CN108181807B - A kind of satellite initial state stage self-adapted tolerance attitude control method - Google Patents
A kind of satellite initial state stage self-adapted tolerance attitude control method Download PDFInfo
- Publication number
- CN108181807B CN108181807B CN201711276396.2A CN201711276396A CN108181807B CN 108181807 B CN108181807 B CN 108181807B CN 201711276396 A CN201711276396 A CN 201711276396A CN 108181807 B CN108181807 B CN 108181807B
- Authority
- CN
- China
- Prior art keywords
- satellite
- attitude
- matrix
- fault
- 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.)
- Active
Links
- 238000000034 method Methods 0.000 title claims abstract description 58
- 238000013507 mapping Methods 0.000 claims abstract description 19
- 230000008569 process Effects 0.000 claims abstract description 19
- 230000003044 adaptive effect Effects 0.000 claims abstract description 13
- 239000011159 matrix material Substances 0.000 claims description 32
- 239000013598 vector Substances 0.000 claims description 13
- 230000001052 transient effect Effects 0.000 claims description 11
- 238000012417 linear regression Methods 0.000 claims description 9
- 238000006243 chemical reaction Methods 0.000 claims description 3
- 238000009434 installation Methods 0.000 claims description 3
- 230000003247 decreasing effect Effects 0.000 claims description 2
- 238000000926 separation method Methods 0.000 claims description 2
- 230000006641 stabilisation Effects 0.000 abstract description 5
- 238000011105 stabilization Methods 0.000 abstract description 5
- 230000004927 fusion Effects 0.000 abstract 1
- 238000013461 design Methods 0.000 description 4
- 238000010586 diagram Methods 0.000 description 3
- 238000005516 engineering process Methods 0.000 description 3
- 238000004364 calculation method Methods 0.000 description 2
- 230000008859 change Effects 0.000 description 2
- 238000012544 monitoring process Methods 0.000 description 2
- 230000009471 action Effects 0.000 description 1
- 238000007792 addition Methods 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000001514 detection method Methods 0.000 description 1
- 230000007257 malfunction Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
- 238000006467 substitution reaction 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
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/244—Spacecraft control systems
Landscapes
- Engineering & Computer Science (AREA)
- Automation & Control Theory (AREA)
- Remote Sensing (AREA)
- General Physics & Mathematics (AREA)
- Evolutionary Computation (AREA)
- Medical Informatics (AREA)
- Software Systems (AREA)
- Physics & Mathematics (AREA)
- Health & Medical Sciences (AREA)
- Computer Vision & Pattern Recognition (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Radar, Positioning & Navigation (AREA)
- Artificial Intelligence (AREA)
- Aviation & Aerospace Engineering (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
A kind of satellite initial state stage self-adapted tolerance attitude control method, including it is successfully separated the rear initial state control stage based on the satellite and the rocket, it establishes Stretching of solar wings process satellite rotary inertia uncertainty models: establishing Satellite Attitude Dynamics model；Using default capabilities function, the Nonlinear Mapping model of posture is established；Based on establishing Stretching of solar wings process satellite rotary inertia uncertainty models, the Nonlinear Mapping model of Satellite Attitude Dynamics model and posture, adaptive fusion device is designed using Backstepping, pass through the attitude stabilization problem this method solve satellite when initial state controls uncertain stage rotary inertia, actuator failures and influenced by external disturbance torque, it ensure that the fault-tolerant ability and robustness of attitude control system, and ensure that posture convergence rate, hyperharmonic convergence error meet preset requirement.
Description
Technical Field
The invention relates to the technical field of spacecraft control, in particular to a satellite initial state stage self-adaptive fault-tolerant attitude control method.
