CN113984069A - Satellite positioning navigation method based on artificial satellite - Google Patents
Satellite positioning navigation method based on artificial satellite Download PDFInfo
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- CN113984069A CN113984069A CN202110484592.9A CN202110484592A CN113984069A CN 113984069 A CN113984069 A CN 113984069A CN 202110484592 A CN202110484592 A CN 202110484592A CN 113984069 A CN113984069 A CN 113984069A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
- G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
- G01C21/005—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 with correlation of navigation data from several sources, e.g. map or contour matching
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
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- Y02D—CLIMATE CHANGE MITIGATION TECHNOLOGIES IN INFORMATION AND COMMUNICATION TECHNOLOGIES [ICT], I.E. INFORMATION AND COMMUNICATION TECHNOLOGIES AIMING AT THE REDUCTION OF THEIR OWN ENERGY USE
- Y02D30/00—Reducing energy consumption in communication networks
- Y02D30/70—Reducing energy consumption in communication networks in wireless communication networks
Abstract
The invention relates to a starlight positioning navigation method based on an artificial satellite, which comprises the steps of firstly adopting a star sensor to observe three satellites in a space, obtaining position coordinates of the three observed satellites according to a satellite ephemeris and calculating the relative distance between any two satellites; measuring unit direction vectors of three satellites relative to the star sensor by using the star sensor, and calculating opening angles of any two satellites relative to the star sensor; calculating the relative distance between the star sensor and each satellite; and calculating the position of the star sensor according to the position coordinates of the three satellites and the relative distance between the star sensor and each satellite, namely realizing the autonomous positioning of the aircraft. The star sensor function is expanded, the autonomous positioning is realized on the basis of the conventional autonomous attitude determination, additional equipment is not added, additional space is not occupied, and the star sensor has high economical efficiency.
Description
Technical Field
The invention belongs to the field of navigation, and relates to a starlight positioning navigation method based on an artificial satellite.
Background
The 'inertia + starlight' composite navigation mode based on the strapdown inertia measurement combination and the star sensor integrates the advantages of two navigation modes, and can realize high dynamic and high precision of autonomous navigation. The star sensor is a device which has small volume, light weight and high reliability and works based on the visible light imaging principle. In the traditional starlight navigation, a star sensor is used for measuring a fixed star and determining the flight attitude of a carrier. Although attitude information with high precision is obtained, the measurement of carrier position information cannot be realized, and the application of star light guidance is very limited.
Disclosure of Invention
The technical problem solved by the invention is as follows: the defects of the prior art are overcome, and a satellite positioning navigation method based on an artificial satellite is provided.
The technical scheme of the invention is as follows:
the satellite positioning navigation method based on the artificial satellite comprises the following steps:
step 1: the star sensor is adopted to observe three satellites in the space, the position coordinates of the three observed satellites are obtained according to the satellite ephemeris, and the relative distance between any two satellites is calculated;
step 2: measuring unit direction vectors of three satellites relative to the star sensor by using the star sensor, and calculating opening angles of any two satellites relative to the star sensor;
and step 3: calculating the relative distance between the star sensor and each satellite according to the relative distance between any two satellites and the field angle of the star sensor;
and 4, step 4: and calculating the position of the star sensor according to the position coordinates of the three satellites and the relative distance between the star sensor and each satellite, namely realizing the autonomous positioning of the aircraft.
In the step 1, a star sensor is adopted to observe three satellites T1, T2 and T3 in space, and satellite ephemeris numbers are established through satellite orbit measurementThe database acquires three satellite position coordinates as Txyz[1]、Txyz[2]、Txyz[3]，
Txyz[1]＝[xT1 yT1 zT1]T
Txyz[2]＝[xT2 yT2 zT2]T
Txyz[3]＝[xT3 yT3 zT3]T。
The relative distance between any two of the three satellites is calculated by the following formula:
p＝|Txyz[2]-Txyz[1]|
q＝|Txyz[3]-Txyz[2]|
r＝|Txyz[1]-Txyz[3]|
wherein p is the relative distance between satellite T1 and satellite T2; q is the relative distance between satellite T2 and satellite T3; r is the relative distance between satellite T1 and satellite T3.
The implementation manner of the step 2 is as follows:
the unit direction vector of three satellites T1, T2 and T3 relative to the star sensor is measured by the star sensor
And calculating the field angle of each two satellites relative to the star sensor:
wherein, A is the opening angle of the satellites T1 and T2 relative to the star sensor, B is the opening angle of the satellites T2 and T3 relative to the star sensor, C is the opening angle of the satellites T1 and T3 relative to the star sensor, and the star sensor is marked as O.
