asWherein e represents attitude error, and is defined asα1
And α2It is positive constant.
Wherein, dynamic control allocation algorithm is as follows in step (4)：
In step (3), the optimization object function J (u, γ) of control allocation algorithm is established, optimum control distribution is calculated
There are equality constraint τ for methoddesign=Q (γ) u, wherein Q (γ)=[- AtIwΩd,Arw],I.e. spacecraft is practical
Control moment Q (γ) u being subject to and design of control law torque τdesignIt is equal, it can obtain l (γ, u, λ) using method of Lagrange multipliers
=J (γ, u)+(τdesign-Q(γ)u)Tλ, wherein λ are Lagrange multipliers；It is the change rate of control distribution output u,It is to draw
The change rate of Ge Lang multipliers λ；Γ∈R7×7With W ∈ R3×3It is symmetrical and positive definite constant matrices；α and β can be calculated
Go out, andζ and φ meets formula αTζ+βTφ+δ=0, whereinSolve ζ and φ.
The spacecraft attitude control system of the present invention includes single frame including controller, dynamic control allocation device, executing agency
Frame control moment gyro and counteraction flyback, spacecraft kinematics model and spacecraft dynamics model etc., controller according to
The posture and angular speed output design control moment, dynamic control allocation algorithm of current spacecraft divide the control signal of controller
It is fitted in each executing agency, and the handoff parameter calculated according to current system conditions, to determine that master plays in which kind of executing agency
It acts on, the control moment of executing agency's output is applied on spacecraft attitude dynamics equation, and then kinetics equation is by angle
Speed is output in kinematical equation, and kinematical equation exports the attitude quaternion of spacecraft, attitude quaternion and spacecraft angle
Speed together forms control law, and advantage is compared with prior art：
(1) present invention can realize the steady switching of mixing executing agency, need to carry out posture fast reserve in spacecraft
When, torque mainly is provided by single-gimbal control momentum gyro, when spacecraft needs to carry out posture orientation, then mainly by reaction
Flywheel provides torque, does so accuracy and stabilization when rapidity and posture when can improve Spacecraft During Attitude Maneuver orient
Property；
(2) traditional control allocation algorithm based on optimization is passed through method of Lagrange multipliers and Li Yapunuo by the present invention
The method of husband's function is converted into the newer dynamic control allocation method of dynamic, the computational efficiency of raising that can be larger.
Description of the drawings
Fig. 1 is the spacecraft attitude dynamic control allocation Method And Principle block diagram for mixing executing agency's switching；
Fig. 2 is the spacecraft attitude dynamic control allocation method flow block diagram for considering the switching of mixing executing agency.
Specific embodiment
The following detailed description of the specific implementation of the present invention, it is necessary to it is indicated herein to be, implement to be only intended to this hair below
Bright further explanation, it is impossible to be interpreted as limiting the scope of the invention, field technology skilled person is according to above-mentioned
Some nonessential modifications and adaptations that invention content makes the present invention, still fall within protection scope of the present invention.
