CN106444815A - Input forming control method of single-shaft maneuverable spacecraft - Google Patents
Input forming control method of single-shaft maneuverable spacecraft Download PDFInfo
- Publication number
- CN106444815A CN106444815A CN201610976806.3A CN201610976806A CN106444815A CN 106444815 A CN106444815 A CN 106444815A CN 201610976806 A CN201610976806 A CN 201610976806A CN 106444815 A CN106444815 A CN 106444815A
- Authority
- CN
- China
- Prior art keywords
- spacecraft
- represent
- attitude
- omega
- satellite
- 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 26
- 239000011159 matrix material Substances 0.000 claims abstract description 26
- 230000001133 acceleration Effects 0.000 claims abstract description 11
- 238000000465 moulding Methods 0.000 claims description 44
- 230000010355 oscillation Effects 0.000 claims description 11
- 230000009471 action Effects 0.000 claims description 7
- 238000004364 calculation method Methods 0.000 claims description 3
- 230000008878 coupling Effects 0.000 claims description 3
- 238000010168 coupling process Methods 0.000 claims description 3
- 238000005859 coupling reaction Methods 0.000 claims description 3
- 238000013016 damping Methods 0.000 claims description 3
- 229910017435 S2 In Inorganic materials 0.000 claims 1
- 230000001268 conjugating effect Effects 0.000 abstract 1
- 230000008569 process Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008901 benefit Effects 0.000 description 1
- 230000009514 concussion Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000006467 substitution reaction 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)
- Moulding By Coating Moulds (AREA)
Abstract
The invention relates to an input forming control method of a single-shaft maneuverable spacecraft. The method comprises the steps that S1, the posture and posture angular velocity motion path of the spacecraft are planned according to the task requirement of the spacecraft with a flexible accessory, and the state equation of posture motion of the spacecraft is set up; S2, a simplified system feature matrix is applied to approximately calculating a feature value corresponding to the flexible accessory of the spacecraft, and the equivalent vibration frequency of the flexible accessory is calculated; S3, the conjugating feature value of an input former is calculated, parameters of the input former are obtained, and design of the input former is completed; S4, an angular acceleration feedback item is added into a controller of the spacecraft, and residual vibration caused by simplification of the posture motion state equation is compensated for. The method is suitable for the single-shaped motorized equation around the Euler shaft, the solving purpose of a nonlinear system equation is achieved by simplifying a system feature matrix, and the residual motion caused by simplification of the system feature matrix is eliminated through angular acceleration compensation.
Description
Technical field
The present invention relates to a kind of molding control method, specifically refer to a kind of being applied to around the motor-driven boat of Euler's axle single shaft
The molding control method of its device, can suppress the vibration exciting in mobile process, realize attitude fast reserve and quickly steady
Fixed.
Background technology
Task and function with modern spacecraft platform become more diversified, to the control accuracy of spacecraft, degree of stability, special
It is not that fast reserve ability is put forward higher requirement.But for the spacecraft with flexible appendage, in fast reserve process
In how to reduce the vibration of flexible appendage and affect on produced by control system, be particularly important.
Traditional molding method is a kind of forming method for opened loop control moment, if directly in closed loop control
The moment outfan of device adds molding device, not only will not eliminate vibration, can lead to the concussion of system on the contrary.And apply at present
Comparison extensive molding method primary limitation in linear control system, for attitude fast reserve control system
Speech, the no matter kinematical equation of the spacecraft or attitude controller of spacecraft is required for adding nonlinear link, and for
Vibration suppressing method containing nonlinear element is it is impossible to apply existing molding method.
Based on above-mentioned, need badly at present and propose a kind of molding control method of single shaft maneuverable spacecraft, machine can be suppressed
The vibration exciting during dynamic, realizes attitude fast reserve and fast and stable.
Content of the invention
It is an object of the invention to provide a kind of molding control method of single shaft maneuverable spacecraft is it is adaptable to around Euler's axle
The motor-driven spacecraft of single shaft, realizes the Solve problems of nonlinear system equation by simplified system eigenmatrix, by angle plus
Velocity compensation is eliminating system features matrix because simplifying brought residual oscillation.
