CN113984069B - Satellite light positioning navigation method based on artificial satellite - Google Patents
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Abstract
The invention relates to a satellite light positioning navigation method based on artificial satellites, which comprises the steps of firstly observing three satellites in a space by adopting a star sensor, obtaining the position coordinates of the three observed satellites according to satellite ephemeris, and calculating the relative distance between any two satellites; measuring unit direction vectors of three satellites relative to the star sensor by adopting the star sensor, and calculating the opening angle 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 has the advantages that the star sensor function is expanded, the autonomous positioning is realized on the basis of realizing the autonomous attitude determination in the prior art, no additional equipment is added, no additional space is occupied, and the star sensor has high economical efficiency.
Description
Technical Field
The invention belongs to the field of navigation, and relates to a satellite positioning navigation method based on an artificial satellite.
Background
The 'inertia and starlight' composite navigation mode based on the strapdown inertial measurement combination and the star sensor integrates the advantages of the two navigation modes, and can realize the 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. The traditional starlight navigation is to measure the star by using a star sensor to determine the carrier flight attitude. Although the attitude information with high precision is obtained, the measurement of the carrier position information cannot be realized, and the application of the star light guidance has a great limitation.
Disclosure of Invention
The invention solves the technical problems that: overcomes the defects of the prior art and provides a satellite light positioning navigation method based on a satellite.
The solution of the invention is as follows:
the satellite light positioning navigation method based on the artificial satellite comprises the following steps:
step 1: three satellites in the space are observed by adopting a star sensor, the position coordinates of the three observed satellites are obtained according to 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 adopting the star sensor, and calculating the opening angle of any two satellites relative to the star sensor;
step 3: calculating the relative distance between the star sensor and each satellite according to the relative distance between any two satellites and the opening angle of the relative star sensor;
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, star sensitivity is adoptedThe device observes three satellites T1, T2 and T3 in space, and obtains three satellite position coordinates as T through a satellite ephemeris database established by satellite orbit measurement xyz [1]、T xyz [2]、T xyz [3]，
T xyz [1]＝[x T1 y T1 z T1 ] T
T xyz [2]＝[x T2 y T2 z T2 ] T
T xyz [3]＝[x T3 y T3 z T3 ] T 。
The relative distance between any two satellites in the three satellites is calculated by using the following formula:
p＝|T xyz [2]-T xyz [1]|
q＝|T xyz [3]-T xyz [2]|
r＝|T xyz [1]-T xyz [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:
measuring unit direction vectors of three satellites T1, T2 and T3 relative to star sensor by adopting star sensor
And calculating the opening angle of each two satellites relative to the star sensor:
wherein A is the opening angle of the satellite T1 and the satellite T2 relative to the star sensor, B is the opening angle of the satellite T2 and the satellite T3 relative to the star sensor, C is the opening angle of the satellite T1 and the satellite 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:
a 2 +b 2 -2a*b*cosA＝p 2
b 2 +c 2 -2b*c*cosB＝q 2
c 2 +a 2 -2c*a*cosC＝r 2
wherein a is the relative distance between the star sensor and the satellite T1; b is the relative distance between the star sensor and satellite T2; c is the relative distance between the star sensor and 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 satellite T1 and the satellite T2 relative to the star sensor, B is the opening angle of the satellite T2 and the satellite T3 relative to the star sensor, and C is the opening angle of the satellite T1 and the satellite T3 relative to the star sensor.
The implementation manner of the step 4 is as follows:
star sensor position is denoted as [ x ] c y c z c ] T ，
The relative distance a between the star sensor and the satellite T1 is denoted as
(x c -x T1 ) 2 +(y c -y T1 ) 2 +(z c -z T1 ) 2 ＝a 2
The relative distance b between the star sensor and satellite T2 is denoted as
(x c -x T2 ) 2 +(y c -y T2 ) 2 +(z c -z T2 ) 2 ＝b 2
The relative distance c between the star sensor and satellite T3 is denoted as
(x c -x T3 ) 2 +(y c -y T3 ) 2 +(z c -z T3 ) 2 ＝c 2
Combining the three formulas, and solving to obtain the position [ x ] of the star sensor c y c z c ] T ；
x T1 ,y T1 ,z T1 Is the position coordinate of satellite T1, x T2 ,y T2 ,z T2 Is the position coordinate of satellite T2, x T3 ,y T3 ,z T3 Is the position coordinates of satellite T3.
The star sensor is installed in a strapdown manner with the aircraft.
When N satellites in the space can be observed by the star sensor, N is more than 3, and the principle of selecting three observation satellites is as follows:
three satellites are selected from N satellites which can be observed by a star sensor and share
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 geometry accuracy factor PDOP is calculated using the following formula:
from all the selection schemes, three satellites corresponding to the selection scheme that minimizes the PDOP are selected as the observation satellites.