The present invention provides a kind of spacecraft attitude dynamic control allocation methods for considering the switching of mixing executing agency, such as scheme
Shown in 2, the specific steps are：
First, the spacecraft attitude control system model using mixing executing agency is established；Then attitude control law is designed,
Ensure the stability of whole system；Later, the optimization object function of control distribution method is established, and energy is designed in object function
Enough handoff parameters for realizing the switching of mixing executing agency；Finally, using suitable parameter, based on liapunov function method and
Method of Lagrange multipliers designs dynamic control allocation method；Specific implementation step is as follows：
The first step establishes spacecraft attitude control system model：
Four element [q of posture0,qv T]T∈R4Represent posture of the spacecraft ontology relative to inertial coodinate system I；qv=[q1 q2
q3]TIt is the vector portion of four element of posture, andex、ey、ezRepresent boat
Components of the unit vector e in three reference axis of inertial coodinate system on Euler's axis of its device rotation, φ represent spacecraft around Europe
The angle that pulling shaft turns over；q0It is the scalar component of four element of posture, andVector portion qvWith scalar component q0It is full
Sufficient equationHere the initial value for taking four element of posture is [0.9-0.3 0.26 0.18]T；ω=[ω1 ω2
ω3]T∈R3It is angular speed of the spacecraft ontology relative to inertial coodinate system I, ω1、ω2、ω3The respectively roll angle speed of satellite
Degree, yaw rate and rate of pitch, the initial value for taking angular speed here are ω (0)=[0 0 0]Trad/s；WhereinTable
Show a kind of skew symmetric matrix about four element vector portion of space attitude, two vectorial vector product calculations can be converted into
The product calculation of matrix and vector has following form：
The kinetic model of spacecraft attitude is as follows：
J represents the moment of inertia matrix of spacecraft, and is 3 × 3 symmetric positive definite matrix, the spacecraft in engineering
Design parameter, J can be chosen for J=[135 56；5 126 7；6 7 131]kg·m2；h∈R3It is the total angular momentum of spacecraft,
And represent on tri- axis of body coordinate system B, total angular momentum is by spacecraft ontology angular momentum J ω, control-moment gyro angular momentum AwIw
Ω and flywheel angular momentum ArwIrwΩrwThree parts form, so h=J ω+AwIwΩ+ArwIrwΩrw；It takes by 4 SGCMG and 3
The mixing executing agency of RW compositions；Aw∈R3×4And At∈R3×4It is the installation matrix of SGCMG, 4 be the number of SGCMG, AwRow
Vector is made of the unit vector in 4 SGCMG frame axis directions, AtColumn vector by the list on 4 SGCMG frame lateral shafts
Bit vector is formed, Arw∈R3×3It is the installation matrix of RW；Matrix A is installedwAnd AtValue depending on SGCMG frame around gimbal axis
The angle γ turned over=[γ1,γ2,γ3,γ4]T, wherein γi, i=1,2,3,4, represent that the frame of i-th of SGCMG turns over
Angle takes initial time γ here0=[0,0,0,0]T；Aw=Aw0[Cosγ]d+At0[Sinγ]d, At=At0[Cosγ]d-Aw0
[Sinγ]d, Aw0And At0Represent that SGCMG installs matrix in γ=[0,0,0,0]TWhen value；[Cosγ]dIt represents with vector
Diagonal matrix of the element of Cos γ for diagonal entry, [Sin γ]dIt represents using the element of vectorial Sin γ as diagonal entry
Diagonal matrix；Wherein Sin (γ)=[sin (γ1),sin(γ2),sin(γ3),sin(γ4)]T, Cos (γ)=[cos
(γ1),cos(γ2),cos(γ3),cos(γ4)]T；Here it takes
Iw∈Rm×mIt is diagonal matrix, diagonal entry is made of the rotary inertia of SGCMG, takes each SGCMG's here
Rotary inertia is all equal and is 0.05kgm2；Ω=[Ω1,Ω2,Ω3,Ω4]TRepresent the angular speed of SGCMG rotors, Ωi, i
=1,2,3,4, it represents the angular speed of i-th of SGCMG rotor, takes Ω herei=6000rpm, i=1,2,3,4；Single flywheel
Angular momentum can be acquired by the integration of the control moment of its output, and the initial value for taking flywheel angular momentum here is zero；ΩdRepresent with to
The element for measuring Ω is the diagonal matrix of diagonal entry；It represents the change rate of γ, is SGCMG frames
Rotational angular velocity, Then represent the rotational angular velocity of the frame of i-th of SGCMG；urw∈RnIt is the control of RW outputs
Torque processed, urw,i, i=1,2,3, then it represents that the control moment of i-th of RW output；ω×It is one kind about spacecraft angular speed
Skew symmetric matrix, can be converted into two vectorial vector product calculations the product calculation of matrix and vector, and form is as follows：
Second step designs the control law of attitude motion of spacecraft：
τdesignIt is design of control law torque, wherein μ, ρ, σ, γkAll it is control law parameter and for more than zero constant, this