The implementation manner of the step 3 is as follows:
the relative distance between the star sensor and each satellite is calculated by the following mathematical model:
a2+b2-2a*b*cos A＝p2
b2+c2-2b*c*cos B＝q2
c2+a2-2c*a*cos C＝r2
wherein a is the relative distance between the star sensor and the satellite T1; b is the relative distance between the star sensor and the satellite T2; c is the relative distance between the star sensor and the satellite T3;
p is the relative distance between satellite T1 and satellite T2; q is the relative distance between satellite T2 and satellite T3; r is the relative distance between satellite T1 and satellite T3;
a is the opening angle of the satellites T1 and T2 relative to the star sensor, B is the opening angle of the satellites T2 and T3 relative to the star sensor, and C is the opening angle of the satellites T1 and T3 relative to the star sensor.
The implementation manner of the step 4 is as follows:
star sensor position is denoted as [ x ]c yc zc]T，
The relative distance a between the star sensor and the satellite T1 is shown as
(xc-xT1)2+(yc-yT1)2+(zc-zT1)2＝a2
The relative distance b between the star sensor and the satellite T2 is shown as
(xc-xT2)2+(yc-yT2)2+(zc-zT2)2＝b2
The relative distance c between the star sensor and the satellite T3 is shown as
(xc-xT3)2+(yc-yT3)2+(zc-zT3)2＝c2
The three formulas are combined to solve to obtain the star sensor position [ x ]c yc zc]T；
xT1,yT1,zT1Is the position coordinate, x, of satellite T1T2,yT2,zT2Is the position coordinate, x, of satellite T2T3,yT3,zT3Is the position coordinates of satellite T3.
And the star sensor is installed with the aircraft in a strapdown manner.
When N satellites in the space can be observed by the star sensor, N is greater than 3, and the principle of selecting three observation satellites is as follows:
selecting three satellites from N satellites capable of being observed by the star sensor, and sharing the same
A group selection scheme;
for each group of selection schemes, calculating the average opening angle D of three satellites relative to the star sensor;
the position geometric dilution of precision PDOP is calculated using the following formula:
and selecting three satellites corresponding to the selection scheme which minimizes the PDOP from all the selection schemes as observation satellites.
Average opening angle of three satellites relative to star sensor
A, B, C are the opening angles of any two satellites relative to the star sensor.
The invention can greatly improve the autonomous navigation precision of the long-time flight aircraft, does not need additional hardware modification for the aircraft adopting the starlight guidance equipment, and autonomously obtains high-precision position navigation information through the observation satellite. The method has the following specific beneficial effects:
(1) the functions of the star sensor are expanded, and on the basis of the traditional realization of autonomous attitude determination, autonomous positioning is realized, no additional equipment is added, no additional space is occupied, and the method has high economical efficiency;
(2) the autonomous positioning is realized based on the observation of the artificial satellite, the anti-interference performance is strong, the reliability is high, and the method is a high-precision brand-new autonomous positioning navigation scheme.
Drawings
FIG. 1 is a schematic diagram of satellite-based star positioning navigation according to the present invention.
Detailed Description
The invention is further elucidated with reference to the drawing.
According to the method, a star sensor is used for observing a plurality of satellites, and the relative distance between the satellites is calculated by combining satellite orbit data provided by a satellite ephemeris; measuring unit direction vectors of each satellite relative to the star sensor by using the star sensor, and calculating to obtain a vector included angle between the satellites; calculating the relative distance between the star sensor and a plurality of satellites; and calculating to obtain the position of the star sensor according to the satellite orbit data and the relative satellite distance, thereby realizing autonomous positioning.
Fig. 1 is a schematic diagram of satellite-based starlight positioning navigation. The star light positioning navigation method based on the artificial satellite realizes high-precision autonomous ranging through the following working steps. The space is not less than 3 satellites T1, T2 and T3, the satellite positions can be obtained through a satellite ephemeris database established by satellite orbit measurement, and the inter-satellite relative distances p, q and r are obtained through calculation. And observing each satellite by using a star sensor, and calculating to obtain an included angle A, B, C between every two satellites. And establishing a calculation model to obtain the distances a, b and c of the star sensor relative to each satellite, thereby positioning the star sensor.
The method comprises the following specific steps:
step 1: and observing three satellites in the space by using the star sensor, obtaining position coordinates of the three observation satellites according to the satellite ephemeris, and calculating the relative distance between the observation satellites.