Background
Aerospace technology gradually affects people's daily life, and satellites play more and more important roles in communication, navigation, weather and other aspects. After the satellite and the arrow are separated, the satellite enters an initial state control stage. In the initial control stage, the solar sailboard installed on the satellite slowly unfolds. The unfolding of the solar panels necessarily changes the rotational inertia of the whole satellite, which brings certain challenges to the attitude stability control of the satellite. Although the initial control phase time is short relative to the whole satellite life, because the initial control is the basis of the satellite steady-state control, the failure of the initial control causes the malfunction of the whole satellite attitude, so that the satellite cannot complete the subsequent task in the orbit period, and therefore, the attitude stability control of the satellite initial phase is particularly important. Meanwhile, the attitude control system of the satellite is a system with a complex structure and works in a severe space environment, so that the possibility of failure of a satellite actuator is increased. Moreover, the satellite is also affected by various disturbance moments from the external environment in space, which requires a certain anti-jamming capability and fault-tolerant capability of the attitude control system. In addition, the transient state and the steady state performance of the attitude are ensured in the initial state stable control of the satellite, the occurrence of large overshoot and steady state errors is avoided, and the realization of the attitude stabilization as soon as possible is the key for ensuring the accuracy and the stability of the satellite attitude control system. Therefore, the satellite attitude control system in the initial state control stage can realize fault-tolerant control under the action of uncertainty of rotational inertia, faults of an actuator and external interference, and the condition that the attitude can be converged according to the preset requirements of transient state and steady state performance is an important task for stably controlling the attitude of the satellite in the initial state stage.
For the problem of uncertainty of rotational inertia of a spacecraft, CN201610369411 firstly establishes a kinematics and dynamics model of a combined spacecraft in a multi-source interference environment, then designs an interference observer to estimate uncertainty of rotational inertia and unmodeled dynamic interference, and finally designs an anti-saturation attitude stabilization controller in combination with the interference observer to realize composite layered anti-interference attitude stabilization of the spacecraft, but neglects the influence of failure of an actuator and does not relate to the problem of ensuring transient and steady performance in attitude control; in the patent CN201611012008, by establishing a fault detection observer adopting an adaptive threshold technology and a fault estimation observer based on the adaptive technology, online real-time monitoring and estimation are implemented on an actuator failure fault and a deviation fault of a spacecraft, and then a backstepping sliding mode fault-tolerant controller is designed to implement attitude stabilization, however, online real-time monitoring and estimation will bring huge calculation amount to an on-board computer, higher requirements are put forward on the calculation capability and storage space of the on-board computer, and the complexity of an attitude control system is increased; in the patent CN201610217207, aiming at the problem that the actuator fault, the external disturbance and the control moment amplitude are limited, an integral sliding mode fault-tolerant controller is designed to effectively improve the stability of the attitude control system and the robustness to the external disturbance when the actuator of the in-orbit spacecraft fails, but the influence of the uncertainty of the rotational inertia on the attitude control system is not considered; in the patent CN201611208106, under the condition that an inertia matrix is unknown and includes external disturbance, a preset performance index is designed, nonlinear mapping is performed on an attitude tracking control model, and finally, a model-free robust controller is designed to realize stable attitude tracking of a spacecraft, but the problem that an actuator fails is not considered. Therefore, the design of the fault-tolerant controller, the compensation of the actuator faults, the robustness on the uncertainty of the rotational inertia of the satellite and the external disturbance, the prior design and the guarantee of the transient state and the steady state performance of the satellite attitude can be realized, and the core problem of ensuring the safe and efficient performance of the satellite space task is solved.
Disclosure of Invention
The invention aims to overcome the defects of the prior art, provides a self-adaptive fault-tolerant attitude control method for a satellite initial state stage, and provides a self-adaptive fault-tolerant attitude control method based on a back-stepping method aiming at the problems of uncertainty of the rotational inertia of the whole satellite caused by the expansion of a solar sailboard and the existence of actuator faults and various external disturbance moments in the initial state control stage of the satellite.
The invention provides a self-adaptive fault-tolerant attitude control method for a satellite initial state stage, which comprises the following steps of:
(1) based on an initial state control stage after the successful separation of the satellite and the arrow, establishing a satellite rotational inertia uncertainty model in the solar sailboard unfolding process:
(2) establishing a satellite attitude dynamic model;
(3) establishing a nonlinear mapping model of the attitude by using a preset performance function;
(4) based on the establishment of a satellite rotational inertia uncertainty model, a satellite attitude dynamics model and a nonlinear mapping model of the attitude in the solar panel unfolding process, a self-adaptive fault-tolerant controller is designed by utilizing a back-stepping method.