In take μ=0.8, ρ=6.4, σ=0.001, γk=0.01；K is the control law parameter of a variation, and change rate isHere
K is taken in the value k of zero moment0=6.2, then it will be apparent thatAnd when k level off to zero when,Also zero is leveled off to, so k is
One successively decrease but consistently greater than zero number；Tanh () is hyperbolic tangent function, for some n-dimensional vector x=[x1,x2,…,
xn]T, Tanh (x)=[tanh (x1),tanh(x2),...,tanh(xn)]T。
Third walks, and designs the optimization object function of the control allocation algorithm based on optimization：
J (u, γ)=J1+J2+J3
Wherein：
Defined in itIt is the output for needing the control acquired distribution and the control for passing to executing agency
Signal processed；η is for realizing the handoff parameter switched between SGCMG and RW Liang Zhong executing agencies；Because the control signal of SGCMG
It is it is expected rotating speed, and the control signal of RW is it is expected torque, so with a by the control signal of Liang Zhong executing agenciesAnd urw's
Unit unitizes, and a=(AtIw)2；ImIt is the unit matrix of a m rank, InIt is the unit matrix of a n rank；ucmg,maxWith-
ucmg,maxIt represents the upper and lower bound of SGCMG frame rotating speeds, takes u herecmg,max=0.6rad/s；urw,maxWith-urw,maxRepresent RW
The upper and lower bound of output torque, takes u hererw,max=0.2Nm；Det () expressions take determinant to square formation；Optimization aim letter
Number J is made of three parts, J1Represent the estimation of executing agency's consumption energy；J2For the inequality of executing agency SGCMG and RW
It constrains and considers in optimization object function,WithZero constant is greater than, is taken hereWhenOr urw,i
Level off to corresponding bound when, J2To level off to infinity, minimize to object function J, obtained control distribution it is defeated
Go outExecuting agency's constraint will be met；J3For considering the singularity of SGCMG mounting configurations,It is one to be more than
Zero constant, takes hereε " is one and is more than zero and smaller constant, takes ε "=10 here-3, work as SGCMG
When leveling off to unusual, det (At) it will level off to zero, andTo level off to it is just infinite, if ε "=0, J3
It will level off to just infinite, minimize to J, J3To be a limited value, so the unusual state of SGCMG will be avoided by；
Executing agency's handoff parameter is：
Wherein K, P are values of the η respectively when spacecraft attitude fast reserve and posture orient, be greater than zero it is normal
Number, usual K are much larger than P, take K=1000, P=0.0022 here；R represents steepness of the η in change procedure, be greater than zero it is normal
Number, takes r=140 here；θ represents current spacecraft attitude state, and is defined asWherein e
It represents attitude error, and is defined asα1And α2It is positive constant, takes α here1=0.3, α2=0.5.
4th step, based on the optimization object function that second step is established, according to method of Lagrange multipliers and Liapunov letter
Several methods, the newer dynamic control allocation method of design dynamic：
In step (3), the optimization object function J (u, γ) of control allocation algorithm is established, optimum control distribution is calculated
There are equality constraint τ for methoddesign=Q (γ) u, wherein Q (γ)=[- AtIwΩd,Arw],I.e. spacecraft is practical
Control moment Q (γ) u and the torque τ of design of control law being subject todesignIt is equal, using method of Lagrange multipliers can obtain l (γ, u,
λ)=J (γ, u)+(τdesign-Q(γ)u)Tλ, wherein λ are Lagrange multipliers；It is the change rate of control distribution output u,
It is the change rate of Lagrange multiplier λ, wherein the initial value of u is taken as u (0)=[0.3825-0.2004-0.2420 0.3409
0.0001 -0.0001 0]T, the initial value of λ is taken as λ (0)=[0.3 0.3]T；0.Γ3∈R7×7With W ∈ R3×3Be it is symmetrical and
The constant matrices of positive definite takes Γ=10 × diag ([0.01 0.01 0.01 0.01 0.01 0.01 0.01]) here, wherein
Diag () is to construct diagonal matrix by diagonal entry of the element of some column vector or row vector, takes W=25 × diag
([0.01 0.01 0.01])；α and β can be calculated, andζ and φ
Meet formula αTζ+βTφ+δ=0, whereinBy constructing linear equation as follows,
Can be in the hope of the value of ζ and φ, wherein υ is a variable to help out when constructing linear equation, takes less than its value：
By system above associative simulation, it can realize the different conditions according to spacecraft attitude between SGCMG and RW
It switches over, can ensure accuracy and stability when rapidity and posture of the system in attitude maneuver orient, and energy
Enough consider executing agency's constraint and the singular problem of SGCMG；In addition, dynamic control allocation method is than control that direct solution optimizes
Distribution method operation processed is faster.