Average opening angle of three satellites relative to star sensor
Wherein A, B, C is the opening angle of any two satellites relative to the star sensor.
The invention can greatly improve the autonomous navigation precision of the long-time aircraft, does not need additional hardware modification for the aircraft adopting the star light guidance equipment, and autonomously acquires the high-precision position navigation information through observing the satellite. The method has the specific beneficial effects that:
(1) The star sensor function is expanded, the autonomous positioning is realized on the basis of realizing the autonomous attitude determination in the prior art, no additional equipment is added, no additional space is occupied, and the star sensor has high economical efficiency;
(2) The autonomous positioning is realized based on the observation of the artificial satellite, the interference immunity is strong, the reliability is high, and the autonomous positioning navigation system is a high-precision and totally new autonomous positioning navigation scheme.
Drawings
Fig. 1 is a schematic diagram of satellite positioning navigation based on artificial satellites in accordance with the present invention.
Detailed Description
The invention is further elucidated below in connection with the accompanying drawings.
According to the invention, a plurality of satellites are observed by using a satellite sensor, and the relative distance between the satellites is calculated by combining satellite orbit data provided by satellite ephemeris; measuring unit direction vectors of the satellites relative to the star sensor by adopting 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 the position of the star sensor according to the satellite orbit data and the relative satellite distance, thereby realizing autonomous positioning.
Fig. 1 shows a satellite positioning navigation schematic diagram based on a satellite. The satellite 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 relative distances p, q and r among satellites are obtained through calculation. And observing each satellite by adopting a star sensor, and calculating to obtain an included angle A, B, C between every two satellites. And (3) establishing a calculation model to obtain distances a, b and c of the star sensor relative to each satellite, so as to position the star sensor.
The method comprises the following specific steps:
step 1: three satellites in the satellite sensor observation space are adopted, the position coordinates of the three observation satellites are obtained according to satellite ephemeris, and the relative distance between the observation satellites is calculated.
Three satellites T1, T2 and T3 in space are observed by adopting a star sensor, and a satellite position coordinate T can be obtained through a satellite ephemeris database established by satellite orbit measurement xyz [1]、T xyz [2]、T xyz [3]。
T xyz [1]＝[x T1 y T1 z T1 ] T
T xyz [2]＝[x T2 y T2 z T2 ] T
T xyz [3]＝[x T3 y T3 z T3 ] T
And calculating the relative distances p, q and r among the three satellites.
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.
Step 2: and measuring unit direction vectors of the three satellites relative to the star sensor by adopting the star sensor, and calculating the opening angle of any two satellites relative to the star sensor.
Measuring unit direction vectors of three satellites T1, T2 and T3 relative to star sensor by adopting star sensor
And calculates the relative opening angle A, B, C of each two satellites relative to the star sensor.
Wherein,,
the unit direction vectors of the satellites T1, T2 and T3 relative to the star sensor are respectively; a is the opening angle of the satellites T1 and T2, 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 denoted as O.
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 opening angle of the relative star sensor.
A mathematical model is built up as follows,
according to the triangle cosine law, in the triangles T1, T2, O,
a 2 +b 2 -2a*b*cosA＝p 2 ………………(3)
similarly, in triangles T2, T3, and O,
b 2 +c 2 -2b*c*cosB＝q 2 ………………(4)
similarly, in triangles T1, T3, and O,
c 2 +a 2 -2c*a*cosC＝r 2 ………………(5)
the combination of (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.
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 installing the star sensor and the aircraft in a strapdown way, so that the autonomous positioning of the aircraft is realized.
Star sensor position is denoted as [ x ] c y c z c ] T 。
The distance between the star sensor and satellite T1 is denoted as
(x c -x T1 ) 2 +(y c -y T1 ) 2 +(z c -z T1 ) 2 ＝a 2 ………………(6)
The distance between the star sensor and satellite T2 is denoted as
(x c -x T2 ) 2 +(y c -y T2 ) 2 +(z c -z T2 ) 2 ＝b 2 ………………(7)
The distance between the star sensor and satellite T3 is denoted as
(x c -x T3 ) 2 +(y c -y T3 ) 2 +(z c -z T3 ) 2 ＝c 2 ………………(8)
Combined type (6) to (8) can solve the position [ x ] of the star sensor c y c z c ] T 。
The star sensor is installed in strapdown with the aircraft, 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 more than 3, and the principle of selecting three observation satellites is as follows:
three satellites are selected from N satellites which can be observed by a star sensor and share
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 geometry accuracy factor PDOP is calculated using the following formula:
from all the selection schemes, three satellites corresponding to the selection scheme that minimizes the PDOP are selected as the observation satellites.