Further, the uncertainty model of the satellite rotational inertia in the step (1) is
J＝Jm-Jnψ(t)
Wherein,the moment of inertia matrix of the whole star is a positive definite symmetric matrix;is a rigid part in the moment of inertia and is an unknown and unchangeable symmetric matrix;is a non-rigid part of the moment of inertia, is time-varying,is the gain of the non-rigid part, is an unknown constant matrix,is known to be time-varying and reflects the movement of the center of mass of the windsurfing board during the unfolding process of the solar windsurfing board.
Further, the first order differential of the whole-star moment of inertia J over time can be written in the form:
further, the building of the satellite attitude dynamics model in the step (2) specifically includes:
wherein,is the quaternion of the attitude unit of the satellite and expresses the rotation attitude of the satellite relative to an inertial coordinate system under a body coordinate system, q0Scalar part of quaternion, qvIs a vector part and satisfiesω＝[ω1,ω2,ω3]TThe attitude angular velocity of the satellite relative to an inertial coordinate system under a body coordinate system; j is a rotational inertia matrix of the whole star; d is belonged to R3×nThe installation matrix of the actuator is represented, the rank (D) is 3, and n is more than or equal to 3;the control torque that is actually output by the actuator is represented,e is an actuator failure matrix, 0 ≦ ei(t). ltoreq.1 represents the failure condition of each actuator when eiWhen t is 1, the actuator works normally, and when 0 < ei(t) < 1 indicates that the actuator has a partial failure condition when ei(t) ═ 0 indicates that the actuator is in complete failure;representing the control moment actually generated by the actuator;indicating an additional fault with the bias torque, assuming it is bounded;representing the disturbance moment in the space environment to which the satellite is subjected, given its bounding, satisfies Representing an identity matrix;representation with respect to arbitrary vectorsOf the following form:
further, the step (3) is specifically:
carrying out nonlinear mapping on the attitude by utilizing a preset performance function, and defining the function:
wherein, epsilon is [ epsilon [ ]1,ε2,ε3]TThe variable is subjected to nonlinear mapping;representing a selected predetermined performance function, strictly non-negative and decreasing, p0＝[ρ10,ρ20,ρ30]TIs an initial value of a performance function, and pi0＞0，ρi∞＝[ρ1∞,ρ2∞,ρ3∞]TRepresents the steady state value of the performance function, and pi∞＞0；l＝[l1,l2,l3]TDetermining the convergence speed of the performance function; q. q.svi(0) The initial value of the attitude quaternion vector is obtained; q. q.svi(t) and ρiThe relationship of (t) satisfies the following condition:
δ＝[δ1,δ2,δ3]Tfor presetting performance parameters, by changing the value thereof, q is reflectedviThe magnitude of overshoot during the convergence transient, when deltaiWhen the attitude quaternion vector is 0, the partial convergence process of the attitude quaternion vector does not overshoot; guaranteeing q by presetting a performance functionviTransient and steady state performance of convergence;
obtaining a variable epsilon after nonlinear mapping by solving an inverse function T of S (epsilon)iHas the following forms:
the form of T is selected as follows:
the first differential over time is as follows:
wherein,
further, by non-linearly mapping the pose, the variable ε converges, making qvAnd converging according to preset transient and steady-state performance, and meeting the requirements of steady-state errors, convergence speed and overshoot.