Although for illustrative purposes, it has been described that exemplary embodiments of the present invention, those skilled in the art
Member it will be understood that, can be in form and details in the case of the scope and spirit for not departing from invention disclosed in appended claims
On the change that carry out various modifications, add and replace etc., and all these changes should all belong to appended claims of the present invention
Protection domain, and each step in each department of claimed product and method, can be in any combination
Form is combined.Therefore, to disclosed in this invention the description of embodiment be not intended to limit the scope of the invention,
But for describing the present invention.Correspondingly, the scope of the present invention is not limited by embodiment of above, but by claim or
Its equivalent is defined.
Claims (5)
1. it is a kind of consider mixing executing agency switching spacecraft attitude dynamic control allocation method, which is characterized in that including with
Lower step：
(1) it is established based on attitude motion of spacecraft with kinetic model using the Spacecraft Attitude Control system for mixing executing agency
System model；
(2) the spacecraft attitude control system model established based on step (1), designs the control of spacecraft attitude control system
Rule ensures the stability of entire spacecraft attitude control system；
(3) consider executing agency's constraint of counteraction flyback RW and single-gimbal control momentum gyro SGCMG and single frame control
The singular problem of moment gyro SGCMG, the optimization object function of design control allocation algorithm, and design can realize executing agency
The handoff parameter of switching；
(4) optimization object function established based on step (3), according to the side of method of Lagrange multipliers and liapunov function
Method has obtained dynamic control allocation method.
2. according to the method described in claim 1, it is characterized in that：The attitude motion of spacecraft model established in step (1)
It is as follows：
Four element [q of posture0,qv T]T∈R4Represent posture of the spacecraft ontology relative to inertial coodinate system I；qv=[q1 q2 q3]T
It is the vector portion of four element of posture, and ex、ey、ezRepresent space flight
Components of the unit vector e in three reference axis of inertial coodinate system on Euler's axis of device rotation, φ represent spacecraft around Euler
The angle that shaft rotation is crossed, q0It is the scalar component of four element of posture, andVector portion qvWith scalar component q0Meet
Equationω=[ω1 ω2 ω3]T∈R3It is angular speed of the spacecraft ontology relative to inertial coodinate system I,
ω1, ω2, ω3Respectively roll angular speed, yaw rate and the rate of pitch of satellite；Wherein qv ×Represent it is a kind of about
Two vectorial vector product calculations are converted into matrix and vector by the skew symmetric matrix of four element vector portion of spacecraft attitude
Product calculation, with following form：
The kinetic model of spacecraft attitude is as follows：
Wherein, J represents the moment of inertia matrix of spacecraft, and is 3 × 3 positive definite symmetric matrices；h∈R3It is the total of spacecraft
Angular momentum, and represent on three axis of body coordinate system B, total angular momentum is by spacecraft ontology angular momentum J ω, control-moment gyro
Angular momentum AwIwΩ and flywheel angular momentum ArwIrwΩrwThree parts form, so h=J ω+AwIwΩ+ArwIrwΩrw；Aw∈R3×m
And At∈R3×mIt is the installation matrix of SGCMG, m is the number of SGCMG, AwColumn vector be by m SGCMG frame axis direction
Unit vector form, AtColumn vector be to be made of the unit vector on m SGCMG frame lateral shaft, Arw∈R3×nIt is RW