For achieving the above object, the present invention provide a kind of molding control method of single shaft maneuverable spacecraft it is adaptable to
Around the motor-driven spacecraft of Euler's axle single shaft, comprise the steps of:
S1, the mission requirements according to the spacecraft containing flexible appendage, plan attitude and the attitude angle speed of this spacecraft
Degree motion path, sets up the state equation of this attitude motion of spacecraft;
The system features matrix that S2, application simplify carrys out the corresponding eigenvalue of approximate calculation spacecraft flexible appendage;
S3, calculate the Con-eigenvalue of molding device, ask for the parameter of molding device, complete molding device
Design;
S4, in the controller of spacecraft add Angular Acceleration Feedback item, compensate attitude motion state equation because simplify institute
The residual oscillation causing.
In described S1, the state equation of the attitude motion of spacecraft containing flexible appendage is:
Wherein, 0 is the null matrix of corresponding dimension;E is the unit matrix of corresponding dimension;
q0Represent Satellite Attitude
The scalar component of state quaternary number;E3Represent 3 rank unit matrixs;ucRepresent control moment;qvRepresent the vector of attitude of satellite quaternary number
Part;Represent the projection in measuring satellite angular velocities three directions under co-ordinates of satellite system;Brot is the coupling moment of flexible appendage
Battle array;ωcIt is the diagonal matrix with each rank flexible vibration frequency as element;I represents the inertia matrix of satellite;ξcResistance for flexible vibration
Buddhist nun's ratio, ηiArray for the mode of oscillation composition of i-th flexible appendage.
In described S2, the attitude motion state equation of simplification is:
Wherein,For the system features matrix simplifying;Solve the attitude motion state equation of this simplification, obtain spacecraft and scratch
The property corresponding eigenvalue λ of adnexa.
In described S3, the parameter of molding device is two, and the action time of one of parameter is 0, and amplitude is 1/
(1+K), the action time of another parameter is T, and amplitude is K/ (1+K);Wherein, the expression formula of K and T is:
ξ and ω according to obtaining after calculating in S3 is Con-eigenvalue, and ξ and ω represents the equivalent resistance of flexible appendage respectively
Buddhist nun's ratio and frequency of vibration, thus obtaining parameter value T and K of molding device, it is Shaper that design completes molding device.
In described S4, in the controller of spacecraft, add angular acceleration molding item, compensate attitude motion state
Equation affects because of the residual oscillation caused by simplification error, obtains the control moment u of controllercFor:
In formula, Shaper represents molding device;KpAnd KdRepresent PD control parameter, qvRepresent attitude of satellite quaternary number
Vector section;qvdRepresent that satellite expects the vector section of attitude quaternion;qv0Represent the arrow of initial time attitude of satellite quaternary number
Amount part;Represent the projection in measuring satellite angular velocities three directions under co-ordinates of satellite system;Represent initial satellite attitude
The projection in angular velocity three directions under co-ordinates of satellite system;adRepresent desired angular acceleration;ωdRepresent desired angular velocity;e0
Represent initial Euler's axle.
In sum, the molding control method of the single shaft maneuverable spacecraft that the present invention provides is it is adaptable to around Euler's axle
The motor-driven spacecraft of single shaft, realizes the Solve problems of nonlinear system equation by simplified system eigenmatrix, by angle plus
Velocity compensation is eliminating system features matrix because simplifying brought residual oscillation.
Brief description
Fig. 1 is the flow chart of the molding control method of single shaft maneuverable spacecraft in the present invention.
Specific embodiment
Below in conjunction with Fig. 1, describe a preferred embodiment of the present invention in detail.
As shown in figure 1, the molding control method of the single shaft maneuverable spacecraft providing for the present invention is it is adaptable to around Euler
The motor-driven spacecraft of axle single shaft, comprises the steps of:
S1, the mission requirements according to the spacecraft containing flexible appendage, plan attitude and the attitude angle speed of this spacecraft
Degree motion path, sets up the state equation of this attitude motion of spacecraft;
The system features matrix that S2, application simplify carrys out the corresponding eigenvalue of approximate calculation spacecraft flexible appendage, and calculates
The equivalent frequency of vibration of flexible appendage;
S3, calculate the Con-eigenvalue of molding device, ask for the parameter of molding device, complete molding device
Design;
S4, in the controller of spacecraft add Angular Acceleration Feedback item, compensate attitude motion state equation because simplify institute
The residual oscillation causing.
