CN116659543B - Satellite position and attitude estimation method and device based on remote sensing satellite orbit number - Google Patents
Satellite position and attitude estimation method and device based on remote sensing satellite orbit number Download PDFInfo
- Publication number
- CN116659543B CN116659543B CN202310743591.0A CN202310743591A CN116659543B CN 116659543 B CN116659543 B CN 116659543B CN 202310743591 A CN202310743591 A CN 202310743591A CN 116659543 B CN116659543 B CN 116659543B
- Authority
- CN
- China
- Prior art keywords
- satellite
- coordinate system
- information
- angle
- under
- 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 26
- 239000013598 vector Substances 0.000 claims abstract description 83
- 238000006243 chemical reaction Methods 0.000 claims abstract description 55
- 238000012545 processing Methods 0.000 claims abstract description 44
- 239000007787 solid Substances 0.000 claims abstract description 7
- 239000011159 matrix material Substances 0.000 claims description 41
- 238000004364 calculation method Methods 0.000 claims description 18
- 230000009466 transformation Effects 0.000 claims description 18
- 230000008569 process Effects 0.000 claims description 8
- 230000005484 gravity Effects 0.000 claims description 6
- 230000001174 ascending effect Effects 0.000 claims description 3
- 238000010276 construction Methods 0.000 claims description 3
- 230000008901 benefit Effects 0.000 abstract description 2
- 238000003384 imaging method Methods 0.000 description 6
- 238000005259 measurement Methods 0.000 description 4
- 238000001514 detection method Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 238000004458 analytical method Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 238000013507 mapping Methods 0.000 description 1
- 230000005855 radiation Effects 0.000 description 1
- 230000000630 rising effect Effects 0.000 description 1
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C23/00—Combined instruments indicating more than one navigational value, e.g. for aircraft; Combined measuring devices for measuring two or more variables of movement, e.g. distance, speed or acceleration
Landscapes
- Engineering & Computer Science (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Navigation (AREA)
Abstract
The invention discloses a satellite position and attitude estimation method and device based on the number of remote sensing satellite orbits, wherein the method comprises the following steps: acquiring downlink data sent by a remote sensing satellite; analyzing the downlink data to obtain six-number information of the satellite orbit; calculating the six-number information by using a satellite motion state estimation model to obtain satellite motion vector information under a first inertial coordinate system; and processing the satellite motion vector information under the first inertial coordinate system by using a ground-to-solid conversion processing model to obtain the position information, the speed information and the attitude angle information of the satellite motion under the ground-center fixed coordinate system. The invention obtains the high-precision motion parameter information of satellite operation by calculating the ephemeris information, reduces the load complexity of the platform, improves the operation reliability and the application efficiency of the system, and has obvious potential economic benefit and technical level jump.
Description
Technical Field
The invention relates to the technical field of satellite remote sensing, in particular to a satellite position and attitude estimation method and device based on the number of remote sensing satellite orbits.
Background
The current satellite remote sensing system is required to complete data processing application in the fields of remote sensing and mapping, and in the application, imaging data and position and attitude information of a satellite platform at imaging moment are required. The information needs to be measured and acquired by high-precision instruments such as a star sensor, a gyroscope and the like with the assistance of a GNSS precise orbit determination technology. Meanwhile, the detection data downloaded by the satellite remote sensing system comprises ephemeris data. The ephemeris data is used to provide parameters such as satellite orbit parameters, accurate time scale, operating speed, etc. How to fully utilize the direct measurement results of the non-instruments, and obtain the position and posture results of the satellite platform under the appointed first inertial system or the earth-centered earth-fixed system through conversion calculation, so as to reduce the direct observation data pressure which can be obtained only by the precise position and posture measurement instrument, reduce the process and technical difficulty of the high-precision posture measurement instrument which is assisted to be mounted on the satellite platform, and improve the use efficiency of the satellite remote sensing system, thus being the problem which needs to be urgently solved at present.
Disclosure of Invention
The invention discloses a satellite position and attitude estimation method and device based on the number of remote sensing satellite orbits, aiming at how to acquire satellite attitude information by utilizing ephemeris data contained in detection data downloaded by a satellite remote sensing system.