Matrix is installed, n is the number of RW；Matrix A is installedwAnd AtValue depend on angle γ=[γ for turning over of SGCMG frames1,γ2…
γm]T, wherein γi, i=1,2 ..., m represent the angle that the frame of i-th of SGCMG turns over, and Aw=Aw0[Cosγ]d+At0
[Sinγ]d, At=At0[Cosγ]d-Aw0[Sinγ]d, Aw0And At0Represent the installation matrix of SGCMG in γ=[0,0 ..., 0]T
∈RmWhen value；[Cosγ]dIt represents using the element of vectorial Cos γ as the diagonal matrix of diagonal entry, [Sin γ]dIt represents
Using the element of vectorial Sin γ as the diagonal matrix of diagonal entry；Wherein Sin (γ)=[sin (γ1),sin(γ2),…,sin
(γm)]T, Cos (γ)=[cos (γ1),cos(γ2),…,cos(γm)]T；
Iw∈Rm×mIt is diagonal matrix, diagonal entry is made of the rotary inertia of m identical SGCMG, Irw∈Rn×nIt is also
Diagonal matrix, diagonal entry are that the rotary inertia of RW communicated by n is formed；Ω=[Ω1,Ω2...Ωm]TRepresent SGCMG
The angular speed of rotor, Ωi, i=1,2 ..., m, the angular speed of i-th of SGCMG rotor of expression；Ωrw∈RnRepresent flywheel rotor
Angular speed, Ωrw,i, i=1,2 ..., m, the angular speed of i-th of flywheel rotor of expression；ΩdRepresent using the element of vectorial Ω to be diagonal
The diagonal matrix of line element；It represents the change rate of γ, is the rotational angular velocity of SGCMG frames,I=
1,2 ..., m, then it represents that the rotational angular velocity of the frame of i-th of SGCMG；urw∈RnBe RW output control moment, urw,i, i=
1,2 ..., n, then it represents that the control moment of i-th of RW output；A kind of skew symmetric matrix of ω × be about spacecraft angular speed,
Two vectorial vector product calculations can be converted into the product calculation of matrix and vector, form is as follows：
3. according to the method described in claim 1, it is characterized in that：In step (2), the control law of attitude control system is as follows：
τdesignIt is the torque of design of control law, wherein μ, ρ, σ, γkBe control law parameter and be greater than zero constant；K is one
The control law parameter of variation, change rate areK is taken in the value of zero momentThen it will be apparent thatAnd
When k level off to zero when,Also level off to zero, thus k be one successively decrease but consistently greater than zero number；Tanh () is tanh
Function, for some n-dimensional vector x=[x1,x2,…,xn]T, Tanh (x)=[tanh (x1),tanh(x2),...,tanh(xn)
]T。
4. according to the method described in claim 1, it is characterized in that：The control allocation algorithm based on optimization is excellent in step (3)
It is as follows to change object function：
J (u, γ)=J1+J2+J3
Wherein：
Defined in itIt is the output for needing the control acquired distribution and the control letter for passing to executing agency
Number；η is for realizing the handoff parameter switched between SGCMG and RW Liang Zhong executing agencies；Because the control signal of SGCMG is the phase
Hope rotating speed, and the control signal of RW is it is expected torque, so with a by the control signal of Liang Zhong executing agenciesAnd urwUnit
It is unitized, and a=(AtIw)2；ImIt is the unit matrix of a m rank, InIt is the unit matrix of a n rank；ucmg,maxWith-ucmg,maxTable
Show the upper and lower bound of SGCMG frame rotating speeds；urw,maxWith-urw,maxRepresent the upper and lower bound of RW output torques；Det () table
Show the determinant for taking square formation；Optimization object function J is made of three parts, J1Represent the estimation of executing agency's consumption energy；J2It is used for
The inequality constraints of executing agency SGCMG and RW are considered in optimization object function,WithZero constant is greater than,
