CN112099519B - Rapid singularity avoidance planning method for spacecraft control moment gyroscope - Google Patents
Rapid singularity avoidance planning method for spacecraft control moment gyroscope Download PDFInfo
- Publication number
- CN112099519B CN112099519B CN202011006501.2A CN202011006501A CN112099519B CN 112099519 B CN112099519 B CN 112099519B CN 202011006501 A CN202011006501 A CN 202011006501A CN 112099519 B CN112099519 B CN 112099519B
- Authority
- CN
- China
- Prior art keywords
- control moment
- moment gyro
- control
- zero motion
- singularity
- 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 40
- 230000033001 locomotion Effects 0.000 claims abstract description 110
- 229920006395 saturated elastomer Polymers 0.000 claims description 12
- 239000011159 matrix material Substances 0.000 claims description 7
- 238000005259 measurement Methods 0.000 claims description 6
- 238000009434 installation Methods 0.000 claims description 3
- 239000004576 sand Substances 0.000 claims description 3
- 230000026058 directional locomotion Effects 0.000 claims 1
- 238000010586 diagram Methods 0.000 description 4
- 230000003321 amplification Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000003199 nucleic acid amplification method Methods 0.000 description 1
Images
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
- G05D1/0833—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using limited authority control
Landscapes
- Engineering & Computer Science (AREA)
- Computer Security & Cryptography (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The invention discloses a fast singularity avoidance planning method for a spacecraft control moment gyro, and belongs to the field of spacecraft attitude planning and control of the control moment gyro. According to the instruction moment of the current control system, the saturation singularity of the current control moment gyro is calculated, the zero motion guide direction of the control moment gyro in the saturation state is calculated, the control moment gyro can keep larger angular momentum, meanwhile, the zero motion direction of the control moment gyro is considered, the singularity avoidance zero motion direction is considered, the zero motion total direction of the control moment gyro is calculated, the magnitude of the zero motion of the control moment gyro is designed, the value of the zero motion is increased when the control moment gyro is close to the singularity state, the zero motion equation of the control moment gyro is obtained, the zero motion equation and the pseudo-inverse control law are combined, the singularity avoidance control law of the control moment gyro is obtained, and therefore the fast singularity avoidance planning of the spacecraft control moment gyro is achieved. The invention has the advantages of high efficiency, high speed, strong safety and high accuracy.
Description
Technical Field
The invention relates to a fast singularity avoidance planning method for a spacecraft control moment gyro, which is suitable for a control moment gyro control law and belongs to the field of spacecraft attitude planning and control moment gyro control.
Background
With the increasing demand for space exploration, space missions that need to be completed by spacecraft are increasingly difficult. These tasks generally require the spacecraft to have a large angular attitude maneuver capability, and therefore place higher demands on the attitude actuator. The control moment gyroscope has a better torque amplification function and is often used for an actuating mechanism of a spacecraft large-angle attitude maneuver. The control moment gyro is used as an actuating mechanism of the spacecraft, a control moment gyro planning algorithm needs to be developed, and a control law of the control moment gyro is designed in advance. And if the control moment gyroscope falls into a singular state, the command moment of the spacecraft attitude control system cannot be output.
Most of the existing spacecraft control moment gyro control laws adopt a gradient method or a robust control method to solve the problem of singularity avoidance of the control moment gyro, and the control moment gyro is ensured to be far away from a singularity state. However, the method does not utilize the saturation characteristic of the control moment gyro, the angular momentum of the control moment gyro is kept small all the time, and the singularity avoidance efficiency is low. The saturated state of the control moment gyro is fully utilized, so that the control moment gyro can keep larger angular momentum, and the utilization rate of the control moment gyro by a control system is improved.