Average opening angle of three satellites relative to star sensor
Wherein A, B, C is the opening angle of any two satellites relative to the star sensor.
The invention relates to an aircraft with autonomous navigation requirements and long-endurance flight, which can realize autonomous navigation positioning by observing artificial satellites by using a star sensor by measuring a plurality of artificial satellites, and is not easy to be interfered by the outside in the whole process by using an autonomous navigation system.
The invention provides an autonomous navigation method for expanding star light guidance into a gesture-determination and positioning dual mode based on a satellite light positioning navigation technology, which realizes equipment multiplexing, is simple and economical, has great significance for long-endurance flying aircrafts with autonomous navigation requirements!
The invention is not described in detail in the field of technical personnel common knowledge.
Claims (9)
1. The satellite light positioning navigation method based on the artificial satellite is characterized by comprising the following steps of:
step 1: three satellites in the space are observed by adopting a star sensor, the position coordinates of the three observed satellites are obtained according to 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 adopting the star sensor, and calculating the opening angle of any two satellites relative to the star sensor;
step 3: calculating the relative distance between the star sensor and each satellite according to the relative distance between any two satellites and the opening angle of the relative star sensor;
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 satellite positioning navigation method according to claim 1, wherein in the step 1, three satellites T1, T2, T3 in space are observed by using a star sensor, and three satellite position coordinates are obtained by using a satellite ephemeris database established by satellite orbit measurement xyz [1]、T xyz [2]、T xyz [3]，
T xyz [1]＝[x T1 y T1 z T1 ] T
T xyz [2]＝[x T2 y T2 z T2 ] T
T xyz [3]＝[x T3 y T3 z T3 ] T 。
3. The satellite based starlight positioning navigation method of claim 2 wherein the relative distance between any two of the three satellites is calculated using the formula:
p＝|T xyz [2]-T xyz [1]|
q＝|T xyz [3]-T xyz [2]|
r＝|T xyz [1]-T xyz [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 implementation manner of step 2 is as follows:
measuring unit direction vectors of three satellites T1, T2 and T3 relative to star sensor by adopting star sensor
And calculating the opening angle of each two satellites relative to the star sensor:
wherein A is the opening angle of the satellite T1 and the satellite T2 relative to the star sensor, B is the opening angle of the satellite T2 and the satellite T3 relative to the star sensor, C is the opening angle of the satellite T1 and the satellite 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 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:
a 2 +b 2 -2a*b*cosA＝p 2
b 2 +c 2 -2b*c*cosB＝q 2
c 2 +a 2 -2c*a*cosC＝r 2
wherein a is the relative distance between the star sensor and the satellite T1; b is the relative distance between the star sensor and satellite T2; c is the relative distance between the star sensor and 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 satellite T1 and the satellite T2 relative to the star sensor, B is the opening angle of the satellite T2 and the satellite T3 relative to the star sensor, and C is the opening angle of the satellite T1 and the satellite T3 relative to the star sensor.
6. The satellite based starlight positioning navigation method according to claim 1, wherein the implementation manner of the step 4 is as follows:
star sensor position is denoted as [ x ] c y c z c ] T ，
The relative distance a between the star sensor and the satellite T1 is denoted as
(x c -x T1 ) 2 +(y c -y T1 ) 2 +(z c -z T1 ) 2 ＝a 2
The relative distance b between the star sensor and satellite T2 is denoted as
(x c -x T2 ) 2 +(y c -y T2 ) 2 +(z c -z T2 ) 2 ＝b 2
The relative distance c between the star sensor and satellite T3 is denoted as
(x c -x T3 ) 2 +(y c -y T3 ) 2 +(z c -z T3 ) 2 ＝c 2
Combining the three formulas, and solving to obtain the position [ x ] of the star sensor c y c z c ] T ；
x T1 ,y T1 ,z T1 Is the position coordinate of satellite T1, x T2 ,y T2 ,z T2 Is the position coordinate of satellite T2, x T3 ,y T3 ,z T3 Is the position coordinates of satellite T3.
7. Satellite based starlight positioning navigation method according to any of claims 1 to 6, wherein the star sensor is mounted strapdown with the aircraft.
8. The satellite positioning navigation method based on artificial satellites according to claim 1, wherein when N satellites in space can be observed by a star sensor, N >3, the principle of selecting three observed satellites is as follows:
three satellites are selected from N satellites which can be observed by a star sensor, and the satellite is C3 in total N 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 geometry accuracy factor PDOP is calculated using the following formula:
from all the selection schemes, three satellites corresponding to the selection scheme that minimizes the PDOP are selected as the observation satellites.
9. The satellite based starlight positioning navigation method of claim 8 wherein the average opening angle of three satellites relative to the star sensor
Wherein A, B, C is the opening angle of any two satellites relative to the star sensor. />
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