The star sensor is adopted to observe three satellites T1, T2 and T3 in space, and the satellite position coordinate T can be obtained through a satellite ephemeris database established by satellite orbit measurementxyz[1]、Txyz[2]、Txyz[3]。
Txyz[1]＝[xT1 yT1 zT1]T
Txyz[2]＝[xT2 yT2 zT2]T
Txyz[3]＝[xT3 yT3 zT3]T
And calculating the relative distance p, q and r among the three satellites.
Wherein p is the relative distance between the satellite T1 and the satellite T2; q is the relative distance between satellite T2 and satellite T3; r is the relative distance between satellite T1 and satellite T3.
Step 2: and measuring unit direction vectors of the three satellites relative to the star sensor by using the star sensor, and calculating the opening angles of any two satellites relative to the star sensor.
The unit direction vector of three satellites T1, T2 and T3 relative to the star sensor is measured by the star sensor
And calculates the relative opening angle A, B, C of each two satellites with respect to the star sensor.
Wherein the content of the first and second substances,
unit direction vectors of the satellites T1, T2 and T3 relative to the star sensor respectively; a is the tension of the satellites T1 and T2Angle, B is the opening angle of the satellites T2 and T3, and C is the opening angle of the satellites T1 and T3. The star sensor is marked as O.
And step 3: and calculating the relative distance between the star sensor and each satellite according to the relative distance between any two satellites and the field angle of the relative star sensor.
The following mathematical model is established as follows,
according to the cosine theorem of triangles T1, T2 and O,
a2+b2-2a*b*cos A＝p2………………(3)
similarly, in the triangles T2, T3 and O,
b2+c2-2b*c*cos B＝q2………………(4)
similarly, in the triangles T1, T3 and O,
c2+a2-2c*a*cos C＝r2………………(5)
the joint formulas (3) to (5) can solve a, b and c.
Wherein a is the distance between the star sensor and the satellite T1; b is the distance between the aircraft and satellite T2; c is the distance between the aircraft and satellite T3.
And 4, step 4: and calculating the position of the star sensor according to the positions of the three satellites and the distance between the star sensor and each satellite, and carrying out strapdown installation on the star sensor and the aircraft, thereby realizing the autonomous positioning of the aircraft.
Star sensor position is denoted as [ x ]c yc zc]T。
The distance between the star sensor and the satellite T1 is represented as
(xc-xT1)2+(yc-yT1)2+(zc-zT1)2＝a2………………(6)
The distance between the star sensor and the satellite T2 is represented as
(xc-xT2)2+(yc-yT2)2+(zc-zT2)2＝b2………………(7)
The distance between the star sensor and the satellite T3 is represented as
(xc-xT3)2+(yc-yT3)2+(zc-zT3)2＝c2………………(8)
The united vertical type (6) - (8) can solve the star sensor position [ x [ ]c yc zc]T。
And the star sensor and the aircraft are installed in a strapdown manner, so that the autonomous positioning of the aircraft is realized.
When N satellites in the space can be observed by the star sensor, N is greater than 3, and the principle of selecting three observation satellites is as follows:
selecting three satellites from N satellites capable of being observed by the star sensor, and sharing the same
A group selection scheme;
for each group of selection schemes, calculating the average opening angle D of three satellites relative to the star sensor;
the position geometric dilution of precision PDOP is calculated using the following formula:
and selecting three satellites corresponding to the selection scheme which minimizes the PDOP from all the selection schemes as observation satellites.
Average opening angle of three satellites relative to star sensor
A, B, C are the opening angles of any two satellites relative to the star sensor.
The invention relates to an aircraft with autonomous navigation requirement and long-term flight, which can realize autonomous navigation positioning by observing an artificial satellite by using a star sensor through measuring a plurality of artificial satellites by using the star sensor.
The invention provides a starlight positioning navigation technology based on an artificial satellite, expands starlight guidance into an autonomous navigation method with attitude determination and positioning modes, realizes equipment reuse, is simple and economic, and has great significance to long-endurance flying aircrafts with autonomous navigation requirements!
The invention is not described in detail and is within the knowledge of a person skilled in the art.