Further, the step (4) is specifically
Introducing two new variables z1＝ε，z2The first derivative of time is determined as follows:
by using the back stepping method, the virtual controller can be obtained:
wherein c is1The gain of the virtual controller is more than 0, and gamma is more than 0;
and (3) performing a conversion model by using a nonlinear regression matrix to obtain:
whereinAbout JmThe amount of linear regression of (a),is related to JnA linear regression quantity; w1，W2，W3Is a matrix corresponding to the linear regression variables of the moment of inertia, satisfying the following form:
based on the virtual controller, the fault-tolerant controller is designed as follows:
the adaptive update law is:
wherein, b [ | | θ |, | | σ |, d |, b [ | | θ |, | | σ | ] |, d |, c, andm]T；Φ＝[||W1||,||W2+W3||,1]T；c2＞0，c3＞0，c4> 0 is a gain parameter associated with the controller; the parameter selection satisfies the condition 1/(2 gammac)2)＜λmin(DEDT)。
Compared with the prior art, the attitude fault-tolerant control method for the initial state control stage of the satellite, which is designed by considering the uncertainty of the rotational inertia, the fault of the actuator and the external disturbance, has the advantages that:
(1) the self-adaptive fault-tolerant attitude control method for the initial state stage of the satellite aims at the solar sailboard unfolding process of the initial state control stage of the satellite, and the moment of inertia can generate larger uncertainty in the process, so that the method has strong engineering significance;
(2) the transient state and steady state performance of the attitude convergence process are considered, the attitude convergence method is restricted by using the preset performance function, so that the larger overshoot is avoided, the convergence error is reduced, the attitude does not exceed the preset convergence speed for convergence, and the space task is ensured to be safely and efficiently performed;
(3) on the basis of ensuring the preset performance of the attitude, failure faults and deviation faults of the actuator in practice are considered, the fault and deviation faults are processed by a self-adaptive method, and specific information of the faults and interference does not need to be known exactly, so that the attitude control system has stronger robustness and fault-tolerant capability on uncertainty of the rotational inertia and faults and interference of the actuator.
Drawings
FIG. 1 is a flow chart of a method for adaptive fault-tolerant attitude control at a satellite initial state stage according to the present invention;
FIG. 2 is a diagram illustrating predetermined performance;
fig. 3 is a schematic diagram of the change of the moment of inertia in the initial state control stage of a satellite with two symmetrically installed solar sailboards.
Detailed Description
Reference will now be made in detail to the embodiments of the present invention, the following examples of which are intended to be illustrative only and are not to be construed as limiting the scope of the invention.
As shown in fig. 1, the method for controlling adaptive fault-tolerant attitude in initial stage of satellite of the present invention comprises the following steps: firstly, establishing a rotational inertia uncertainty model when a solar panel is unfolded in the initial state control process of a satellite; then, based on a rotational inertia uncertainty model, a satellite attitude dynamics model containing actuator faults and external interference is established; then, establishing a nonlinear mapping model of the attitude by using a preset performance function; finally, a self-adaptive fault-tolerant controller is designed by adopting a backstepping method, the functional block diagram of the whole system is shown in fig. 1, and the specific implementation steps are as follows (the attitude fault-tolerant control in the initial state control stage of the satellite with two solar panels and six thrusters symmetrically arranged is taken as an example to illustrate the specific implementation of the method):
first, as shown in fig. 3, a satellite initial state control stage is established, and a model of variation of the moment of inertia of the satellite during the unfolding process of the solar panel is:
J＝Jm-Jnψ(t)
whereinThe moment of inertia matrix of the whole star is a positive definite symmetric matrix;is a rigid part in the moment of inertia and is an unknown and unchangeable symmetric matrix;is a non-rigid part of the moment of inertia, is time-varying,is the gain of the non-rigid part, is an unknown constant matrix,is known to be time-varying and reflects the movement of the center of mass of the windsurfing board during the unfolding process of the solar windsurfing board. The first order differential of the moment of inertia J of the whole star over time can be written in the form:
according to the actual satelliteDesign parameters, take Jm＝[950 10 5；10 600 30；5 30 360]Tkg·m2Considering that two solar sailboards with the same mass are symmetrically arranged on two sides of the satellite, and the mass is m1＝m2100kg, and in the process of unfolding the solar sailboards, the change rule of the mass centers of the two solar sailboards along with time is known as follows: e＝{ex,ey,ezand is a unit base coordinate related to the satellite body coordinate system, when the time is less than or equal to 10s, k is 1, when the time is greater than 10s, k is 10/t, the unfolding time of the solar panel is 10s, and the non-rigid part in the moment of inertia can be expressed asI.e. non-rigid part of the moment of inertia gain Jn＝-[100I3,100I3]Law of motion of center of mass of sailboard