Disclosure of Invention
Aiming at the problem of singularity avoidance of the control moment gyro, the invention discloses a rapid singularity avoidance planning method of a spacecraft control moment gyro, which aims to solve the technical problems that: in a spacecraft attitude maneuver task, a control moment gyro output instruction moment is used for designing a control moment gyro singularity avoiding method, the control moment gyro is guided to approach to a saturation state, the control moment gyro keeps larger angular momentum, the utilization rate of a control system to the control moment gyro is improved, singularity constraint is quickly avoided and instruction moment is output by designing a control moment gyro singularity avoiding method, and the safety and singularity avoiding efficiency of a spacecraft control system are improved. The invention has the advantages of high efficiency, high speed, strong safety and high accuracy.
The purpose of the invention is realized by the following technical scheme:
the invention discloses a rapid singularity avoidance planning method of a spacecraft control moment gyro, which comprises the steps of calculating the saturation singularity of the current control moment gyro according to the instruction moment of the current control system, the zero motion guide direction of the control moment gyro in the saturation state is calculated according to the saturation singularity, so that the control moment gyro can keep larger angular momentum, meanwhile, the zero motion direction of the control moment gyro is considered, the singular avoidance zero motion direction is considered, the total zero motion direction of the control moment gyro is calculated, the magnitude of the zero motion of the control moment gyro is designed, the value of the zero motion becomes large when the control moment gyro is close to a singular state, the zero motion equation of the control moment gyro is obtained, the zero motion equation and the pseudo-inverse control law are combined, the singular avoidance control law of the control moment gyro is obtained, and therefore the rapid singular avoidance planning of the spacecraft control moment gyro is achieved.
The invention discloses a rapid singularity avoidance planning method of a spacecraft control moment gyroscope, which comprises the following steps:
the method comprises the following steps: and calculating the saturation singularity of the current control moment gyro according to the command moment of the current control system.
Calculating the saturation singularity delta of the current control moment gyro according to the command moment u of the current control systemb。
The angular momentum H expression of the control moment gyro is as follows:
H＝h0(Asinδ+Bcosδ)E (1)
wherein h is0Is the angular momentum of the individual control moment gyro, a and B are the installation configuration matrices, E is the unit vector of dimension n, δ is the current frame angle combination.
C expression of a moment matrix of the control moment gyro:
C＝Acosδ-Bsinδ (2)
according to the command torque u of the current control system, in a singular state:
u·C＝0 (3)
substituting the formulas (1) and (2) into the formula (3) according to the corresponding conditions of the saturated singularity and singularity to obtain the saturated singularity delta of the front control moment gyrob：
δb＝arctan(uA/uB) (4)
Step two: and calculating the zero motion guide direction of the control moment gyro in the saturation state based on the saturation singularity of the control moment gyro obtained in the step one. By guiding the control moment gyro frame to rotate towards the saturation singular point, the control moment gyro can keep larger angular momentum, and the utilization rate of the control moment gyro by a control system is improved.
Saturated singularity delta of control moment gyro obtained based on step onebCalculating the zero motion guide direction delta of the control moment gyro in the saturation stateΔ. Angular saturation singularity delta of gyro frame by guiding control momentbAnd the control moment gyro can keep larger angular momentum by rotating, so that the utilization rate of the control system to the control moment gyro is improved.
Make full use of control torqueThe saturation state of the gyroscope guides the angular saturation singularity approach of the control moment gyroscope frame, and the frame angle difference delta towards the saturation singularityΔAs the direction of zero movement of the control moment gyro.
δΔ＝δb-δ (5)
Controlling the orientation delta of a moment gyroΔThe control moment gyro keeps larger angular momentum by moving in the zero movement direction, and the utilization rate of the control system to the control moment gyro is improved.
Step three: and (4) based on the zero motion direction of the control moment gyro in the step two, considering the singularity avoidance zero motion direction, and calculating the zero motion total direction of the control moment gyro, namely, by considering the singularity avoidance of the control moment gyro, the singularity constraint can be avoided and the instruction moment is output, so that the safety and the accuracy of the control system are improved.