Claims (9)
1. The satellite positioning navigation method based on the artificial satellite is characterized by comprising the following steps:
step 1: the star sensor is adopted to observe three satellites in the space, the position coordinates of the three observed satellites are obtained according to the satellite ephemeris, and the relative distance between any two satellites is calculated;
step 2: measuring unit direction vectors of three satellites relative to the star sensor by using the star sensor, and calculating opening angles of any two satellites relative to the star sensor;
and step 3: calculating the relative distance between the star sensor and each satellite according to the relative distance between any two satellites and the field angle of the star sensor;
and 4, step 4: and calculating the position of the star sensor according to the position coordinates of the three satellites and the relative distance between the star sensor and each satellite, namely realizing the autonomous positioning of the aircraft.
2. The method as claimed in claim 1, wherein in step 1, the star sensor is used to observe three satellites in space, T1, T2 and T3, and a satellite ephemeris database established by satellite orbit measurement is used to obtain the position coordinates of the three satellites as Txyz[1]、Txyz[2]、Txyz[3]，
Txyz[1]＝[xT1 yT1 zT1]T
Txyz[2]＝[xT2 yT2 zT2]T
Txyz[3]＝[xT3 yT3 zT3]T。
3. The satellite-based starlight positioning navigation method according to claim 2, wherein the relative distance between any two of the three satellites is calculated by using the following formula:
p＝|Txyz[2]-Txyz[1]|
q＝|Txyz[3]-Txyz[2]|
r＝|Txyz[1]-Txyz[3]|
wherein p is the relative distance between satellite T1 and satellite T2; q is the relative distance between satellite T2 and satellite T3; r is the relative distance between satellite T1 and satellite T3.
4. The satellite-based starlight positioning navigation method according to claim 1, wherein the step 2 is implemented as follows:
the unit direction vector of three satellites T1, T2 and T3 relative to the star sensor is measured by the star sensor
And calculating the field angle of each two satellites relative to the star sensor:
wherein, A is the opening angle of the satellites T1 and T2 relative to the star sensor, B is the opening angle of the satellites T2 and T3 relative to the star sensor, C is the opening angle of the satellites T1 and T3 relative to the star sensor, and the star sensor is marked as O.
5. The satellite-based starlight positioning navigation method according to claim 1, wherein the step 3 is implemented as follows:
the relative distance between the star sensor and each satellite is calculated by the following mathematical model:
a2+b2-2a*b*cosA＝p2
b2+c2-2b*c*cosB＝q2
c2+a2-2c*a*cosC＝r2
wherein a is the relative distance between the star sensor and the satellite T1; b is the relative distance between the star sensor and the satellite T2; c is the relative distance between the star sensor and the satellite T3;
p is the relative distance between satellite T1 and satellite T2; q is the relative distance between satellite T2 and satellite T3; r is the relative distance between satellite T1 and satellite T3;
a is the opening angle of the satellites T1 and T2 relative to the star sensor, B is the opening angle of the satellites T2 and T3 relative to the star sensor, and C is the opening angle of the satellites T1 and T3 relative to the star sensor.
6. The satellite-based starlight positioning navigation method according to claim 1, wherein the step 4 is implemented as follows:
star sensor position is denoted as [ x ]c yc zc]T，
The relative distance a between the star sensor and the satellite T1 is shown as
(xc-xT1)2+(yc-yT1)2+(zc-zT1)2＝a2
The relative distance b between the star sensor and the satellite T2 is shown as
(xc-xT2)2+(yc-yT2)2+(zc-zT2)2＝b2
The relative distance c between the star sensor and the satellite T3 is shown as
(xc-xT3)2+(yc-yT3)2+(zc-zT3)2＝c2
The three formulas are combined to solve to obtain the star sensor position [ x ]c yc zc]T；
xT1,yT1,zT1Is the position coordinate, x, of satellite T1T2,yT2,zT2Is the position coordinate, x, of satellite T2T3,yT3,zT3Is the position coordinates of satellite T3.
7. The satellite based starlight positioning navigation method according to any one of claims 1 to 6, characterized in that the star sensor is installed in a strapdown with an aircraft.
8. The method according to claim 1, wherein when there are N satellites in space that can be observed by the star sensor, N >3, the principle of selecting three observation satellites is as follows:
selecting three satellites from N satellites capable of being observed by the star sensor, and the total number is C3NA group selection scheme;
for each group of selection schemes, calculating the average opening angle D of three satellites relative to the star sensor;
the position geometric dilution of precision PDOP is calculated using the following formula:
and selecting three satellites corresponding to the selection scheme which minimizes the PDOP from all the selection schemes as observation satellites.
9. The satellite based star optical positioning and navigation method according to claim 8, wherein the average field angle of three satellites with respect to the star sensor
A, B, C are the opening angles of any two satellites relative to the star sensor.
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