Secondly, establishing an attitude dynamics model of the satellite attitude control system in an initial state control stage as follows:
wherein,is the quaternion of the attitude unit of the satellite and expresses the rotation attitude of the satellite relative to an inertial coordinate system under a body coordinate system, q0Is a scalar quantity, qvIs a vector, satisfyThe initial attitude was taken to be q (0) — [0.2, -0.15, -0.25,0.9354]T；ω＝[ω1,ω2,ω3]TFor the attitude angular velocity of the satellite relative to the inertial coordinate system in the body coordinate system, the initial angular velocity is ω (0) — [0.05, -0.05, 0.01%]Trad/s; j is the moment of inertia of the whole star, and the value is shown in the previous step; d is belonged to R3×6Is a mounting matrix of the thruster, satisfies the rank (D) 3, and is taken as The control torque that is actually output by the actuator is represented,e is an actuator failure matrix, 0 ≦ ei(t). ltoreq.1 represents the failure condition of each actuator when eiWhen t is 1, the actuator is in normal working state, and when 0 < ei(t) < 1 indicates that the actuator has a partial failure condition when ei(t) ═ 0 represents that the actuator is in a complete failure state, and the values are:
the control moment actually generated by the thruster is represented, namely the control moment is a fault-tolerant controller required to be designed;and (3) representing an additional fault caused by the bias torque, and taking the following values:
representing the disturbance moment of the satellite from the space environment, and meetingdmFor the maximum value of the additional fault and the external interference, take as Representing an identity matrix;representation with respect to arbitrary vectorsOf the following form:
thirdly, nonlinear mapping is carried out on the posture by utilizing a preset performance function, and a model is established as follows:
wherein, epsilon is [ epsilon [ ]1,ε2,ε3]TIs a converted variable;is strictly non-negative and degressive, and the initial value of the performance function is rho0＝[0.4,0.35,0.45]TThe steady state value of the performance function is taken as rho∞＝[0.0001,0.0001,0.0001]T，l＝[0.1,0.1,0.1]TDetermining a performance function convergence speed; q. q.svi(0) Is the initial value of the attitude quaternion vector, taken as q as shown in the previous stepv(0)＝[0.2,-0.15,-0.25]T；qvi(t) and ρiThe relationship of (t) satisfies the following form:
δ＝[δ1,δ2,δ3]Tfor presetting performance parameters, by changing the value thereof, q is reflectedviThe overshoot in the convergence process is taken to be δ ═ 0,0]TAnd indicating that the attitude q convergence process is not overshot.
According to the nonlinear mapping model, the form of T is selected as follows:
the first differential over time is as follows:
wherein,
by nonlinear mapping of the attitude, the variable epsilon converges to make qvAnd convergence is carried out according to preset transient state and steady state performance, and requirements on aspects of steady state errors, convergence speed, overshoot and the like are met.
Fourthly, considering a self-adaptive backstepping fault-tolerant control method of the uncertainty of the rotational inertia, the fault of the thruster and the external interference, and designing a self-adaptive controller by using a backstepping method:
first, two new variables z are introduced1＝ε，z2The first derivative of time is determined as follows:
based on the back-stepping method, a virtual controller is obtained:
wherein, c1The gains of the virtual controllers are more than 0 and gamma more than 0, and c is respectively selected1＝0.2，γ＝0.01。
And (3) performing a conversion model by using a nonlinear regression matrix to obtain:
whereinAbout JmThe amount of linear regression of (a),is related to JnA linear regression quantity; w1，W2，W3Is a matrix corresponding to the linear regression variables of the moment of inertia, satisfying the following form:
based on the virtual controller, designing a fault-tolerant controller and a self-adaptive updating law respectively as follows:
wherein，b＝[||θ||,||σ||,dm]T；Φ＝[||W1||,||W2+W3||,1]T；c2＞0，c3＞0，c4> 0 is a gain parameter associated with the controller, selected by the designer as c2＝270，c3＝0.001，c410; minimum eigenvalue λ of the product of installation matrix and failure matrixmin(DEDT) 0.3920, and 1/(2 γ c) can be calculated2) 0.1852, so that condition 1/(2 γ c) is satisfied2)＜λmin(DEDT)。
By Matlab simulation, a satellite initial state stage self-adaptive fault-tolerant attitude control method can be obtained, the satellite attitude can be converged according to preset transient state and steady state performance under the consideration of the uncertainty of the rotational inertia caused by the unfolding of the solar sailboard, the fault of a thruster and the external interference in the initial state control stage, the method has strong robustness and fault-tolerant capability, and can meet the performance requirements of small enough attitude steady state error, fast convergence rate, no overshoot and the like.