Zero motion direction delta based on step two control moment gyroscopeΔZero direction of motion delta taking into account singular avoidancesCalculating the total direction delta of zero motion of the control moment gyrozNamely, by considering the singularity avoidance of the control moment gyro, the singularity constraint can be avoided and the instruction moment can be output, so that the safety and the accuracy of the control system are improved.
The parameters for evaluating the singular state of the control moment gyroscope are singular measurement m:
m＝det(CCT) (6)
the zero motion direction for controlling the moment gyro to avoid singularity is the far direction delta of the singularity metric ms：
Zero motion direction delta for gyroscopic singularity avoidance according to control momentsAnd controlling the zero motion direction delta of the moment gyro towards the saturation singularityΔTo obtain the total zero motion direction delta of the control moment gyroz：
δz＝k1δs+k2δz (8)
Wherein k is1And k2For the scale factor, the moment gyro is controlled along deltazThe direction is rotated, so that the moment gyro can be controlled to avoid singularity, and larger angular momentum can be kept, and the safety and the accuracy of the control moment gyro are further improved.
Step four: and designing the magnitude of the zero motion of the control moment gyro based on the total direction of the zero motion of the control moment gyro obtained in the step three to obtain a zero motion equation of the control moment gyro. By considering the magnitude of zero motion of the control moment gyro, the control moment gyro is quickly avoided near the singular state, the safety of the control moment gyro is improved, and the rapidity and the high efficiency of the control moment gyro singular avoidance planning method are improved.
Zero motion total direction delta of control moment gyroscope obtained based on step threezDesigning the magnitude alpha of the zero motion of the control moment gyro to obtain the zero motion equation of the control moment gyro
By considering the magnitude of zero motion of the control moment gyro, the control moment gyro is quickly avoided near the singular state, the safety of the control moment gyro is improved, and the rapidity and the high efficiency of the control moment gyro singular avoidance planning method are improved.
To the total direction of zero movement deltazAnd analyzing, wherein the control moment gyro contains a component close to a saturation singular point and a component avoiding singularity, and the control moment gyro needs to quickly avoid singularity, so that the size alpha of zero motion of the control moment gyro is designed to be larger when the control moment gyro is close to singularity.
α＝k3e-m (9)
Wherein k is3As a proportional coefficient, the closer to the singular state, the larger the value of the zero motion size alpha of the control moment gyro is, and the quick singularity avoidance is facilitated.
Will control the total direction delta of zero motion of the moment gyrozThe zero motion of the control moment gyro is obtained by combining the control moment gyro with the zero motion alphaFang Cheng
Wherein E isnIs an n-order identity matrix, controlling a moment gyro according to a zero motion equation
And (5) rotating the top.
Step five: and combining the zero motion equation and the pseudo-inverse manipulation law of the control moment gyroscope obtained in the step four to obtain the manipulation law of the control moment gyroscope. The control moment gyro rotates according to the manipulation law from the initial frame angle, and therefore the control moment gyro can rapidly perform singular avoidance planning. The control moment gyro zero motion equation and the pseudo-inverse control law are combined to obtain the control law of the control moment gyro, the singularity of the control moment gyro is avoided quickly, and a large angular momentum is kept.
The control moment gyro zero motion equation obtained in the step four
And pseudo-inverse law of operation
The combination is carried out to obtain the control law of the rapid singular evasion planning algorithm of the control moment gyro
The control moment gyro rotates according to the manipulation law from the initial frame angle, and therefore the control moment gyro can rapidly perform singular avoidance planning.
By considering the singularity avoidance of the control moment gyro, the control moment gyro can avoid singularity constraint and output instruction moment, and the safety and the accuracy of the control moment gyro are improved. The control moment gyro zero motion equation and the pseudo-inverse control law are combined to obtain the control law of the control moment gyro, the control moment gyro can rapidly avoid singularity and keep larger angular momentum, and the control moment gyro has the advantages of high efficiency, high speed, strong safety and high accuracy.