The invention discloses a satellite position and attitude estimation method based on the number of remote sensing satellite orbits, which comprises the following steps:
S1, acquiring downlink data sent by a remote sensing satellite;
S2, analyzing the downlink data to obtain six-number information of the satellite orbit; the six-root information comprises a semi-long axis a of a satellite orbit, an inclination angle i of the satellite orbit, a near-place radial angle omega of the satellite orbit, an eccentricity e of the satellite orbit, a true near-point angle v of the satellite orbit and an ascending intersection point right angle omega of the satellite orbit;
S3, calculating the six-root information by using a satellite motion state estimation model to obtain satellite motion vector information under a first inertial coordinate system; the satellite motion state estimation model comprises a satellite motion parameter estimation model and a satellite motion parameter conversion model; the satellite motion vector information comprises satellite speed vector information and satellite position vector information; the first inertial coordinate system takes the gravity center of the earth as an origin, the Z axis of the first inertial coordinate system points to the north pole, the X axis of the first inertial coordinate system points to the spring point and is positioned on the equatorial plane, and the Y axis of the first inertial coordinate system is positioned on the equatorial plane and is perpendicular to the X axis;
S4, processing satellite motion vector information under the first inertial coordinate system by using a ground-to-solid conversion processing model to obtain position information, speed information and attitude angle information of satellite motion under a ground-center fixed coordinate system; the earth center fixed coordinate system has an origin at the gravity center of the earth and moves along with the revolution of the earth, a Z axis points to the north pole and rotates along with the rotation of the earth, an X axis points to the direction of the primary meridian and rotates along with the rotation of the earth, a Y axis is positioned at the equatorial plane and is vertical to the X axis, and the Y axis rotates along with the rotation of the earth;
The calculating the six pieces of information by using a satellite motion state estimation model to obtain satellite motion vector information under a first inertial coordinate system comprises the following steps:
Calculating the six-number information by using a satellite motion parameter estimation model to obtain satellite motion vector information under a first satellite orbit coordinate system; the first satellite orbit coordinate system takes a focus of a satellite orbit, which is close to a near place, as an origin, takes a connecting line of the focus, which is close to the place, pointing to the near place as an X axis, takes a satellite orbit plane as an XOY plane, wherein a Y axis is positioned on the XOY plane and is perpendicular to the X axis, and a Z axis points to the angular momentum direction of satellite operation and is perpendicular to the XOY plane.
Calculating satellite position vector information under the first satellite orbit coordinate system by using a satellite motion parameter conversion model to obtain satellite motion vector information under the first inertial coordinate system;
The satellite motion parameter estimation model has the following calculation expression:
wherein (x 1, y1, z 1) represents satellite position vector information in the first satellite orbit coordinate system, (vx 1, vy1, vz 1) represents satellite velocity vector information in the first satellite orbit coordinate system, and u is a gravitational constant;
the expression of the satellite motion parameter conversion model is as follows:
Wherein, Rot 3(-Ω)、Rot1(-i)、Rot3 (-omega) is a first conversion matrix, a second conversion matrix and a third conversion matrix of the satellite motion parameter conversion model respectively; (x 2, y2, z 2) represents satellite position vector information in the first inertial coordinate system, and (vx 2, vy2, vz 2) represents satellite velocity vector information in the first inertial coordinate system;
The processing the satellite motion vector information under the first inertial coordinate system by using the earth-solid conversion processing model to obtain the position information, the speed information and the attitude angle information of the satellite motion under the earth-center fixed coordinate system comprises the following steps:
s41, processing satellite motion vector information under the first inertial coordinate system by using a ground fixed position conversion processing model to obtain position information and speed information of satellite motion under a ground fixed coordinate system;
S42, processing the position information and the speed information of the satellite motion under the geocentric fixed coordinate system by using a geodetic angle conversion processing model to obtain the attitude angle information of the satellite motion under the geodetic fixed coordinate system.
The processing the satellite motion vector information under the first inertial coordinate system by using the earth fixed position conversion processing model to obtain the position information and the speed information of the satellite motion under the earth fixed coordinate system comprises the following steps:
Processing the current time information by using a satellite included angle estimation model to obtain a satellite included angle theta; the satellite included angle theta is an included angle between the meridian where the satellite point at the current time of the satellite is located and the meridian of the spring passing point;
Constructing a ground fixed satellite motion conversion matrix by utilizing the satellite included angle theta; the earth-fixed satellite motion conversion matrix has the expression:
Calculating satellite motion vector information under the first inertial coordinate system by using the earth-fixed satellite motion conversion matrix to obtain position information and speed information of satellite motion under an earth-fixed coordinate system; the calculation expression of the calculation process is as follows:
Where (x 3, y3, z 3) represents position information of satellite motion in the geocentric fixed coordinate system, and (vx 3, vy3, vz 3) represents velocity information of satellite motion in the geocentric fixed coordinate system.
The satellite included angle estimation model has a calculation expression as follows:
θ＝θG(t)+θP，
θG(t)＝θG0(Tu)+ωeΔt，
Tu＝du/36525，
wherein θ G (T) is an angle of rotation of the geocentric fixed coordinate system around the Z axis of the first inertial coordinate system at the current time T, that is, an included angle between the Z axis of the geocentric fixed coordinate at the current time T and the Z axis of the first inertial coordinate system, λ P represents a longitude value of a satellite's point, Δt represents a length of time of the coordinated universal time elapsed since julian days corresponding to the current time T, ω e represents an angular rate of rotation of the earth, ω e＝7.29×10-5rad/sec,θG0(Tu) represents a first angular variable corresponding to T u, T u represents a normalized coordinated universal time day corresponding to the current time T, and d u is a number of coordinated universal time days elapsed since julian days corresponding to the current time T;
The processing the position information and the speed information of the satellite motion under the geocentric fixed coordinate system by using the geodetic angle conversion processing model to obtain the attitude angle information of the satellite motion under the geodetic fixed coordinate system comprises the following steps:
determining the position information of satellite motion under the geocentric fixed coordinate system as a position vector x G, wherein the expression of the position vector x G is as follows:
Determining the speed information of satellite motion under the geocentric fixed coordinate system as a speed vector v G, wherein the expression of the speed vector v G is regarded as:
Constructing and obtaining an angle transformation matrix by using the position vector x G and the speed vector v G
The angle transformation matrixThe construction process of (1) comprises:
Wherein r k, R ω is the angle transformation matrix/>, respectivelyThe angle transformation matrix/>, a second column vector and a third column vectorA matrix of 3 rows and 3 columns;
using the angle transformation matrix Calculating to obtain the attitude angle omega of satellite motion under the earth center fixed coordinate system,And k; ω represents the rotation angle of the X-axis around the geocentric fixed coordinate system while the satellite is in motion; /(I)A rotation angle of a Y-axis around a geocentric fixed coordinate system when the satellite moves; k represents the rotation angle of the Z axis around the geocentric fixed coordinate system while the satellite is in motion; the attitude angle omega,/>And k, the calculation formulas are respectively as follows:
wherein r qj represents an angle transformation matrix Q=1, 2,3, j=1, 2, 3; attitude angles omega,/>, of satellite motion under geocentric fixed coordinate systemAnd k, forming attitude angle information of satellite motion under a geocentric fixed coordinate system.