Although exemplary embodiments of the present invention have been described for illustrative purposes, those skilled in the art will appreciate that various modifications, additions, substitutions and the like can be made in form and detail without departing from the scope and spirit of the invention as disclosed in the accompanying claims, all of which are intended to fall within the scope of the claims, and that various steps in the various sections and methods of the claimed product can be combined together in any combination. Therefore, the description of the embodiments disclosed in the present invention is not intended to limit the scope of the present invention, but to describe the present invention. Accordingly, the scope of the present invention is not limited by the above embodiments, but is defined by the claims or their equivalents.
Claims (6)
1. A satellite initial state stage self-adaptive fault-tolerant attitude control method is characterized by comprising the following steps:
(1) based on an initial state control stage after the successful separation of the satellite and the arrow, establishing a satellite rotational inertia uncertainty model in the solar sailboard unfolding process:
(2) establishing a satellite attitude dynamic model;
(3) establishing a nonlinear mapping model of the attitude by using a preset performance function;
(4) designing a self-adaptive fault-tolerant controller by utilizing a backstepping method based on establishing a satellite rotational inertia uncertainty model, a satellite attitude dynamics model and a nonlinear mapping model of the attitude in the solar panel unfolding process;
the establishment of the satellite attitude dynamics model in the step (2) specifically comprises the following steps:
Q(q)＝q0I3+qv ×
wherein,is the quaternion of the attitude unit of the satellite and expresses the rotation attitude of the satellite relative to an inertial coordinate system under a body coordinate system, q0Scalar part of quaternion, qvIs a vector part and satisfiesω＝[ω1,ω2,ω3]TThe attitude angular velocity of the satellite relative to an inertial coordinate system under a body coordinate system; j is a rotational inertia matrix of the whole satellite system; d is belonged to R3×nThe installation matrix of the actuator is represented, the rank (D) is 3, and n is more than or equal to 3;the control torque that is actually output by the actuator is represented,e is an actuator failure matrix, 0 ≦ ei(t) 1 or less indicates the failure condition of each actuatorWhen e isiWhen t is 1, the actuator works normally, and when 0 < ei(t) < 1 indicates that the actuator has a partial failure condition when ei(t) ═ 0 indicates that the actuator is in complete failure;representing the control moment actually generated by the actuator;indicating an additional fault with the bias torque, assuming it is bounded;representing the disturbance moment in the space environment to which the satellite is subjected, given its bounding, satisfies Representing an identity matrix;representation with respect to arbitrary vectorsOf the following form:
2. the method according to claim 1, wherein the adaptive fault-tolerant attitude control method comprises: the uncertainty model of the satellite rotational inertia in the step (1) is as follows:
J＝Jm-Jnψ(t)
wherein,the moment of inertia matrix of the whole star is a positive definite symmetric matrix;is a rigid part in the moment of inertia and is an unknown and unchangeable symmetric matrix;is a non-rigid part of the moment of inertia, is time-varying,is the gain of the non-rigid part, is an unknown constant matrix,is known to be time-varying and reflects the movement of the center of mass of the windsurfing board during the unfolding process of the solar windsurfing board.
3. The method according to claim 2, wherein the adaptive fault-tolerant attitude control method in the initial stage of the satellite comprises: the first order differential of the moment of inertia J of the whole star over time can be written in the form:
4. the method according to claim 1, wherein the adaptive fault-tolerant attitude control method comprises: the step (3) is specifically as follows:
carrying out nonlinear mapping on the attitude by utilizing a preset performance function, and defining the function:
wherein, epsilon is [ epsilon [ ]1,ε2,ε3]TThe variable is subjected to nonlinear mapping;representing a selected predetermined performance function, strictly non-negative and decreasing, p0＝[ρ10,ρ20,ρ30]TIs an initial value of a performance function, and pi0＞0，ρi∞＝[ρ1∞,ρ2∞,ρ3∞]TRepresents the steady state value of the performance function, and pi∞＞0；l＝[l1,l2,l3]TDetermining the convergence speed of the performance function; q. q.svi(0) The initial value of the attitude quaternion vector is obtained; q. q.svi(t) and ρiThe relationship of (t) satisfies the following condition:
δ＝[δ1,δ2,δ3]Tfor presetting performance parameters, by changing the value thereof, q is reflectedviThe magnitude of overshoot during the convergence transient, when deltaiWhen the attitude quaternion vector is 0, the partial convergence process of the attitude quaternion vector does not overshoot; guaranteeing q by presetting a performance functionviTransient and steady state performance of convergence;
obtaining a variable epsilon after nonlinear mapping by solving an inverse function T of S (epsilon)iHas the following forms:
the form of T is selected as follows:
the first differential over time is as follows:
wherein,
5. the method according to claim 4, wherein the adaptive fault-tolerant attitude control method in the initial stage of the satellite comprises: by non-linearly mapping the attitude, the variable ε converges, making qvAnd converging according to preset transient and steady-state performance, and meeting the requirements of steady-state errors, convergence speed and overshoot.