Has the advantages that:
1. according to the rapid singularity avoidance planning method for the spacecraft control moment gyroscope, disclosed by the invention, the control moment gyroscope can keep larger angular momentum by guiding the angular saturation singularity rotation of the control moment gyroscope frame, and the utilization rate of a control system to the control moment gyroscope is improved.
2. According to the rapid singularity avoidance planning method for the spacecraft control moment gyro, the singularity avoidance of the control moment gyro is considered, the control moment gyro can avoid singularity constraint and output a command moment, and the safety and the accuracy of a control system are improved.
3. According to the rapid singularity avoidance planning method for the spacecraft control moment gyroscope, disclosed by the invention, the control moment gyroscope can rapidly avoid near a singularity state by considering the magnitude of zero motion of the control moment gyroscope, so that the safety of the control moment gyroscope is improved, and the rapidity and the high efficiency of a planning algorithm are improved.
4. The invention discloses a rapid singularity avoidance planning method of a spacecraft control moment gyro, which combines a zero motion equation and a pseudo-inverse control law of the control moment gyro to obtain the control law of the control moment gyro. The method has good application prospect in the field of spacecraft attitude planning and control moment gyro control.
Drawings
FIG. 1 is a general flow chart of a method for fast singular avoidance planning of a spacecraft control moment gyro;
fig. 2 is a diagram of singularity measurement of a control moment gyro obtained by a fast singularity avoidance planning method of a spacecraft control moment gyro.
Fig. 3 is a schematic diagram of angular rotation of a control moment gyro frame obtained by a fast singularity avoidance planning method of a spacecraft control moment gyro.
Detailed Description
For a better understanding of the objects and advantages of the present invention, reference should be made to the following detailed description taken in conjunction with the accompanying drawings and examples.
In order to verify the feasibility of the method, the control moment gyroscope is set to be in a pentagonal pyramid structure, and the initial frame angles of the control moment gyroscope are [0.39,2.06,2.16,2.70,1.73 and 0.22 ]]T。
As shown in fig. 1, the method for fast singular avoidance planning of a spacecraft control moment gyro disclosed in this embodiment includes the following steps:
the method comprises the following steps: and calculating the saturation singularity of the current control moment gyro according to the command moment of the current control system.
Calculating the saturation singularity delta of the current control moment gyro according to the command moment u of the current control systemb。
The angular momentum H expression of the control moment gyro is as follows:
H＝h0(Asinδ+Bcosδ)E (1)
wherein h is01000N · m · s is the angular momentum of the individual control moment gyro, a and B are the installation configuration matrix of the pentagonal pyramid control moment gyro, E is the unit vector of 6 dimensions, δ is the current frame angle combination, δ is [0.39,2.06,2.16,2.70,1.73,0.22 ] for the initial state]T。
C expression of a moment matrix of the control moment gyro:
C＝Acosδ-Bsinδ (2)
according to the command torque u of the current control system, in a singular state:
u·C＝0 (3)
substituting the formulas (1) and (2) into the formula (3) according to the corresponding conditions of the saturated singularity and singularity to obtain the saturated singularity delta of the front control moment gyrob：
δb＝arctan(uA/uB) (4)
Step two: and calculating the zero motion guide direction of the control moment gyro in the saturation state based on the saturation singularity of the control moment gyro obtained in the step one. By guiding the control moment gyro frame to rotate towards the saturation singular point, the control moment gyro can keep larger angular momentum, and the utilization rate of the control moment gyro by a control system is improved.
Saturated singularity delta of control moment gyro obtained based on step onebCalculating the zero motion guide direction delta of the control moment gyro in the saturation stateΔ. Angular saturation singularity delta of gyro frame by guiding control momentbAnd the control moment gyro can keep larger angular momentum by rotating, so that the utilization rate of the control system to the control moment gyro is improved.