In described S1, the state equation of the attitude motion of spacecraft containing flexible appendage is:
Wherein, 0 is the null matrix of corresponding dimension;E is the unit matrix of corresponding dimension;Ic=E-Brot BrotT;H=
(E-BrotTI-1Brot)-1;q0Represent the scalar component of attitude of satellite quaternary number;E3Represent 3 rank units
Matrix;ucRepresent control moment;qvRepresent the vector section of attitude of satellite quaternary number;Represent measuring satellite angular velocities in satellite
The projection in lower three directions of coordinate system;Brot is the coupling matrix of flexible appendage;ωcIt is with each rank flexible vibration frequency as element
Diagonal matrix;I represents the inertia matrix of satellite;ξcFor the damping ratio of flexible vibration, ηiMode of oscillation for i-th flexible appendage
The array of composition.
In described S2, the attitude motion state equation of simplification is:
Wherein,For the system features matrix simplifying;Solve the attitude motion state equation of this simplification, obtain spacecraft and scratch
The property corresponding eigenvalue λ of adnexa.
In described S3, the parameter (minimum pulse number) of molding device is two, the action time of one of parameter A
For 0, amplitude is 1/ (1+K), and the action time of another parameter B is T, and amplitude is K/ (1+K), specifically refers to following table:
Molding device parameter | A | B |
Action time | 0 | T |
Amplitude | 1/(1+K) | K/(1+K) |
Wherein, the expression formula of K and T is:
ξ and ω according to obtaining after calculating in S3 is Con-eigenvalue, and ξ and ω represents the equivalent resistance of flexible appendage respectively
Buddhist nun's ratio and frequency of vibration, thus obtaining parameter value T and K of molding device, it is Shaper that design completes molding device.
In described S4, in the controller of spacecraft, add angular acceleration molding item, compensate attitude motion state
Equation affects because of the residual oscillation caused by simplification error, obtains the control moment u of controllercFor:
In formula, Shaper represents molding device;KpAnd KdRepresent PD (ratio according to designed by system performance for the designer
Example differential) control parameter, qvRepresent the vector section of attitude of satellite quaternary number;qvdRepresent that satellite expects the vector of attitude quaternion
Part;qv0Represent the vector section of initial time attitude of satellite quaternary number;Represent measuring satellite angular velocities in co-ordinates of satellite system
The projection in lower three directions;Represent the projection in initial satellite attitude angular velocity three directions under co-ordinates of satellite system;adRepresent
Desired angular acceleration;ωdRepresent desired angular velocity;e0Represent initial Euler's axle.
The molding control method of the single shaft maneuverable spacecraft that the present invention provides, compared with prior art, has following
Advantage and beneficial effect:1st, it is applied to the control system containing nonlinear element;2nd, can completely suppress motor-driven in theory
During the vibration that excites;3rd, fast reserve and stable effect can be reached;4th, because vibration will be pressed down after motor-driven completely
System, therefore can be bigger by the bandwidth Design of spacecraft;5th, reliability is high, and algorithm is simple, and on star, software is easily realized.
Although present disclosure has been made to be discussed in detail by above preferred embodiment, but it should be appreciated that above-mentioned
Description is not considered as limitation of the present invention.After those skilled in the art have read the above, for the present invention's
Multiple modifications and substitutions all will be apparent from.Therefore, protection scope of the present invention should be limited to the appended claims.
Claims (5)
1. a kind of molding control method of single shaft maneuverable spacecraft is it is characterised in that be applied to motor-driven around Euler's axle single shaft
Spacecraft, comprise the steps of:
S1, the mission requirements according to the spacecraft containing flexible appendage, plan attitude and the attitude angular velocity fortune of this spacecraft
Dynamic path, sets up the state equation of this attitude motion of spacecraft;
The system features matrix that S2, application simplify carrys out the corresponding eigenvalue of approximate calculation spacecraft flexible appendage, and calculates flexibility
The equivalent frequency of vibration of adnexa;
S3, calculate the Con-eigenvalue of molding device, ask for the parameter of molding device, complete setting of molding device
Meter;
S4, in the controller of spacecraft, add Angular Acceleration Feedback item, compensate attitude motion state equation caused because simplifying
Residual oscillation.
