CN106628258B - A kind of satellite spin attitude determination method based on solar vector information - Google Patents
A kind of satellite spin attitude determination method based on solar vector information Download PDFInfo
- Publication number
- CN106628258B CN106628258B CN201610885091.0A CN201610885091A CN106628258B CN 106628258 B CN106628258 B CN 106628258B CN 201610885091 A CN201610885091 A CN 201610885091A CN 106628258 B CN106628258 B CN 106628258B
- Authority
- CN
- China
- Prior art keywords
- solar vector
- satellite
- spin
- sun
- solar
- 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 16
- 238000005259 measurement Methods 0.000 claims description 15
- 230000006340 racemization Effects 0.000 abstract description 3
- 229920006395 saturated elastomer Polymers 0.000 description 3
- 238000004364 calculation method Methods 0.000 description 2
- 235000013399 edible fruits Nutrition 0.000 description 2
- 238000005070 sampling Methods 0.000 description 2
- 230000029777 axis specification Effects 0.000 description 1
- 230000007812 deficiency Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000005484 gravity Effects 0.000 description 1
- 210000001503 joint Anatomy 0.000 description 1
Classifications
-
- 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
-
- 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
- B64G1/245—Attitude control algorithms for spacecraft attitude control
Landscapes
- Engineering & Computer Science (AREA)
- Remote Sensing (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Radar, Positioning & Navigation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Automation & Control Theory (AREA)
- Navigation (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
A kind of satellite spin attitude determination method based on solar vector information, first measures the sun using sun sensor in a time series, obtains the solar vector sequence under celestial body coordinate system.Then, according to the precision of sun sensor, the solar vector for participating in calculating is chosen.The spin axis of satellite and the size of spin angle velocity are finally calculated according to the variation of solar vector.The method of the present invention can be implemented to provide input information for racemization and the precession control for guaranteeing energy security and taking, and can assess measure and provide foundation, have very strong engineering operability and realizability.
Description
Technical field
The invention belongs to satellite gravity anomaly fields, are related to a kind of determination method of satellite spin posture.
Background technique
When satellite in-orbit period, when causing celestial body angular speed excessive due to exception, in order to guarantee the safety of satellite system,
It needs to understand satellite energy state in time；Meanwhile adopting an effective measure in time for safeguards system energy security, it is required to
Obtain the azimuth information of satellite spin axis and solar vector.
Existing method relies primarily on gyro and measures to the angular speed of satellite, and then takes the measures such as jet to satellite
Racemization is carried out, but the range of gyro is limited, under the speed conditions of satellite big angle, gyro can be saturated, so as to cause that can not obtain
Take the angular speed of satellite.
Summary of the invention
Technical problem solved by the present invention is having overcome the deficiencies of the prior art and provide a kind of based on solar vector information
Satellite spin attitude determination method, using change in location of the solar vector under celestial body system, solve celestial body spin axis and
Angular speed size solves in the prior art when the in-orbit exception of satellite causes angular speed is excessive to cause gyro saturation that can not obtain and defend
The problem of star spin information.
The technical solution of the invention is as follows: a kind of satellite spin attitude determination method based on solar vector information, packet
It includes:
(1) the solar vector S described under celestial body coordinate system is obtained using sun sensor measurementb, and remember at one
Between the solar vector sequence that measures in sequence be Sbk, k=1,2,3 ..., N, N is positive integer；
(2) in SbkIn, positive integer m, n, the p for meeting n > m, p > 1 are taken for K, and accordingly extract sun vector measurement value sequence
Four value S in columnbm、Sbm+p、Sbn、Sbn+p, celestial body spin axis is calculatedAnd spin angle velocity size | ω |:
ΔSbmp=Sbm+p-Sbm、ΔSbnp=Sbn+p-Sbn
Wherein Δ t is the time interval in solar vector sequence between latter two first measured value.
The advantages of the present invention over the prior art are that:
(1) the method for the present invention proposes based entirely on solar vector information, avoids in the prior art due to big angular speed
The problem of gyro is saturated and can not know satellite angular velocity information, and can be generalized in the case of gyro free and determine Satellite Angle
Speed；
(2) the method for the present invention, which gives according to the measurement accuracy of sun sensor, chooses measured value calculating satellite angular speed
Strategy, especially can still be able to maintain meter in the lesser situation of satellite angular speed during satellite despun control
Calculate precision.
Detailed description of the invention
Fig. 1 is the flow diagram of the method for the present invention；
Fig. 2 is position view of the spin axis under celestial body system in the embodiment of the present invention；
Fig. 3 is the big logotype of the spin angle velocity in the embodiment of the present invention.