The saturation state of the control moment gyro is fully utilized to guide the frame angle of the control moment gyro to approach to a saturation singular point, and the frame angle difference delta facing the saturation singular pointΔAs the direction of zero movement of the control moment gyro.
δΔ＝δb-δ (5)
Controlling the orientation delta of a moment gyroΔThe control moment gyro keeps larger angular momentum by moving in the zero movement direction, and the utilization rate of the control system to the control moment gyro is improved.
Step three: and (4) based on the zero motion direction of the control moment gyro in the step two, considering the singularity avoidance zero motion direction, and calculating the zero motion total direction of the control moment gyro, namely, by considering the singularity avoidance of the control moment gyro, the singularity constraint can be avoided and the instruction moment is output, so that the safety and the accuracy of the control system are improved.
Zero motion direction delta based on step two control moment gyroscopeΔZero direction of motion delta taking into account singular avoidancesCalculating the total direction delta of zero motion of the control moment gyrozNamely, by considering the singularity avoidance of the control moment gyro, the singularity constraint can be avoided and the instruction moment can be output, so that the safety and the accuracy of the control system are improved.
And evaluating the parameter of the control moment gyroscope in the singular state as singular measurement m, and when the control force singular measurement m is greater than 0, indicating that the control moment gyroscope is not in the singular state.
m＝det(CCT) (6)
The zero motion direction for controlling the moment gyro to avoid singularity is the far direction delta of the singularity metric ms：
Zero motion direction delta for gyroscopic singularity avoidance according to control momentsAnd controlling the zero motion direction delta of the moment gyro towards the saturation singularityΔTo obtain the total zero motion direction delta of the control moment gyroz：
δz＝k1δs+k2δz (8)
Wherein k is1And k2Taking 2 and 1 respectively as proportionality coefficients, and controlling moment gyro along deltazThe direction rotation can control the moment gyro to avoid singularity and keep larger angular momentum, thereby improving the safety and the accuracy of the control moment gyro.
Step four: and designing the magnitude of the zero motion of the control moment gyro based on the total direction of the zero motion of the control moment gyro obtained in the step three to obtain a zero motion equation of the control moment gyro. By considering the magnitude of zero motion of the control moment gyro, the control moment gyro is quickly avoided near the singular state, the safety of the control moment gyro is improved, and the rapidity and the high efficiency of the control moment gyro singular avoidance planning method are improved.
Zero motion total direction delta of control moment gyroscope obtained based on step threezDesigning the zero motion alpha of the control moment gyro to obtain the control momentZero equation of motion for a gyroscope
By considering the magnitude of zero motion of the control moment gyro, the control moment gyro is quickly avoided near the singular state, the safety of the control moment gyro is improved, and the rapidity and the high efficiency of the control moment gyro singular avoidance planning method are improved.
To the total direction of zero movement deltazAnd analyzing, wherein the control moment gyro contains a component close to a saturation singular point and a component avoiding singularity, and the control moment gyro needs to quickly avoid singularity, so that the size alpha of zero motion of the control moment gyro is designed to be larger when the control moment gyro is close to singularity.
α＝k3e-m (9)
Wherein k is3Is the proportionality coefficient, k3The value of (2) is 2, and the closer to the singular state, the larger the value of the zero motion size alpha of the control moment gyro is, which is beneficial to quickly avoiding singularity.
Will control the total direction delta of zero motion of the moment gyrozThe zero motion equation of the control moment gyro is obtained by combining the control moment gyro with the zero motion magnitude alpha
Wherein E isnIs an n-order identity matrix, controlling a moment gyro according to a zero motion equation
And (5) rotating the top.