2. the molding control method of single shaft maneuverable spacecraft as claimed in claim 1 is it is characterised in that described S1
In, the state equation of the attitude motion of spacecraft containing flexible appendage is:
Wherein, 0 is the null matrix of corresponding dimension;E is the unit matrix of corresponding dimension;Ic=E-Brot BrotT;H=(E-
BrotTI-1Brot)-1;q0Represent the scalar component of attitude of satellite quaternary number;E3Represent 3 rank unit squares
Battle array;ucRepresent control moment;qvRepresent the vector section of attitude of satellite quaternary number;Represent that measuring satellite angular velocities are sat in satellite
The projection in lower three directions of mark system;Brot is the coupling matrix of flexible appendage;ωcIt is with each rank flexible vibration frequency as element
Diagonal matrix;I represents the inertia matrix of satellite;ξcFor the damping ratio of flexible vibration, ηiMode of oscillation group for i-th flexible appendage
The array becoming.
3. the molding control method of single shaft maneuverable spacecraft as claimed in claim 2 is it is characterised in that described S2
In, the attitude motion state equation of simplification is:
Wherein, Pt *For the system features matrix simplifying;Solve the attitude motion state equation of this simplification, obtain spacecraft flexibility attached
The corresponding eigenvalue λ of part.
4. the molding control method of single shaft maneuverable spacecraft as claimed in claim 3 is it is characterised in that described S3
In, the parameter of molding device is two, and the action time of one of parameter is 0, and amplitude is 1/ (1+K), another parameter
Action time be T, amplitude be K/ (1+K);Wherein, the expression formula of K and T is:
Be Con-eigenvalue according to ξ and ω that obtain after calculating in S3, ξ and ω represent respectively the equivalent damping ratio of flexible appendage and
Frequency of vibration, thus obtaining parameter value T and K of molding device, it is Shaper that design completes molding device.
5. the molding control method of single shaft maneuverable spacecraft as claimed in claim 4 is it is characterised in that described S4
In, add angular acceleration molding item in the controller of spacecraft, compensate attitude motion state equation because of simplification error institute
The residual oscillation impact causing, obtains the control moment u of controllercFor:
In formula, Shaper represents molding device;KpAnd KdRepresent PD control parameter, qvRepresent the vector of attitude of satellite quaternary number
Part;qvdRepresent that satellite expects the vector section of attitude quaternion;qv0Represent the vector portion of initial time attitude of satellite quaternary number
Point;Represent the projection in measuring satellite angular velocities three directions under co-ordinates of satellite system;Represent initial satellite attitude angle speed
The projection in degree three directions under co-ordinates of satellite system;adRepresent desired angular acceleration;ωdRepresent desired angular velocity;e0Represent
Initial Euler's axle.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN201610976806.3A CN106444815B (en) | 2016-11-07 | 2016-11-07 | A kind of molding control method of single shaft maneuverable spacecraft |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN201610976806.3A CN106444815B (en) | 2016-11-07 | 2016-11-07 | A kind of molding control method of single shaft maneuverable spacecraft |
Publications (2)
Publication Number | Publication Date |
---|---|
CN106444815A true CN106444815A (en) | 2017-02-22 |
CN106444815B CN106444815B (en) | 2019-01-22 |
Family
ID=58180322
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
CN201610976806.3A Active CN106444815B (en) | 2016-11-07 | 2016-11-07 | A kind of molding control method of single shaft maneuverable spacecraft |
Country Status (1)
Country | Link |
---|---|
CN (1) | CN106444815B (en) |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN107807657A (en) * | 2017-11-29 | 2018-03-16 | 南京理工大学 | A kind of Flexible Spacecraft self-adaptation control method based on path planning |