Specific embodiment
As shown in Figure 1, a kind of satellite spin attitude determination method based on solar vector information, includes the following steps:
S1 is obtained using sun sensor measurement in celestial body coordinate system (usual celestial body coordinate system is defined as: origin is in satellite
Mass center, butt joint ring direction are -X direction, and windsurfing direction is Y-direction, and Z-direction meets the right-hand rule) under the solar vector S that describesb,
Measurement obtains solar vector sequence S in a time seriesbk, k=1,2,3 ..., N, N is positive integer.Without loss of generality,
Assuming that the time interval of measurement sequence is isometric, it is Δ t；
S2, in SbkIn, appoint and take continuous three sun vector measurement values, celestial body spin axis direction can be calculated and sat in celestial body
Unit direction vector under mark systemAnd spin angle velocity size | ω |, calculation method is as follows:
Assuming that the corresponding three sun vector measurement values of continuous observation moment T1, T1+ Δ t, T1+2* Δ t are SbT1、SbT2、
SbT3, then:
ΔSb21=SbT2-SbT1、ΔSb32=SbT3-SbT2 (1)
The principle of above three formula from: around one fixing axle in space rotation unit direction vector, rotary shaft
Direction and angular velocity of rotation can be calculated by the unit direction vector variable quantity and change rate.In the present invention, the sun is sweared
In inertial space be in the amount short time it is fixed, the variable quantity of solar vector therefore can be led to as caused by the rotation of satellite
The variable quantity for crossing calculating solar vector obtains the rotary shaft and angular velocity of rotation of satellite.Wherein sun arrow is calculated in formula (1)
The variable quantity of amount, formula (2) calculate the spin axis of satellite according to the variable quantity of solar vector, and formula (3) is according to solar vector
The spin angle velocity size of change rate calculating satellite.
S3 causes the solar vector between double sampling to exist when celestial body spin angle velocity can gradually decrease during racemization
Variation under celestial body coordinate system is also smaller and smaller, due to being influenced by sun sensor measurement accuracy, by what is be spaced twice in succession
Data, which carry out spin axis determination, will to determine that error becomes increasing.
In order to overcome influence of the sun sensor to precision is determined, according to sun sensor measurement error to determining precision
The accuracy requirement of impact analysis and measurement needs to adjust data break according to spin angle velocity size, and strategy is as follows:
Positive integer m, n, p are taken, n > m, p > 1 are met, then takes four value S in sun vector measurement value sequencebm、Sbm+p、Sbn、
Sbn+p, calculation formula is as follows:
ΔSbmp=Sbm+p-Sbm、ΔSbnp=Sbn+p-Sbn (4)
The principle of above three formula is identical as formula (1)~(3), the difference is that the variation meter of solar vector
It counts in.When satellite spin angular speed is larger, solar vector variable quantity is big in the unit time, can be calculated with formula (1), when defending
When star spin angle velocity is smaller, solar vector variable quantity is smaller in the unit time, then the measurement noise of sun sensor is for knot
Fruit is affected, and in order to obtain biggish solar vector variable quantity, needs to elongate the time interval measured twice, this is namely
The difference of formula (4) and formula (1).
Embodiment
By taking certain satellite in orbit as an example, celestial body angular speed is excessive to cause all gyros to be all saturated, and is not used to judgement and defends
Star spin states.The satellite configures two sun sensors, there is effective measured value in the case, is as follows:
It is calculated according to upper table using the method for the present invention.
Firstly, available solar vector at every sampling moment, such as in star 100483763,100483765,
100483767, solar vector can be calculated respectively are as follows:
Secondly, using formula (1)~(3), available satellite spin axis unit direction vector and spin angle velocity are as follows:
| ω |=384.8315 °/s
Finally, doing above-mentioned calculating to the data in table, the calculating knot of satellite spin axis and spin angle velocity can be obtained
Fruit ordered series of numbers, the celestial body spin posture information finally obtained are as shown in Figures 2 and 3.
The content that description in the present invention is not described in detail belongs to the well-known technique of those skilled in the art.