The invention also discloses a device for estimating the satellite position and the attitude based on the number of remote sensing satellite orbits, which comprises:
a memory storing executable program code;
a processor coupled to the memory;
The processor calls the executable program codes stored in the memory to execute the satellite position and attitude estimation method based on the remote sensing satellite orbit number.
The invention also discloses a computer storage medium which stores computer instructions, and the computer instructions are used for executing the satellite position and attitude estimation method based on the remote sensing satellite orbit number when being called.
The beneficial effects of the invention are as follows:
the method obtains the high-precision motion parameter information of satellite operation by calculating the ephemeris information, is hopeful to replace or reduce the technical requirements of high-precision pose measuring instruments of satellite-borne remote sensing systems, reduces the load complexity of platforms, improves the operation reliability and the application efficiency of the system, and has obvious potential economic benefit and technical level jump.
Drawings
FIG. 1 is a flow chart of the method of the present invention;
FIG. 2 is a schematic view of the rotation angle between ECR and the first inertial frame;
FIG. 3 is a schematic diagram of position and velocity state vectors for computing an attitude rotation matrix under an ECR system.
Detailed Description
For a better understanding of the present disclosure, an embodiment is presented herein.
FIG. 1 is a flow chart of the method of the present invention; FIG. 2 is a schematic view of the rotation angle between ECR and the first inertial frame; FIG. 3 is a schematic diagram of position and velocity state vectors for computing an attitude rotation matrix under an ECR system.
The invention obtains the final expected satellite position vector and pointing gesture under the earth center earth fixed system by utilizing the number of the orbit 6 of the earth remote sensing satellite through a series of calculation processing, thereby providing data support for the earth observation of the subsequent remote sensing satellite.
The invention discloses a satellite position and attitude estimation method based on the number of remote sensing satellite orbits, which comprises the following steps:
S1, acquiring downlink data sent by a remote sensing satellite;
S2, analyzing the downlink data to obtain six-number information of the satellite orbit; the six-root information comprises a semi-long axis a of a satellite orbit, an inclination angle i of the satellite orbit, a near-place radial angle omega of the satellite orbit, an eccentricity e of the satellite orbit, a true near-point angle v of the satellite orbit and an ascending intersection point right angle omega of the satellite orbit; the satellite orbit is a remote sensing satellite orbit; the satellites in the application are all remote sensing satellites. The analysis processing of the downlink data can be realized by adopting professional remote sensing data processing software.
After six number information of the satellite orbit is obtained, error discrimination operation is carried out on the six number information of the satellite orbit, and the expression is as follows:
∣X2-X20∣≤a,
∣X3-X30∣≤b,
∣(X2-X20)(X3-X30)∣≤a2,
Wherein a and b are error thresholds of the dip angle and the near-site argument of the satellite orbit respectively, the satellite angle error area constraint threshold is a2, X20 and X30 are standard values of the dip angle and the near-site argument of the satellite orbit respectively, and X2 and X3 represent the dip angle and the near-site argument of the satellite orbit in six pieces of information of the satellite orbit respectively; if the six-root information of the satellite orbit meets the error discrimination expression, reserving the six-root information; if not, deleting the information and acquiring six pieces of information at the next moment.
S3, calculating the six-root information by using a satellite motion state estimation model to obtain satellite motion vector information under a first inertial coordinate system; the satellite motion state estimation model comprises a satellite motion parameter estimation model and a satellite motion parameter conversion model; the satellite motion vector information comprises satellite speed vector information and satellite position vector information; the first inertial coordinate system takes the gravity center of the earth as an origin, the Z axis of the first inertial coordinate system points to the north pole, the X axis of the first inertial coordinate system points to the spring point and is positioned on the equatorial plane, and the Y axis of the first inertial coordinate system is positioned on the equatorial plane and is perpendicular to the X axis; the X axis, the Y axis and the Z axis meet the right rule; the right rule is a rule followed by a Cartesian coordinate system.