CN108958275A (en) * | 2018-06-25 | 2018-12-07 | 南京理工大学 | A kind of hard and soft liquid coupled system attitude controller and motor-driven path combined optimization method |
Citations (12)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5333819A (en) * | 1993-03-12 | 1994-08-02 | General Electric Company | Self tuning motion/vibration suppression system |
US5378974A (en) * | 1993-07-02 | 1995-01-03 | The United States Of America As Represented By The Secretary Of The Air Force | Vibration damping system |
JP2000289697A (en) * | 1999-04-06 | 2000-10-17 | Ishikawajima Harima Heavy Ind Co Ltd | Pay-load damping mechanism |
CN102073276A (en) * | 2011-02-21 | 2011-05-25 | 哈尔滨工业大学 | Method for controlling flexible structure and self-adaptive changing structure by radial basis function (RBF) neural network |
CN102298390A (en) * | 2011-06-24 | 2011-12-28 | 北京航空航天大学 | Anti-disturbance flexible spacecraft attitude and vibration composite control method |
CN102654773A (en) * | 2012-05-15 | 2012-09-05 | 北京航空航天大学 | Method for controlling flexible spacecraft based on ZVDD and PWM (pulse-width modulation) mixing input former |
CN102662403A (en) * | 2012-03-21 | 2012-09-12 | 西北工业大学 | Instruction design method for changing configuration of assembly based on input shaping |
CN102736518A (en) * | 2012-07-24 | 2012-10-17 | 北京航空航天大学 | Composite anti-interference controller comprising measurement and input time delay for flexible spacecraft |
US8532847B1 (en) * | 2012-09-28 | 2013-09-10 | Fukashi Andoh | Vibration suppressing device for spacecraft |
US20140246949A1 (en) * | 2013-03-01 | 2014-09-04 | Seiko Epson Corporation | Mems device, electronic device, electronic apparatus, and moving object |
CN104932509A (en) * | 2015-05-15 | 2015-09-23 | 上海新跃仪表厂 | Ground testing system for active vibration abatement of flexible spacecraft |
CN105843237A (en) * | 2016-03-22 | 2016-08-10 | 北京航空航天大学 | Spacecraft attitude reference instruction generation method for suppressing flexible vibration |
-
2016
- 2016-11-07 CN CN201610976806.3A patent/CN106444815B/en active Active
Patent Citations (12)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5333819A (en) * | 1993-03-12 | 1994-08-02 | General Electric Company | Self tuning motion/vibration suppression system |
US5378974A (en) * | 1993-07-02 | 1995-01-03 | The United States Of America As Represented By The Secretary Of The Air Force | Vibration damping system |
JP2000289697A (en) * | 1999-04-06 | 2000-10-17 | Ishikawajima Harima Heavy Ind Co Ltd | Pay-load damping mechanism |
CN102073276A (en) * | 2011-02-21 | 2011-05-25 | 哈尔滨工业大学 | Method for controlling flexible structure and self-adaptive changing structure by radial basis function (RBF) neural network |
CN102298390A (en) * | 2011-06-24 | 2011-12-28 | 北京航空航天大学 | Anti-disturbance flexible spacecraft attitude and vibration composite control method |
CN102662403A (en) * | 2012-03-21 | 2012-09-12 | 西北工业大学 | Instruction design method for changing configuration of assembly based on input shaping |
CN102654773A (en) * | 2012-05-15 | 2012-09-05 | 北京航空航天大学 | Method for controlling flexible spacecraft based on ZVDD and PWM (pulse-width modulation) mixing input former |
CN102736518A (en) * | 2012-07-24 | 2012-10-17 | 北京航空航天大学 | Composite anti-interference controller comprising measurement and input time delay for flexible spacecraft |
US8532847B1 (en) * | 2012-09-28 | 2013-09-10 | Fukashi Andoh | Vibration suppressing device for spacecraft |
US20140246949A1 (en) * | 2013-03-01 | 2014-09-04 | Seiko Epson Corporation | Mems device, electronic device, electronic apparatus, and moving object |
CN104932509A (en) * | 2015-05-15 | 2015-09-23 | 上海新跃仪表厂 | Ground testing system for active vibration abatement of flexible spacecraft |
CN105843237A (en) * | 2016-03-22 | 2016-08-10 | 北京航空航天大学 | Spacecraft attitude reference instruction generation method for suppressing flexible vibration |
Non-Patent Citations (2)
Title |
---|