WhenOr urw,iLevel off to corresponding bound when, J2It will level off to infinity, minimize to object function J, obtained control
The output of distributionExecuting agency's constraint will be met；J3For considering the singularity of SGCMG mounting configurations,It is
One be more than zero constant, ε " be one be more than zero and smaller constant, when SGCMG levels off to it is unusual when, det (At) will
Level off to zero, andTo level off to it is just infinite, if ε "=0, J3It will level off to just infinite, J will be asked most
Small value, J3To be a limited value, so the unusual state of SGCMG will be avoided by；
Executing agency's handoff parameter is：
Wherein K, P are values of the η respectively when spacecraft attitude fast reserve and posture orient, and are greater than zero constant, lead to
Normal K is much larger than P；R represents steepness of the η in change procedure, is greater than zero constant；θ represents the state of current spacecraft attitude,
And it is defined asWherein e represents attitude error, and is defined asα1And α2
It is positive constant.
5. according to the method described in claim 1, it is characterized in that：Dynamic control allocation algorithm is as follows in step (4)：
In step (3), the optimization object function J (u, γ) of control allocation algorithm, control allocation algorithm presence etc. are established
Formula constrains τdesign=Q (γ) u, wherein Q (γ)=[- AtIwΩd,Arw],That is the practical control being subject to of spacecraft
Torque Q (γ) u processed and design of control law torque τdesignIt is equal, using method of Lagrange multipliers can obtain l (γ, u, λ)=J (γ,
u)+(τdesign-Q(γ)u)Tλ, wherein λ are Lagrange multipliers；It is the change rate of control distribution output u,It is Lagrange
The change rate of multiplier λ；Γ∈R7×7With W ∈ R3×3It is symmetrical and positive definite constant matrices；α and β can be calculated, andζ and φ meets formula αTζ+βTφ+δ=0, wherein
Solve ζ and φ.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810269654.2A CN108227728B (en) | 2018-03-29 | 2018-03-29 | Spacecraft attitude dynamic control distribution method considering hybrid actuating mechanism switching |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201810269654.2A CN108227728B (en) | 2018-03-29 | 2018-03-29 | Spacecraft attitude dynamic control distribution method considering hybrid actuating mechanism switching |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN108227728A true CN108227728A (en) | 2018-06-29 |
| CN108227728B CN108227728B (en) | 2020-01-10 |
Family
ID=62659228
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201810269654.2A Active CN108227728B (en) | 2018-03-29 | 2018-03-29 | Spacecraft attitude dynamic control distribution method considering hybrid actuating mechanism switching |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN108227728B (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109116730A (en) * | 2018-07-11 | 2019-01-01 | 南京航空航天大学 | A kind of energy-optimised management method of mixing executing agency based on TU cooperative game |
| CN109164822A (en) * | 2018-09-26 | 2019-01-08 | 北京航空航天大学 | It is a kind of based on have mixing executing agency Spacecraft Attitude Control method |
| CN109668710A (en) * | 2018-12-18 | 2019-04-23 | 大连理工大学 | Strut shoe formula dummy vehicle multi-dimensional vibration control method |
| CN109823572A (en) * | 2019-01-30 | 2019-05-31 | 武汉大学 | Actuating mechanism configuration and control method for reciprocating and rapid swinging of agile satellite attitude |
| CN112357120A (en) * | 2020-10-22 | 2021-02-12 | 北京航空航天大学 | Reinforced learning attitude constraint control method considering installation deviation of actuating mechanism |