6. The method according to claim 1, wherein the adaptive fault-tolerant attitude control method comprises: the step (4) is specifically that
Introducing two new variables z1＝ε，z2The first derivative of time is determined as follows:
by using the back stepping method, the virtual controller can be obtained:
wherein, c1The gain of the virtual controller is more than 0, and gamma is more than 0;
and (3) performing a conversion model by using a nonlinear regression matrix to obtain:
whereinAbout JmThe amount of linear regression of (a),is related to JnA linear regression quantity; w1，W2，W3Is a matrix corresponding to the linear regression variables of the moment of inertia, satisfying the following form:
based on the virtual controller, the fault-tolerant controller is designed as follows:
the adaptive update law is:
wherein, b [ | | θ |, | | σ |, d |, b [ | | θ |, | | σ | ] |, d |, c, andm]T；Φ＝[||W1||,||W2+W3||,1]T；c2＞0，c3＞0，c4> 0 is a gain parameter associated with the controller;the parameter selection satisfies the condition 1/(2 gammac)2)＜λmin(DEDT)。
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201711276396.2A CN108181807B (en) | 2017-12-06 | 2017-12-06 | A kind of satellite initial state stage self-adapted tolerance attitude control method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201711276396.2A CN108181807B (en) | 2017-12-06 | 2017-12-06 | A kind of satellite initial state stage self-adapted tolerance attitude control method |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN108181807A CN108181807A (en) | 2018-06-19 |
| CN108181807B true CN108181807B (en) | 2019-03-29 |
Family
ID=62545692
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201711276396.2A Active CN108181807B (en) | 2017-12-06 | 2017-12-06 | A kind of satellite initial state stage self-adapted tolerance attitude control method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN108181807B (en) |
Families Citing this family (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110340898B (en) * | 2019-08-22 | 2020-12-01 | 北京航空航天大学 | Self-adaptive fault-tolerant control method for free floating space manipulator |
| CN110667890B (en) * | 2019-09-10 | 2022-06-28 | 北方信息控制研究院集团有限公司 | Nonlinear attitude stabilization method for spacecraft |
| CN110618694B (en) * | 2019-09-10 | 2022-08-12 | 北方信息控制研究院集团有限公司 | Robust attitude control method for space interceptor |
| CN110733668B (en) * | 2019-10-16 | 2022-08-30 | 中山大学 | Flexible satellite active vibration suppression method based on component force synthesis and fixed time convergence |
| CN111045441B (en) * | 2019-12-19 | 2022-04-08 | 南京航空航天大学 | Hypersonic aircraft sensor composite fault self-healing control method |
| CN111258325B (en) * | 2020-01-08 | 2022-08-02 | 哈尔滨工业大学 | High-precision and high-performance attitude fault-tolerant control method and device for ground remote sensing satellite and computer storage medium |
| CN111158398A (en) * | 2020-01-15 | 2020-05-15 | 哈尔滨工业大学 | Adaptive control method of hypersonic aircraft considering attack angle constraint |
| CN111562794B (en) * | 2020-04-08 | 2021-09-14 | 中南大学 | Spacecraft attitude control method for actuator fault and input quantization |
| CN115959307B (en) * | 2023-01-18 | 2024-05-31 | 西北工业大学深圳研究院 | Method for stably controlling attitude of tethered satellite in preset time |
| CN116395148B (en) * | 2023-03-20 | 2023-09-15 | 银河航天(北京)网络技术有限公司 | Satellite and rocket separation control system and control method |
| CN117622521B (en) * | 2023-12-09 | 2024-06-21 | 北京航空航天大学 | Self-adaptive sliding mode control method for microsatellite with coiled extension arm |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104898683A (en) * | 2015-05-20 | 2015-09-09 | 哈尔滨工业大学 | Flexible satellite neural network backstepping sliding mode attitude control method |
| CN106802660A (en) * | 2017-03-09 | 2017-06-06 | 北京航天自动控制研究所 | One kind is combined strong anti-interference attitude control method |