S4, processing satellite motion vector information under the first inertial coordinate system by using a ground-to-solid conversion processing model to obtain position information, speed information and attitude angle information of satellite motion under a ground-center fixed coordinate system; the earth center fixed coordinate system (ECR for short) has an origin at the gravity center of the earth and moves along with the revolution of the earth, a Z axis pointing to the north pole and rotating along with the rotation of the earth, an X axis pointing to the direction of the primary meridian and rotating along with the rotation of the earth, and a Y axis located at the equatorial plane and perpendicular to the X axis and rotating along with the rotation of the earth;
In the six-number information, a represents a semi-major axis of a satellite orbit, and the semi-major axis of the satellite orbit refers to a distance between a near point (a point where the satellite orbit is closest to the earth) and a far point (a point where the satellite orbit is farthest from the earth). i denotes the inclination of the satellite orbit, which is the angle between the plane of the satellite orbit and the equatorial plane, and conventionally is a number between 0 ° and 180 °. ω represents the perigee argument of the satellite orbit, which is the angle between the node line and the semi-major axis, and is the operational angle value from the rising intersection point to the perigee. The nodal line is the intersection between the orbital plane and the equatorial plane. e denotes the eccentricity of the satellite orbit, which is a number between 0 and 1 describing the shape of the satellite elliptical orbit: when e is equal to 0, the ellipse is a circle, and when e is very close to 1, the ellipse is very long and thin, and e is used for calculating the ratio between the focal length and the main axis; v represents the true near point angle of the satellite, which is the running angle value of the satellite from the near point to the current position, and in one rotation, the value of the running angle value is from 0 to 360 degrees; it is defined as 0 ° at the near point and thus 180 ° at the far point. Omega represents the right ascent and descent of the satellite orbit, which is the angle from the earth center to the spring point to the line from the earth center to the right ascent and descent point in the equatorial plane, and the value of the right ascent and descent of the satellite orbit is in the range of 0 DEG to 360 deg.
The calculating the six pieces of information by using a satellite motion state estimation model to obtain satellite motion vector information under a first inertial coordinate system comprises the following steps:
Calculating the six-number information by using a satellite motion parameter estimation model to obtain satellite motion vector information under a first satellite orbit coordinate system; the first satellite orbit coordinate system takes a focus of a satellite orbit, which is close to a near place, as an origin, takes a connecting line of the focus, which is close to the place, pointing to the near place as an X axis, takes a satellite orbit plane as an XOY plane, wherein a Y axis is positioned on the XOY plane and is perpendicular to the X axis, and a Z axis points to the angular momentum direction of satellite operation and is perpendicular to the XOY plane.
Calculating satellite position vector information under the first satellite orbit coordinate system by using a satellite motion parameter conversion model to obtain satellite motion vector information under the first inertial coordinate system;
The satellite motion parameter estimation model has the following calculation expression:
Wherein, (x 1, y1, z 1) represents satellite position vector information in the first satellite orbit coordinate system, (vx 1, vy1, vz 1) represents satellite velocity vector information in the first satellite orbit coordinate system, u is a gravitational constant, and is 398600.4418 ±0.0008km 3s-2;
the expression of the satellite motion parameter conversion model is as follows:
Wherein, Rot 3(-Ω)、Rot1(-i)、Rot3 (-omega) is a first conversion matrix, a second conversion matrix and a third conversion matrix of the satellite motion parameter conversion model respectively; (x 2, y2, z 2) represents satellite position vector information in the first inertial coordinate system, and (vx 2, vy2, vz 2) represents satellite velocity vector information in the first inertial coordinate system;
The processing the satellite motion vector information under the first inertial coordinate system by using the earth-solid conversion processing model to obtain the position information, the speed information and the attitude angle information of the satellite motion under the earth-center fixed coordinate system comprises the following steps:
s41, processing satellite motion vector information under the first inertial coordinate system by using a ground fixed position conversion processing model to obtain position information and speed information of satellite motion under a ground fixed coordinate system;
S42, processing the position information and the speed information of the satellite motion under the geocentric fixed coordinate system by using a geodetic angle conversion processing model to obtain the attitude angle information of the satellite motion under the geodetic fixed coordinate system.
The processing the satellite motion vector information under the first inertial coordinate system by using the earth fixed position conversion processing model to obtain the position information and the speed information of the satellite motion under the earth fixed coordinate system comprises the following steps:
processing the current time information by using a satellite included angle estimation model to obtain a satellite included angle theta; the satellite included angle theta is an included angle between the meridian where the satellite point at the current time of the satellite is located and the meridian of the spring passing point; the current time information can be obtained through time measurement equipment;
Constructing a ground fixed satellite motion conversion matrix by utilizing the satellite included angle theta; the earth-fixed satellite motion conversion matrix has the expression:
Calculating satellite motion vector information under the first inertial coordinate system by using the earth-fixed satellite motion conversion matrix to obtain position information and speed information of satellite motion under an earth-fixed coordinate system; the calculation expression of the calculation process is as follows:
Where (x 3, y3, z 3) represents position information of satellite motion in the geocentric fixed coordinate system, and (vx 3, vy3, vz 3) represents velocity information of satellite motion in the geocentric fixed coordinate system.