| CN113156987A (en) * | 2021-04-15 | 2021-07-23 | 哈尔滨工业大学 | Spacecraft actuating mechanism combining double-frame shear type moment gyroscope and flywheel and control method thereof |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101934863A (en) * | 2010-09-29 | 2011-01-05 | 哈尔滨工业大学 | Satellite posture all-round controlling method based on magnetic moment device and flywheel |
| CN102627151A (en) * | 2012-05-09 | 2012-08-08 | 哈尔滨工业大学 | Moment distribution method for rapid maneuvering satellite based on mixed actuating mechanism |
| CN103092208A (en) * | 2013-01-09 | 2013-05-08 | 哈尔滨工业大学 | Spacecraft high-accuracy speediness attitude maneuver method based on single gimbal control moment gyro (SGCMG) and reaction wheel (RW) |
| CN105867401A (en) * | 2016-04-28 | 2016-08-17 | 北京航空航天大学 | Spacecraft posture fault tolerance control method of single gimbal control moment gyroscope groups |
| CN106218922A (en) * | 2016-07-27 | 2016-12-14 | 中国科学院长春光学精密机械与物理研究所 | The associating actuating mechanism controls method of flexible quick satellite |
-
2018
- 2018-03-29 CN CN201810269654.2A patent/CN108227728B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101934863A (en) * | 2010-09-29 | 2011-01-05 | 哈尔滨工业大学 | Satellite posture all-round controlling method based on magnetic moment device and flywheel |
| CN102627151A (en) * | 2012-05-09 | 2012-08-08 | 哈尔滨工业大学 | Moment distribution method for rapid maneuvering satellite based on mixed actuating mechanism |
| CN103092208A (en) * | 2013-01-09 | 2013-05-08 | 哈尔滨工业大学 | Spacecraft high-accuracy speediness attitude maneuver method based on single gimbal control moment gyro (SGCMG) and reaction wheel (RW) |
| CN105867401A (en) * | 2016-04-28 | 2016-08-17 | 北京航空航天大学 | Spacecraft posture fault tolerance control method of single gimbal control moment gyroscope groups |
| CN106218922A (en) * | 2016-07-27 | 2016-12-14 | 中国科学院长春光学精密机械与物理研究所 | The associating actuating mechanism controls method of flexible quick satellite |
Non-Patent Citations (4)
| Title |
|---|
| BOYAN JIANG,等: "Fixed-Time Attitude Control for Rigid Spacecraft With Actuator Saturation and Faults", 《IEEE TRANSACTIONS ON CONTROL SYSTEMS TECHNOLOGY》 * |
| YAJIE MA，等: "Actuator failure compensation and attitude control for rigid satellite by adaptive control using quaternion feedback", 《JOURNAL OF THE FRANKLIN INSTITUTE》 * |
| 叶东，等: "考虑安装偏差的联合执行机构自适应控制算法", 《哈尔滨工业大学学报》 * |
| 金磊，等: "采用单框架控制力矩陀螺和动量轮的航天器姿态跟踪控制研究", 《宇航学报》 * |
Cited By (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109116730A (en) * | 2018-07-11 | 2019-01-01 | 南京航空航天大学 | A kind of energy-optimised management method of mixing executing agency based on TU cooperative game |
| CN109116730B (en) * | 2018-07-11 | 2021-10-12 | 南京航空航天大学 | Hybrid execution mechanism energy optimization management method based on TU cooperative game |
| CN109164822A (en) * | 2018-09-26 | 2019-01-08 | 北京航空航天大学 | It is a kind of based on have mixing executing agency Spacecraft Attitude Control method |
| CN109164822B (en) * | 2018-09-26 | 2020-02-21 | 北京航空航天大学 | Spacecraft attitude control method based on hybrid actuating mechanism |
| CN109668710A (en) * | 2018-12-18 | 2019-04-23 | 大连理工大学 | Strut shoe formula dummy vehicle multi-dimensional vibration control method |
| CN109668710B (en) * | 2018-12-18 | 2020-04-07 | 大连理工大学 | Multi-dimensional vibration control method for strut tail support type aircraft model |
| US10969298B2 (en) | 2018-12-18 | 2021-04-06 | Dalian University Of Technology | Multi-dimensional vibration control method for the model of strut tail-supported aircraft |
| CN109823572A (en) * | 2019-01-30 | 2019-05-31 | 武汉大学 | Actuating mechanism configuration and control method for reciprocating and rapid swinging of agile satellite attitude |
| CN112357120A (en) * | 2020-10-22 | 2021-02-12 | 北京航空航天大学 | Reinforced learning attitude constraint control method considering installation deviation of actuating mechanism |