Step five: and combining the zero motion equation and the pseudo-inverse manipulation law of the control moment gyroscope obtained in the step four to obtain the manipulation law of the control moment gyroscope. The control moment gyro rotates according to the manipulation law from the initial frame angle, and therefore the control moment gyro can rapidly perform singular avoidance planning. The control moment gyro zero motion equation and the pseudo-inverse control law are combined to obtain the control law of the control moment gyro, the singularity of the control moment gyro is avoided quickly, and a large angular momentum is kept.
The control moment gyro zero motion equation obtained in the step four
And pseudo-inverse law of operation
The combination is carried out to obtain the control law of the rapid singular evasion planning algorithm of the control moment gyro
The control moment gyro rotates from the initial frame angle according to the manipulation law, and the control moment gyro can rapidly and singularly avoid planning. The obtained singularity measurement schematic diagram of the control moment gyro is shown in fig. 2, the control moment gyro can better avoid singularity, the angle rotation schematic diagram of the control moment gyro frame is shown in fig. 3, and the control moment gyro can stably rotate.
By considering the singularity avoidance of the control moment gyro, the control moment gyro can avoid singularity constraint and output instruction moment, and the safety and the accuracy of the control moment gyro are improved. The control moment gyro zero motion equation and the pseudo-inverse control law are combined to obtain the control law of the control moment gyro, the control moment gyro can rapidly avoid singularity and keep larger angular momentum, and the control moment gyro has the advantages of high efficiency, high speed, strong safety and high accuracy.
The above detailed description is intended to illustrate the objects, aspects and advantages of the present invention, and it should be understood that the above detailed description is only exemplary of the present invention and is not intended to limit the scope of the present invention, and any modifications, equivalents, improvements and the like within the spirit and principle of the present invention should be included in the scope of the present invention.
Claims (6)
1. A fast singularity avoidance planning method for a spacecraft control moment gyro is characterized by comprising the following steps: comprises the following steps of (a) carrying out,
the method comprises the following steps: calculating the saturation singularity of the current control moment gyro according to the instruction moment of the current control system;
step two: calculating the zero motion guide direction of the control moment gyro in a saturated state based on the obtained saturated singular point of the control moment gyro; the control moment gyro can keep larger angular momentum by guiding the control moment gyro frame to rotate towards the saturation singular point;
step three: based on the zero motion guide direction of the control moment gyro in the step two, the total zero motion direction of the control moment gyro is calculated by considering the singular avoiding zero motion direction;
step four: designing the magnitude of the zero motion of the control moment gyroscope based on the total direction of the zero motion of the control moment gyroscope obtained in the step three to obtain a zero motion equation of the control moment gyroscope;
step five: combining the zero motion equation and the pseudo-inverse manipulation law of the control moment gyroscope obtained in the step four to obtain the manipulation law of the control moment gyroscope; the control moment gyro rotates from the initial frame angle according to the manipulation law of the control moment gyro, and therefore the control moment gyro can rapidly and singularly avoid planning.
2. The method for rapid singularity avoidance planning for a spacecraft controlled moment gyroscope of claim 1, wherein: the first implementation method comprises the following steps of,
calculating the saturation singularity delta of the current control moment gyro according to the command moment u of the current control systemb；
The angular momentum H expression of the control moment gyro is as follows:
H＝h0(A sinδ+B cosδ)E (1)
wherein h is0The angular momentum of a single control moment gyroscope, A and B are installation configuration matrixes, E is a unit vector of n dimensions, and delta is the current frame angle combination;
c expression of a moment matrix of the control moment gyro:
C＝A cosδ-B sinδ (2)
according to the command torque u of the current control system, in a singular state:
u·C＝0 (3)
substituting the formula (2) into the formula (3) according to the corresponding conditions of the saturated singularity and singularity to obtain the saturated singularity delta of the control moment gyrob：
δb＝arctan(uA/uB) (4)。
3. The method for rapid singularity avoidance planning for a spacecraft controlled moment gyroscope of claim 2, wherein: the second step is realized by the method that,
saturated singularity delta of control moment gyro obtained based on step onebCalculating the zero motion guide direction delta of the control moment gyro in the saturation stateΔ；
The saturation state of the control moment gyro is fully utilized to guide the angular saturation singularity of the control moment gyro frame to approach, and the zero motion guide direction delta of the control moment gyro in the saturation state is generatedΔ；
δΔ＝δb-δ (5)
Along deltaΔThe directional movement enables the control moment gyro to keep larger angular momentum, and the utilization rate of the control moment gyro by a control system is improved.