BING LIL,ETC: "Vibration Suppression of a 3-PRR Flexible Parallel Manipulator Using Input Shaping", 《PROCEEDINGS OFTHE 2009 IEEE INTERNATIONAL CONFERENCE ON MECHATRONICS AND AUTOMATION 》 * |
姚雨晗,等: "基于输入成型法的空间站变构型过程挠性振动抑制策略", 《宇宙学报》 * |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN107807657A (en) * | 2017-11-29 | 2018-03-16 | 南京理工大学 | A kind of Flexible Spacecraft self-adaptation control method based on path planning |
CN107807657B (en) * | 2017-11-29 | 2021-01-26 | 南京理工大学 | Flexible spacecraft attitude self-adaptive control method based on path planning |
CN108958275A (en) * | 2018-06-25 | 2018-12-07 | 南京理工大学 | A kind of hard and soft liquid coupled system attitude controller and motor-driven path combined optimization method |
CN108958275B (en) * | 2018-06-25 | 2023-09-26 | 南京理工大学 | Rigid-flexible liquid coupling system attitude controller and maneuvering path joint optimization method |
Also Published As
Publication number | Publication date |
---|---|
CN106444815B (en) | 2019-01-22 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
CN103412491B (en) | A kind of Spacecraft feature axis attitude maneuver index time-varying sliding-mode control | |
CN102636139B (en) | Six-displacement-sensor dynamic measuring method of space six-degree-of-freedom movement | |
CN103204193B (en) | A kind of under-actuated bipod robot ambulation control method | |
CN104133479B (en) | A kind of employing single-axle air bearing table simulates the method for testing of flexible satellite three-axis attitude coupled motions | |
CN104567930A (en) | Transfer alignment method capable of estimating and compensating wing deflection deformation | |
CN107703742A (en) | A kind of flexible spacecraft sensor fault adjusting method | |
US11872701B2 (en) | Total centroid state estimation method, humanoid robot and computer readable storage medium using the same | |
CN104723341A (en) | Positioning control method for flexibility joint mechanical arm based on connection and damping configuration | |
Banerjee et al. | Efficient simulation of large overall motion of beams undergoing large deflection | |
Hu et al. | Dynamics and vibration suppression of space structures with control moment gyroscopes | |
Zhiyu et al. | Identification of the state-space model and payload mass parameter of a flexible space manipulator using a recursive subspace tracking method | |
CN105956348A (en) | Spacecraft dynamics modeling method | |
CN103786901B (en) | A kind of method improving Spacecraft Attitude Control performance and vibration-isolating platform | |
CN104656447A (en) | Differential geometry nonlinear control method for aircraft anti-interference attitude tracking | |
CN111098299B (en) | Method for compensating disturbance of space robot | |
CN106249616A (en) | One server mechanical arm dynamic modeling method and system in-orbit | |
CN105912007A (en) | Differential geometry nonlinear control method of spatial mechanical arm anti-interference attitude stabilization | |
CN114779799B (en) | Flexible spacecraft attitude and orbit tracking control method based on expansion disturbance observer | |
CN106444815A (en) | Input forming control method of single-shaft maneuverable spacecraft | |
CN108427272A (en) | Flexible Spacecraft control based on flexible mode observation and vibration suppressing method | |
CN105022402A (en) | Method for determining shortest time of coupled rigid-body spacecraft rapid maneuver | |
CN109634111A (en) | A kind of high-speed overload Robotic Dynamic Method for Calculating Deformation | |
CN109815637A (en) | A kind of emulation mode calculating complete soft Manipulator Dynamics response | |
CN107908855A (en) | The modeling method of Coupled Rigid-flexible space ribbon cable system system | |
CN109084757A (en) | A kind of movement of aircraft wing couples velocity error calculation method with dynamic deformation |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
C06 | Publication | ||
PB01 | Publication | ||
C10 | Entry into substantive examination | ||
SE01 | Entry into force of request for substantive examination | ||
GR01 | Patent grant | ||
GR01 | Patent grant |