Claims (1)
1. a kind of satellite spin attitude determination method based on solar vector information, characterized by comprising:
(1) the solar vector S described under celestial body coordinate system is obtained using sun sensor measurementb, and remember in a time series
The upper obtained solar vector sequence that measures is Sbk, k=1,2,3 ..., N, N is positive integer；
(2) in SbkIn, positive integer m, n, the p for meeting n > m, p > 1 are taken for K, and accordingly extract in sun vector measurement value sequence
Four value Sbm、Sbm+p、Sbn、Sbn+p, celestial body spin axis is calculatedAnd spin angle velocity size | ω |:
ΔSbmp=Sbm+p-Sbm、ΔSbnp=Sbn+p-Sbn
Wherein Δ t is the time interval in solar vector sequence between latter two first measured value.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610885091.0A CN106628258B (en) | 2016-10-10 | 2016-10-10 | A kind of satellite spin attitude determination method based on solar vector information |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610885091.0A CN106628258B (en) | 2016-10-10 | 2016-10-10 | A kind of satellite spin attitude determination method based on solar vector information |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN106628258A CN106628258A (en) | 2017-05-10 |
| CN106628258B true CN106628258B (en) | 2019-05-24 |
Family
ID=58853795
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201610885091.0A Active CN106628258B (en) | 2016-10-10 | 2016-10-10 | A kind of satellite spin attitude determination method based on solar vector information |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN106628258B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113386979B (en) * | 2021-06-03 | 2022-12-13 | 长光卫星技术股份有限公司 | Data transmission attitude planning method for self-adaptive sun avoidance |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5020744A (en) * | 1990-01-12 | 1991-06-04 | General Electric Company | Method for acquiring three-axis earth pointing attitude for an initially spinning spacecraft |
| EP0731401A2 (en) * | 1995-03-06 | 1996-09-11 | Space Systems / Loral, Inc. | Spacecraft acquisition of orientation by scan of earth sensor field of view |
| CN101586954A (en) * | 2009-05-27 | 2009-11-25 | 北京航空航天大学 | Digital sun sensor for stable-spinning micro/nano satellite |
| CN103072701A (en) * | 2013-01-30 | 2013-05-01 | 北京控制工程研究所 | Racemization control method for under-actuated satellite |
| CN103438886A (en) * | 2013-08-02 | 2013-12-11 | 国家卫星气象中心 | Determination method for attitudes of spinning stabilized meteorological satellite based on coarse-fine attitude relation model |
-
2016
- 2016-10-10 CN CN201610885091.0A patent/CN106628258B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5020744A (en) * | 1990-01-12 | 1991-06-04 | General Electric Company | Method for acquiring three-axis earth pointing attitude for an initially spinning spacecraft |
| EP0731401A2 (en) * | 1995-03-06 | 1996-09-11 | Space Systems / Loral, Inc. | Spacecraft acquisition of orientation by scan of earth sensor field of view |
| CN101586954A (en) * | 2009-05-27 | 2009-11-25 | 北京航空航天大学 | Digital sun sensor for stable-spinning micro/nano satellite |
| CN103072701A (en) * | 2013-01-30 | 2013-05-01 | 北京控制工程研究所 | Racemization control method for under-actuated satellite |
| CN103438886A (en) * | 2013-08-02 | 2013-12-11 | 国家卫星气象中心 | Determination method for attitudes of spinning stabilized meteorological satellite based on coarse-fine attitude relation model |
Also Published As
| Publication number | Publication date |
|---|---|
| CN106628258A (en) | 2017-05-10 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN106482734A (en) | A kind of filtering method for IMU Fusion | |
| CN106915477A (en) | A kind of attitude control method | |
| CN104165638B (en) | Multi-position self-calibration method for biaxial rotating inertial navigation system | |
| CN103940425A (en) | Magnetic-inertial combination strapdown measuring method | |
| CN106153073B (en) | A kind of nonlinear initial alignment method of full posture Strapdown Inertial Navigation System | |
| CN104697521B (en) | A method of high-speed rotary body posture and angular speed are measured using gyro redundancy oblique configuration mode | |
| CN104596543B (en) | Gyroscope combined error coefficient scaling method under a kind of benchmark uncertain condition | |
| CN105588562B (en) | The method of carrier angular movement is isolated in a kind of rotation modulation inertial navigation system | |
| CN103900607B (en) | Rotation type strapdown inertial navigation system transposition method based on inertial system | |
| CN102937450B (en) | A kind of relative attitude defining method based on gyro to measure information | |
| CN100559190C (en) | A kind of method of demarcating that the accelerometer zero drift is carried out at rail | |
| CN103759729A (en) | Initial attitude acquisition method for ground test for soft lunar landing by using SINS (serial inertial navigation system) | |
| CN107228674A (en) | A kind of improved method for star sensor and gyro Federated filter | |
| CN103148854A (en) | Attitude measurement method of micro-electro mechanical system (MEMS) inertial navigation system based on single-shaft forward revolution and reverse revolution | |
| CN106248082B (en) | A kind of aircraft autonomous navigation system and air navigation aid | |
| CN105241319B (en) | A kind of guided cartridge of spin at a high speed real-time alignment methods in the air | |
| CN105180728A (en) | Front data based rapid air alignment method of rotary guided projectiles | |
| CN106767925A (en) | The location parameter of inertial navigation system three identification alignment methods with twin shaft indexing mechanism | |
| CN106628258B (en) | A kind of satellite spin attitude determination method based on solar vector information | |
| CN100498227C (en) | Method for online real time clearing constant value drift of gyroscope of strapdown inertial navigation system | |
| CN104101345B (en) | Multisensor attitude fusion method based on complementary reconstruction technology | |
| CN104931069B (en) | Demarcate the method, apparatus and system of gyroscope | |
| CN103575300B (en) | Rotating modulation-based rocking base inertial navigation system coarse alignment method | |
| CN109029499A (en) | A kind of accelerometer bias iteration optimizing estimation method based on gravity apparent motion model | |
| CN103591960A (en) | Rotating modulation based coarse alignment method of static base inertial navigation system |
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 |