4. The method for rapid singularity avoidance planning for a spacecraft controlled moment gyroscope of claim 3, wherein: the third step is to realize the method as follows,
zero motion guide direction delta of control moment gyro based on step two saturation statesΔZero direction of motion delta taking into account singular avoidancesCalculating the total direction delta of zero motion of the control moment gyroz；
The parameters for evaluating the singular state of the control moment gyroscope are singular measurement m:
m＝det(CCT) (6)
generating a singularly evaded zero-motion direction delta from a singularity metric ms：
Zero motion direction delta for gyroscopic singularity avoidance according to control momentsAnd zero motion guiding direction delta of control moment gyroΔTo obtain the total zero motion direction delta of the control moment gyroz：
δz＝k1δs+k2δz (8)
Wherein k is1And k2Is a scaling factor.
5. The method for rapid singularity avoidance planning for a spacecraft controlled moment gyroscope of claim 4, wherein: the implementation method of the fourth step is that,
zero motion total direction delta of control moment gyroscope obtained based on step threezDesigning the magnitude alpha of the zero motion of the control moment gyro to obtain the zero motion equation of the control moment gyro
For the total direction delta of zero movement of the control moment gyrozAnalyzing, wherein the control moment gyro comprises a component close to a saturation singular point and a component avoiding singularity, and the control moment gyro needs to quickly avoid singularity, so that the size alpha of zero motion of the control moment gyro is designed to be larger when the control moment gyro is close to singularity;
α＝k3e-m (9)
wherein k is3Is a proportion systemCounting;
will control the total direction delta of zero motion of the moment gyrozThe zero motion equation of the control moment gyro is obtained by combining the control moment gyro with the zero motion magnitude alpha
Wherein E isnIs an n-order identity matrix, controlling a moment gyro according to a zero motion equation
And (5) rotating the top.
6. The method for rapid singularity avoidance planning for a spacecraft controlled moment gyroscope of claim 5, wherein: the fifth step is to realize that the method is that,
the control moment gyro zero motion equation obtained in the step four
And pseudo-inverse law of operation
The combination is carried out to obtain the control law of the rapid singular evasion planning algorithm of the control moment gyro
The control moment gyro rotates from the initial frame angle according to the manipulation law of the control moment gyro, and therefore the control moment gyro can rapidly and singularly avoid planning.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011006501.2A CN112099519B (en) | 2020-09-23 | 2020-09-23 | Rapid singularity avoidance planning method for spacecraft control moment gyroscope |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011006501.2A CN112099519B (en) | 2020-09-23 | 2020-09-23 | Rapid singularity avoidance planning method for spacecraft control moment gyroscope |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN112099519A CN112099519A (en) | 2020-12-18 |
| CN112099519B true CN112099519B (en) | 2021-09-03 |
Family
ID=73755892
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011006501.2A Active CN112099519B (en) | 2020-09-23 | 2020-09-23 | Rapid singularity avoidance planning method for spacecraft control moment gyroscope |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112099519B (en) |
Family Cites Families (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| KR101474274B1 (en) * | 2013-03-25 | 2014-12-18 | 한국항공우주연구원 | Control moment gyroscope |
| CN103941741B (en) * | 2014-04-28 | 2016-06-01 | 北京控制工程研究所 | Based on the control moment gyro frame corners speed control method for determination of amount of zero motion |
| CN105223961B (en) * | 2015-10-16 | 2018-04-13 | 北京机械设备研究所 | It is a kind of to be used for the unusual Spacecraft Attitude Control method evaded of control-moment gyro |
| CN105235917B (en) * | 2015-10-26 | 2017-06-06 | 上海新跃仪表厂 | A kind of control-moment gyro group for tending to target framework angle manipulates rule implementation method |
| CN105388902B (en) * | 2015-11-30 | 2018-04-27 | 北京控制工程研究所 | A kind of unusual bypassing method of control-moment gyro based on instruction Torque vector control |
| CN109871025B (en) * | 2019-02-28 | 2020-09-18 | 北京控制工程研究所 | Method for designing control law of variable-speed control moment gyroscope |
| CN110609565B (en) * | 2019-08-23 | 2023-03-31 | 上海航天控制技术研究所 | Error analysis and precision evaluation method for control moment gyro system |
| CN110658837B (en) * | 2019-09-19 | 2020-10-23 | 北京控制工程研究所 | Steady reconstruction method under control moment gyro fault condition |
-
2020
- 2020-09-23 CN CN202011006501.2A patent/CN112099519B/en active Active
Also Published As
| Publication number | Publication date |
|---|---|
| CN112099519A (en) | 2020-12-18 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Chwa | Sliding-mode-control-based robust finite-time antisway tracking control of 3-D overhead cranes | |
| Zheng et al. | Finite-time path following control for a stratospheric airship with input saturation and error constraint | |
| Shtessel et al. | Guidance and control of missile interceptor using second-order sliding modes | |
| Chwa et al. | Adaptive nonlinear guidance law considering control loop dynamics | |
| JP4783366B2 (en) | Method and system for optimizing torque in a CMG array | |
| CN103092208A (en) | Spacecraft high-accuracy speediness attitude maneuver method based on single gimbal control moment gyro (SGCMG) and reaction wheel (RW) | |
| JP2008189235A (en) | Method, system and program for controlling attitude change, and program recording medium | |
| CN111624878B (en) | Integral sliding mode acquisition method and system for autonomous water surface robot trajectory tracking | |
| Wang et al. | Terminal guidance for a hypersonic vehicle with impact time control | |
| Shao et al. | Model‐assisted extended state observer and dynamic surface control–based trajectory tracking for quadrotors via output‐feedback mechanism | |
| CN105629732A (en) | Spacecraft attitude output feedback tracking control method giving consideration to limited control | |
| Romagnoli et al. | High performance two degrees of freedom attitude control for solar sails | |
| CN110926278B (en) | Preposition guiding method adopting multi-preposition-angle superposition and tail end correction | |
| Schaub et al. | Redundant reaction wheel torque distribution yielding instantaneous l2 power-optimal spacecraft attitude control | |
| CN112099519B (en) | Rapid singularity avoidance planning method for spacecraft control moment gyroscope | |
| Mondal et al. | Constrained quasi-spectral MPSP with application to high-precision missile guidance with path constraints | |
| US9567112B1 (en) | Method and apparatus for singularity avoidance for control moment gyroscope (CMG) systems without using null motion | |
| Jiao et al. | Autonomous homing of parafoil and payload system based on ADRC | |
| CN102749846B (en) | Design method of double parallel configuration VSDGCMGs singularity avoidance steering law | |
| CN113280800B (en) | Equivalent analysis method for angular momentum envelope frame of magnetically suspended control sensitive gyroscope | |
| Özkan et al. | Performance comparison of the notable acceleration-and angle-based guidance laws for a short-range air-to-surface missile | |
| Su et al. | A novel biased proportional navigation guidance law for close approach phase | |
| CN109445448B (en) | Self-adaptive integral sliding-mode attitude controller for wheel-controlled minisatellite | |
| Li et al. | Null‐space‐based optimal control reallocation for spacecraft stabilization under input saturation | |
| Sagliano et al. | Booster dispersion area management through aerodynamic guidance and control |
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 |