| CN106814746A (en) * | 2017-03-24 | 2017-06-09 | 哈尔滨工业大学 | A kind of spacecraft appearance rail integration Backstepping Tracking Control |
-
2017
- 2017-12-06 CN CN201711276396.2A patent/CN108181807B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104898683A (en) * | 2015-05-20 | 2015-09-09 | 哈尔滨工业大学 | Flexible satellite neural network backstepping sliding mode attitude control method |
| CN106802660A (en) * | 2017-03-09 | 2017-06-06 | 北京航天自动控制研究所 | One kind is combined strong anti-interference attitude control method |
| CN106814746A (en) * | 2017-03-24 | 2017-06-09 | 哈尔滨工业大学 | A kind of spacecraft appearance rail integration Backstepping Tracking Control |
Also Published As
| Publication number | Publication date |
|---|---|
| CN108181807A (en) | 2018-06-19 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN108181807B (en) | A kind of satellite initial state stage self-adapted tolerance attitude control method | |
| Zheng et al. | Path following of a surface vessel with prescribed performance in the presence of input saturation and external disturbances | |
| Zheng et al. | Error-constrained LOS path following of a surface vessel with actuator saturation and faults | |
| Xiao et al. | A novel disturbance estimation scheme for formation control of ocean surface vessels | |
| Li et al. | Observer-based fault-tolerant attitude control for rigid spacecraft | |
| CN106773679B (en) | A kind of spacecraft fault tolerant control method based on angular speed observer | |
| Zhang et al. | Adaptive sliding mode control for re-entry attitude of near space hypersonic vehicle based on backstepping design | |
| CN106292681B (en) | A kind of satellite Active Fault-tolerant Control Method distributed based on observer and On-line Control | |
| Li et al. | Fault-tolerant attitude stabilization incorporating closed-loop control allocation under actuator failure | |
| Qinglei et al. | Adaptive integral-type sliding mode control for spacecraft attitude maneuvering under actuator stuck failures | |
| Zhang et al. | Hybrid fuzzy adaptive fault-tolerant control for a class of uncertain nonlinear systems with unmeasured states | |
| An et al. | Sliding mode disturbance observer-enhanced adaptive control for the air-breathing hypersonic flight vehicle | |
| Lee | Fault-tolerant finite-time controller for attitude tracking of rigid spacecraft using intermediate quaternion | |
| Chunodkar et al. | Switching angular velocity observer for rigid-body attitude stabilization and tracking control | |
| Bu et al. | Novel auxiliary error compensation design for the adaptive neural control of a constrained flexible air-breathing hypersonic vehicle | |
| CN104880948A (en) | Fault tolerant control method for a spacecraft equipped with a flywheel based on robust adaptation | |
| CN109507890A (en) | A kind of unmanned plane dynamic inverse generalized predictive controller based on ESO | |
| CN113619814B (en) | Relative attitude and orbit coupling control method for final approaching section of intersection butt joint | |
| Chen et al. | Sliding-mode-based attitude tracking control of spacecraft under reaction wheel uncertainties | |
| Wang et al. | Path following of the autonomous airship with compensation of unknown wind and modeling uncertainties | |
| Lian et al. | Full attitude control of a quadrotor using fast nonsingular terminal sliding mode with angular velocity planning | |
| Cheng et al. | Disturbance observer-based nonsingular fixed-time sliding mode tracking control for a quadcopter | |
| Hao et al. | Full-state-constrained non-certainty-equivalent adaptive control for satellite swarm subject to input fault | |
| Nadda et al. | Improved quadrotor altitude control design using second-order sliding mode | |
| Lv et al. | Strong fixed-time dynamic inverse adaptive LQR integrated control strategy for dynamic positioning of ship |
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 |