The satellite included angle estimation model has a calculation expression as follows:
θ＝θG(t)+λP，
θG(t)＝θG0(Tu)+ωeΔt，
Tu＝du/36525，
Wherein θ G (T) is an angle of rotation of the geocentric fixed coordinate system around the Z axis of the first inertial coordinate system at the current time T, that is, an included angle between the Z axis of the geocentric fixed coordinate at the current time T and the Z axis of the first inertial coordinate system, λ P represents a longitude value of a satellite's point, Δt represents a length of time of the coordinated universal time elapsed since julian days corresponding to the current time T, ω e represents an angular rate of rotation of the earth, ω e＝7.29×10-5rad/sec,θG0(Tu) represents a first angular variable corresponding to T u, T u represents a normalized coordinated universal time day corresponding to the current time T, and d u is a number of coordinated universal time days elapsed since julian days corresponding to the current time T; the julian day, JD 2451545.0 for short, corresponds to epoch 2000, month 1, day 1, 12h UT1 (JR 2000).
The processing the position information and the speed information of the satellite motion under the geocentric fixed coordinate system by using the geodetic angle conversion processing model to obtain the attitude angle information of the satellite motion under the geodetic fixed coordinate system comprises the following steps:
determining the position information of satellite motion under the geocentric fixed coordinate system as a position vector x G, wherein the expression of the position vector x G is as follows:
Determining the speed information of satellite motion under the geocentric fixed coordinate system as a speed vector v G, wherein the expression of the speed vector v G is regarded as:
Constructing and obtaining an angle transformation matrix by using the position vector x G and the speed vector v G
The angle transformation matrixThe construction process of (1) comprises:
Wherein r k, R ω is the angle transformation matrix/>, respectivelyThe angle transformation matrix/>, a second column vector and a third column vectorA matrix of 3 rows and 3 columns;
using the angle transformation matrix Calculating to obtain the attitude angle omega of satellite motion under the earth center fixed coordinate system,And k; ω represents the rotation angle of the X-axis around the geocentric fixed coordinate system while the satellite is in motion; /(I)A rotation angle of a Y-axis around a geocentric fixed coordinate system when the satellite moves; k represents the rotation angle of the Z axis around the geocentric fixed coordinate system while the satellite is in motion; the attitude angle omega,/>And k, the calculation formulas are respectively as follows:
wherein r qj represents an angle transformation matrix Q=1, 2,3, j=1, 2, 3; attitude angles omega,/>, of satellite motion under geocentric fixed coordinate systemAnd k, forming attitude angle information of satellite motion under a geocentric fixed coordinate system.
The invention also discloses a device for estimating the satellite position and the attitude based on the number of remote sensing satellite orbits, which comprises:
a memory storing executable program code;
a processor coupled to the memory;
The processor calls the executable program codes stored in the memory to execute the satellite position and attitude estimation method based on the remote sensing satellite orbit number.
The invention also discloses a computer storage medium which stores computer instructions, and the computer instructions are used for executing the satellite position and attitude estimation method based on the remote sensing satellite orbit number when being called.
In the detection data transmitted by the remote sensing satellite under earth observation, additional metadata is generally provided besides the imaging data of the sensor to explain the imaging geometrical radiation imaging parameters, the imaging data and the satellite operation state of the sensor. The metadata file format has a plurality of forms of specifications such as plain text, HDF (hierarchical data format file), XML (extensible markup language), DIMAP (Digital Image Map) and the like, and the contents necessarily comprise UTC time marks, orbit parameters, state vectors and the like, and the following table lists the metadata content description of internationally common remote sensing satellites.
Table 1 metadata file main content and representation of five common remote sensing satellite systems (abbreviation: rs=reference system; pos=position; vel=speed; att=attitude)
The following uses MISR satellite remote sensor metadata file as an example to describe the position and posture result and effect under the ECR coordinate system calculated by the conversion of the method of the present patent. For simplicity, only 10 state vectors and corresponding conversion calculations in 256 pieces of raw data are given here.
The position coordinates in the first inertial coordinate system can be calculated from the metadata ephemeris of EOS-AM1 of the MISR using the orbit root number as follows
The columns of data are: time (in seconds), X, Y, Z (in meters). After conversion to ECR system, position (X, Y, Z) is
During this conversion, the coordinate system rotation is performed only around the Z-axis. Obviously, the results are also presented below in terms of geographical coordinates (latitude and longitude, height in meters), it being apparent that longitude changes significantly during the conversion process, while latitude and height remain relatively stable.
Position in ECR series of positions in first inertial series
The corresponding pose results are calculated from the position and velocity state vectors in the ECR system as follows
Together with the results in the first inertial frame, the same attitude is
The external orientation value in the first inertial frame is (only the derived attitude angle corresponding to the first piece of data is listed):
in an ECR system, the corresponding results are:
Thus, the position and the attitude of the satellite are obtained.
The foregoing is merely exemplary of the present application and is not intended to limit the present application. Various modifications and variations of the present application will be apparent to those skilled in the art. Any modification, equivalent replacement, improvement, etc. which come within the spirit and principles of the application are to be included in the scope of the claims of the present application.
Claims (2)
1. A satellite position and attitude estimation method based on the number of remote sensing satellite orbits is characterized by comprising the following steps:
S1, acquiring downlink data sent by a remote sensing satellite;
S2, analyzing the downlink data to obtain six-number information of the satellite orbit; the six-root information comprises a semi-long axis a of a satellite orbit, an inclination angle i of the satellite orbit, a near-place radial angle omega of the satellite orbit, an eccentricity e of the satellite orbit, a true near-point angle v of the satellite orbit and an ascending intersection point right angle omega of the satellite orbit;
S3, calculating the six-root information by using a satellite motion state estimation model to obtain satellite motion vector information under a first inertial coordinate system; the satellite motion state estimation model comprises a satellite motion parameter estimation model and a satellite motion parameter conversion model; the satellite motion vector information comprises satellite speed vector information and satellite position vector information; the first inertial coordinate system takes the gravity center of the earth as an origin, the Z axis of the first inertial coordinate system points to the north pole, the X axis of the first inertial coordinate system points to the spring point and is positioned on the equatorial plane, and the Y axis of the first inertial coordinate system is positioned on the equatorial plane and is perpendicular to the X axis;
S4, processing satellite motion vector information under the first inertial coordinate system by using a ground-to-solid conversion processing model to obtain position information, speed information and attitude angle information of satellite motion under a ground-center fixed coordinate system; the earth center fixed coordinate system has an origin at the gravity center of the earth and moves along with the revolution of the earth, a Z axis points to the north pole and rotates along with the rotation of the earth, an X axis points to the direction of the primary meridian and rotates along with the rotation of the earth, a Y axis is positioned at the equatorial plane and is vertical to the X axis, and the Y axis rotates along with the rotation of the earth;
The calculating the six pieces of information by using a satellite motion state estimation model to obtain satellite motion vector information under a first inertial coordinate system comprises the following steps:
calculating the six-number information by using a satellite motion parameter estimation model to obtain satellite motion vector information under a first satellite orbit coordinate system; the first satellite orbit coordinate system takes a focus of a satellite orbit, which is close to a near place, as an origin, takes a connecting line of the focus, which is close to the place, pointing to the near place as an X axis, takes a satellite orbit plane as an XOY plane, wherein the Y axis is positioned on the XOY plane and is perpendicular to the X axis, and the Z axis points to the angular momentum direction of satellite operation and is perpendicular to the XOY plane;
Calculating satellite position vector information under the first satellite orbit coordinate system by using a satellite motion parameter conversion model to obtain satellite motion vector information under the first inertial coordinate system;
The satellite motion parameter estimation model has the following calculation expression:
wherein (x 1, y1, z 1) represents satellite position vector information in the first satellite orbit coordinate system, (vx 1, vy1, vz 1) represents satellite velocity vector information in the first satellite orbit coordinate system, and u is a gravitational constant;
the expression of the satellite motion parameter conversion model is as follows:
Wherein, Rot 3(-Ω)、Rot1(-i)、Rot3 (- ω) is a first conversion matrix, a second conversion matrix, and a third conversion matrix of the satellite motion parameter conversion model, respectively, (x 2, y2, z 2) represents satellite position vector information under the first inertial coordinate system, (vx 2, vy2, vz 2) represents satellite velocity vector information under the first inertial coordinate system, and u is a gravitational constant;
The processing the satellite motion vector information under the first inertial coordinate system by using the earth-solid conversion processing model to obtain the position information, the speed information and the attitude angle information of the satellite motion under the earth-center fixed coordinate system comprises the following steps:
s41, processing satellite motion vector information under the first inertial coordinate system by using a ground fixed position conversion processing model to obtain position information and speed information of satellite motion under a ground fixed coordinate system;
S42, processing the position information and the speed information of the satellite motion under the geocentric fixed coordinate system by using a geodetic angle conversion processing model to obtain attitude angle information of the satellite motion under the geodetic fixed coordinate system;
The processing the satellite motion vector information under the first inertial coordinate system by using the earth fixed position conversion processing model to obtain the position information and the speed information of the satellite motion under the earth fixed coordinate system comprises the following steps:
Processing the current time information by using a satellite included angle estimation model to obtain a satellite included angle theta; the satellite included angle theta is an included angle between the meridian where the satellite point at the current time of the satellite is located and the meridian of the spring passing point;
Constructing a ground fixed satellite motion conversion matrix by utilizing the satellite included angle theta; the earth-fixed satellite motion conversion matrix has the expression:
Calculating satellite motion vector information under the first inertial coordinate system by using the earth-fixed satellite motion conversion matrix to obtain position information and speed information of satellite motion under an earth-fixed coordinate system; the calculation expression of the calculation process is as follows:
Wherein (x 3, y3, z 3) represents position information of satellite motion in the geocentric fixed coordinate system, and (vx 3, vy3, vz 3) represents velocity information of satellite motion in the geocentric fixed coordinate system;
The satellite included angle estimation model has a calculation expression as follows:
θ＝θG(t)+λP，
θG(t)＝θG0(Tu)+ωeΔt，
Tu＝du/36525，
Wherein θ G (T) is a rotation angle of the geocentric fixed coordinate system around the Z-axis of the first inertial coordinate system at the current time T, λ P is a longitude value of a satellite's undersea point, Δt is a time length of coordinated universal time elapsed since julian day corresponding to the current time T, ω e is a rotation angular rate of the earth, ω e＝7.29×10-5rad/sec,θG0(Tu) is a first angle variable corresponding to T u, T u is a normalized coordinated universal time day, and d u is a coordinated universal time day elapsed since julian day corresponding to the current time T;
The processing the position information and the speed information of the satellite motion under the geocentric fixed coordinate system by using the geodetic angle conversion processing model to obtain the attitude angle information of the satellite motion under the geodetic fixed coordinate system comprises the following steps:
determining the position information of satellite motion under the geocentric fixed coordinate system as a position vector x G, wherein the expression of the position vector x G is as follows:
Determining the speed information of satellite motion under the geocentric fixed coordinate system as a speed vector v G, wherein the expression of the speed vector v G is regarded as:
Constructing and obtaining an angle transformation matrix by using the position vector x G and the speed vector v G
The angle transformation matrixThe construction process of (1) comprises:
Wherein r k, R ω is the angle transformation matrix/>, respectivelyThe angle transformation matrix/>, a second column vector and a third column vectorA matrix of 3 rows and 3 columns;
using the angle transformation matrix Calculating to obtain attitude angles omega and/of satellite motion under a geocentric fixed coordinate systemAnd k; ω represents the rotation angle of the X-axis around the geocentric fixed coordinate system while the satellite is in motion; /(I)A rotation angle of a Y-axis around a geocentric fixed coordinate system when the satellite moves; k represents the rotation angle of the Z axis around the geocentric fixed coordinate system while the satellite is in motion; the attitude angle omega,/>And k, the calculation formulas are respectively as follows:
wherein r qj represents an angle transformation matrix Q=1, 2,3, j=1, 2, 3; attitude angles omega,/>, of satellite motion under geocentric fixed coordinate systemAnd k, forming attitude angle information of satellite motion under a geocentric fixed coordinate system.
2. An apparatus for satellite position attitude estimation based on a remote sensing satellite orbit count, the apparatus comprising:
a memory storing executable program code;
a processor coupled to the memory;
The processor invokes the executable program code stored in the memory to perform the satellite position attitude estimation method based on the number of remote sensing satellite orbits as claimed in claim 1.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310743591.0A CN116659543B (en) | 2023-06-21 | 2023-06-21 | Satellite position and attitude estimation method and device based on remote sensing satellite orbit number |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310743591.0A CN116659543B (en) | 2023-06-21 | 2023-06-21 | Satellite position and attitude estimation method and device based on remote sensing satellite orbit number |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN116659543A CN116659543A (en) | 2023-08-29 |
| CN116659543B true CN116659543B (en) | 2024-05-07 |
Family
ID=87717006
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310743591.0A Active CN116659543B (en) | 2023-06-21 | 2023-06-21 | Satellite position and attitude estimation method and device based on remote sensing satellite orbit number |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116659543B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116910107B (en) * | 2023-09-13 | 2023-12-15 | 北京航天驭星科技有限公司 | Track data extraction method, track calculation method, track data extraction device, track data storage medium and track calculation device |
Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108710379A (en) * | 2018-06-14 | 2018-10-26 | 上海卫星工程研究所 | Fixed statellite is imaged Yaw steering angle computational methods |
| CN108958272A (en) * | 2018-06-15 | 2018-12-07 | 上海卫星工程研究所 | Yaw steering method is imaged in fixed statellite |
| CN109781059A (en) * | 2019-01-08 | 2019-05-21 | 上海卫星工程研究所 | Spaceborne spot beam anternma pointing accuracy assessment system over the ground |
| CN110162855A (en) * | 2019-05-10 | 2019-08-23 | 哈尔滨工业大学 | Spin load Dynamic Accuracy Analysis and error distribution method on remote sensing satellite star |
| CN111427002A (en) * | 2020-03-19 | 2020-07-17 | 上海卫星工程研究所 | Azimuth angle calculation method for ground measurement and control antenna pointing satellite |
| CN112966211A (en) * | 2021-02-04 | 2021-06-15 | 上海卫星工程研究所 | Method and system for calculating visual angle under satellite-to-target observation |
| CN113641182A (en) * | 2021-07-27 | 2021-11-12 | 东方红卫星移动通信有限公司 | High-precision aiming and pointing method and system for inter-satellite laser communication system |
| CN113761664A (en) * | 2021-09-15 | 2021-12-07 | 长光卫星技术有限公司 | Flywheel rotating speed optimization method during remote sensing satellite imaging period |
| CN114802818A (en) * | 2022-06-22 | 2022-07-29 | 上海航天空间技术有限公司 | Morning and evening orbit satellite and sun attitude calculation method and guidance method thereof |
| CN115061171A (en) * | 2022-05-07 | 2022-09-16 | 中国空间技术研究院 | Satellite atmospheric perturbation measuring and calculating method and device |
| CN115982938A (en) * | 2022-11-17 | 2023-04-18 | 中国空间技术研究院 | Method and device for calculating satellite attitude pointing accuracy of autonomous maneuvering trajectory planning |
-
2023
- 2023-06-21 CN CN202310743591.0A patent/CN116659543B/en active Active
Patent Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108710379A (en) * | 2018-06-14 | 2018-10-26 | 上海卫星工程研究所 | Fixed statellite is imaged Yaw steering angle computational methods |
| CN108958272A (en) * | 2018-06-15 | 2018-12-07 | 上海卫星工程研究所 | Yaw steering method is imaged in fixed statellite |
| CN109781059A (en) * | 2019-01-08 | 2019-05-21 | 上海卫星工程研究所 | Spaceborne spot beam anternma pointing accuracy assessment system over the ground |
| CN110162855A (en) * | 2019-05-10 | 2019-08-23 | 哈尔滨工业大学 | Spin load Dynamic Accuracy Analysis and error distribution method on remote sensing satellite star |
| CN111427002A (en) * | 2020-03-19 | 2020-07-17 | 上海卫星工程研究所 | Azimuth angle calculation method for ground measurement and control antenna pointing satellite |
| CN112966211A (en) * | 2021-02-04 | 2021-06-15 | 上海卫星工程研究所 | Method and system for calculating visual angle under satellite-to-target observation |
| CN113641182A (en) * | 2021-07-27 | 2021-11-12 | 东方红卫星移动通信有限公司 | High-precision aiming and pointing method and system for inter-satellite laser communication system |
| CN113761664A (en) * | 2021-09-15 | 2021-12-07 | 长光卫星技术有限公司 | Flywheel rotating speed optimization method during remote sensing satellite imaging period |
| CN115061171A (en) * | 2022-05-07 | 2022-09-16 | 中国空间技术研究院 | Satellite atmospheric perturbation measuring and calculating method and device |
| CN114802818A (en) * | 2022-06-22 | 2022-07-29 | 上海航天空间技术有限公司 | Morning and evening orbit satellite and sun attitude calculation method and guidance method thereof |
| CN115982938A (en) * | 2022-11-17 | 2023-04-18 | 中国空间技术研究院 | Method and device for calculating satellite attitude pointing accuracy of autonomous maneuvering trajectory planning |
Non-Patent Citations (3)
| Title |
|---|
| 基于轨道动态模型的卫星天线指向功能测试方法;尹卿;于澎;白力舸;胡帆;方凯;李振东;;航天器工程;20200215(第01期);全文 * |
| 敏捷遥感卫星新型姿态控制精度评估方法;凌琼;董天舒;谭志云;侯锐;矫轲;张雷;赵峭;;航天器工程;20191215(第06期);8-14 * |
| 面向自主运行的深空探测航天器体系结构设计及自主任务规划方法;冯小恩;李玉庆;杨晨;何熊文;徐勇;朱立颖;;控制理论与应用;20191215(第12期);全文 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN116659543A (en) | 2023-08-29 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| JP5068073B2 (en) | Inertial GPS navigation system using input alignment data for inertial system | |
| EP3460399B1 (en) | Methods, apparatuses, and computer programs for estimating the heading of an axis of a rigid body | |
| CN100501331C (en) | Navigation satellite autonomous navigation system and method based on X-ray pulsar | |
| CN103969672B (en) | A kind of multi-satellite system and strapdown inertial navigation system tight integration air navigation aid | |
| Walter et al. | Astrometry of fundamental catalogues: the evolution from optical to radio reference frames | |
| US10309786B2 (en) | Navigational and location determination system | |
| US20080122785A1 (en) | Portable display with improved functionality | |
| EP2557394A1 (en) | Method and system for processing pulse signals within an interital navigation system | |
| CN110221332A (en) | A kind of the dynamic lever arm estimation error and compensation method of vehicle-mounted GNSS/INS integrated navigation | |
| US8837782B1 (en) | Geoposition determination using satellite ephemerides | |
| CN116659543B (en) | Satellite position and attitude estimation method and device based on remote sensing satellite orbit number | |
| CN105371844A (en) | Initialization method for inertial navigation system based on inertial / celestial navigation interdependence | |
| CN112197765B (en) | Method for realizing fine navigation of underwater robot | |
| CN110440827A (en) | A kind of scaling method of parameter error, device and storage medium | |
| US20140249750A1 (en) | Navigational and location determination system | |
| CN106772483A (en) | A kind of data post processing method and device based on CORS systems | |
| Kumar | Integration of inertial navigation system and global positioning system using kalman filtering | |
| US10006770B2 (en) | Remote location determination system | |
| CN112781617B (en) | Error estimation method, integrated navigation processing system and storage medium | |
| EP3748294A1 (en) | Direction finder | |
| Zhou | Low-cost MEMS-INS/GPS integration using nonlinear filtering approaches | |
| Lee et al. | An efficient integrated attitude determination method using partially available doppler measurement under weak GPS environment | |
| Falletti et al. | The Kalman Filter and its Applications in GNSS and INS | |
| Ford et al. | MEMS Inertial on an RTK GPS Receiver: Integration Options and Test Results | |
| Harke et al. | Gyroscope Drift Estimation of a GPS/MEMSINS Smartphone Sensor Integration Navigation System for Kayaking |
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 |