| CN112357120B (en) * | 2020-10-22 | 2022-05-20 | 北京航空航天大学 | Reinforced learning attitude constraint control method considering installation deviation of actuating mechanism |
| CN113156987A (en) * | 2021-04-15 | 2021-07-23 | 哈尔滨工业大学 | Spacecraft actuating mechanism combining double-frame shear type moment gyroscope and flywheel and control method thereof |
| CN113156987B (en) * | 2021-04-15 | 2022-05-31 | 哈尔滨工业大学 | Spacecraft actuating mechanism combining double-frame shear type moment gyroscope and flywheel and control method thereof |
Also Published As
| Publication number | Publication date |
|---|---|
| CN108227728B (en) | 2020-01-10 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN108227728A (en) | A kind of spacecraft attitude dynamic control allocation method for considering the switching of mixing executing agency | |
| CN105867401B (en) | The spacecraft attitude fault tolerant control method of single-gimbal control moment gyros | |
| CN106933241B (en) | Single-gimbal control momentum gyro spacecraft fault tolerant control method based on fault de couple | |
| CN106292681B (en) | A kind of satellite Active Fault-tolerant Control Method distributed based on observer and On-line Control | |
| US8131409B2 (en) | Gyroless transfer orbit sun acquisition using only wing current measurement feedback | |
| Hu et al. | Robust finite-time control allocation in spacecraft attitude stabilization under actuator misalignment | |
| CN107168357B (en) | A kind of spacecraft attitude maneuver control method considering posture restraint and anti-unwinding | |
| CN104960674B (en) | A kind of sensing tracking and controlling method of moving target | |
| CN107688295A (en) | A kind of quadrotor finite time self-adaptation control method based on fast terminal sliding formwork | |
| CN102163057B (en) | Tracking control method of plane paths of self-managing airship | |
| Hu et al. | Nonlinear proportional-derivative control incorporating closed-loop control allocation for spacecraft | |
| CN103092208A (en) | Spacecraft high-accuracy speediness attitude maneuver method based on single gimbal control moment gyro (SGCMG) and reaction wheel (RW) | |
| CN105159083A (en) | High-precision friction compensation control method of double-frame magnetic levitation CMG frame system | |
| CN103034237B (en) | Spacecraft attitude maneuver control method using two single-frame control moment gyroscopes | |
| Hu et al. | Velocity-free fault-tolerant control allocation for flexible spacecraft with redundant thrusters | |
| CN112572835B (en) | Satellite in-orbit angular momentum management and control method with attitude switching function | |
| CN110333733A (en) | A kind of the tandem variable universe fuzzy attitude control system and method for quadrotor | |
| US10202208B1 (en) | High control authority variable speed control moment gyroscopes | |
| CN110282121A (en) | Coordinate system conversion method, its device, aircraft course control method and aircraft | |
| Sun | Constrained adaptive fault-tolerant attitude tracking control of rigid spacecraft | |
| CN109164819A (en) | The contragradience adaptive sliding mode large angle attitude control method of rigid body spacecraft | |
| CN113815903B (en) | Flywheel zero crossing avoidance method for remote sensing satellite | |
| CN111781833A (en) | Spacecraft online optimal attitude avoidance control method based on state dependence decomposition | |
| CN108427429B (en) | Spacecraft visual axis maneuvering control method considering dynamic pointing constraint | |
| CN111026154A (en) | Six-degree-of-freedom cooperative control method for preventing collision in spacecraft formation |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |