Target positioning method suitable for pointing satellite by ground survey station antenna
阅读说明：本技术 适用于地面测站天线对卫星指向的目标定位方法 (Target positioning method suitable for pointing satellite by ground survey station antenna ) 是由 吕旺 郑峰 司力琼 顾强 俞航 王田野 徐晔 杨珺 于 2020-03-19 设计创作，主要内容包括：本发明提供了一种适用于地面测站天线对卫星指向的目标定位方法,通过给定目标时刻的卫星星历数据,经过卫星轨道的相关计算以及相关坐标系的转换计算,得到卫星在地固系下的位置,从而完成卫星的定位过程。本发明考虑卫星实际运行情况计算地面测站天线与卫星的位置关系,不依赖于仿真软件或过多假设内容,有效解决了用于地面站接收天线对卫星指向的的卫星定位问题,而且达到了比较高的定位精度。(The invention provides a target positioning method suitable for a ground survey station antenna to point a satellite. The invention takes the actual running condition of the satellite into consideration to calculate the position relation between the ground station antenna and the satellite, does not depend on simulation software or excessive assumed contents, effectively solves the problem of satellite positioning for the ground station receiving antenna to point to the satellite, and achieves higher positioning precision.)
1. A target positioning method suitable for a ground observation station antenna to point to a satellite is characterized in that the position of the satellite under a terrestrial fixation system is obtained through satellite ephemeris data of a given target moment and through satellite orbit correlation calculation and conversion calculation of a correlation coordinate system, and therefore the positioning process of the satellite is completed.
2. The method of claim 1, comprising the steps of:
calculating the position of the inertial system satellite: according to a given target time t1Computing t from ephemeris information of1Position R of time satellite under inertial systemwECI；
Calculating a target time second counting value: according to a given target time t1Calculating a second count value t of epoch J2000.0 to a predetermined target timec；
And a step of calculating the position of the earth-fixed satellite: according to the epoch J2000.0 obtained by calculation to the second counting value t of the given target timecCalculating t1Position R of time satellite under earth's fixationwECF。
3. The method for locating an object pointed to by a satellite via an antenna of a ground station as claimed in claim 2, wherein the step of calculating the position of the inertial system satellite calculates t1Position R of time satellite under inertial systemwECIThe method comprises the following steps:
inputting a target time t1Includes: the device comprises a track semi-major axis a, a track eccentricity e, a track inclination angle i, a rising intersection declination omega, an argument omega of a near place and an average and near point angle M;
calculating t1Position R of time satellite under inertial systemwECI：
RwECI＝Q*rp
Wherein the rotation matrix Q is described in a 3-1-3 rotation order:
vector rp：
Wherein M is1True proximal angle:
。
4. the method as claimed in claim 2, wherein the step of calculating the second counting value at the target time comprises calculating epoch J2000.0 to the second counting value t at the given target timecThe method comprises the following steps:
input t1Year of the moment, month, day, hour, min, sec, calculate julian day JD:
wherein, floor () is a round-down operation;
calculating a second count value t from epoch J2000.0 to a given target time based on the julian day JDc：
tc＝(JD-2455197.5)×86400+315547200。
5. The method for positioning an object by pointing an antenna of a ground station to a satellite according to claim 2, wherein the calculation t in the step of calculating the position of the earth-fixed satellite is1Position R of time satellite under earth's fixationwECFThe method comprises the following steps:
calculating an earth rotation matrix ER, a nutation matrix NR and a precision matrix PR;
calculating a conversion matrix M from an inertia system to a ground-fixed system according to the calculated earth rotation matrix ER, nutation matrix NR and time difference matrix PRECI2ECF；
T is obtained according to the calculation1Position R of time satellite under inertial systemwECICalculating t1Position R of time satellite under earth's fixationwECF。
6. The method for positioning the target pointed by the ground station antenna to the satellite according to claim 5, wherein the method for calculating the earth rotation matrix ER, the nutation matrix NR and the time matrix PR is as follows:
epoch J2000.0 (second counting value t to given target time) obtained from the calculationcCalculating a yellow meridian nutation delta psi, a yellow-red intersection angle and an intersection angle nutation delta:
wherein, T2kRelative epoch J2000.0, julian century:
the earth rotation matrix ER calculation method comprises the following steps:
calculating the declination nutation delta mu:
Δμ＝Δψ*cos
calculating Greenwich mean time
Calculating greenwich mean time SG：
Calculating an earth rotation matrix ER:
the nutation matrix NR calculation method comprises the following steps:
NR＝RX(--Δ)RZ(-Δψ)RX()
wherein the content of the first and second substances,
the calculation method of the age matrix PR comprises the following steps:
calculating the age constant ζA、θA、ZA：
Calculating a time offset matrix PR:
PR＝RZ(-ZA)RY(θA)RZ(-ζA)
wherein the content of the first and second substances,
。
7. the method of claim 5 applied to groundThe method for positioning the target pointed by the satellite by the surface station antenna is characterized in that a conversion matrix M from an inertia system to a ground-fixed system is calculated according to a calculated earth rotation matrix ER, a nutation matrix NR and a time-of-flight matrix PRECI2ECFThe calculation method of (2) is as follows:
MECI2ECF＝ER*NR*PR 。
8. the method of claim 5 wherein the t is calculated based on the computed position of the target1Position R of time satellite under inertial systemwECICalculating t1Position R of time satellite under earth's fixationwECFThe calculation method of (2) is as follows:
RwECF＝MECI2ECF*RwECI。
Technical Field
The invention relates to the field of satellite orbit calculation and control, in particular to a target positioning method suitable for pointing a satellite by a ground survey station antenna.
Background
The radar is applied to a plurality of important fields such as guidance and beyond-the-horizon detection, plays a very important role in scientific and technological construction, and with the requirements of detection and control of an outer space target, the tracking and searching capabilities of a ground observation station radar antenna provide higher and higher requirements, because the signal-to-noise ratio of link transmission information is reduced due to the satellite-ground pointing deviation, and even a signal loss phenomenon occurs if the maximum station position tolerance is exceeded. This requires that the positioning of the satellites be more rapid and accurate. Therefore, the satellite positioning calculation method with high design precision has general practical significance.
The positioning calculation of the satellite mainly calculates the position of the satellite under an inertial system according to the orbit information and the time information of the satellite, and further calculates the position of the satellite under a geostationary system. The existing related algorithm research in China mostly focuses on the optimization design of the orientation of a satellite to a ground survey station under the condition that the position of the ground survey station is fixed, and the research on the orientation positioning of a ground survey station antenna to the satellite is less. Aiming at the actual situation, the invention provides a satellite real-time positioning method with higher precision for a ground survey station antenna, and the target positioning pointed by the ground survey station antenna to a satellite is realized.
The patent "a method for controlling the pointing direction of a dual-axis antenna to the ground around a moon satellite" (patent number: CN101204994A) describes a method for calculating the pointing direction of a satellite to the earth center around a moon satellite, which estimates the position of the satellite according to ephemeris data on the ground, calculates the visible area of the satellite to the earth, and calculates the pointing angle of the dual-axis antenna. The patent refers to the field of 'teaching or training'. The invention is different from the method in that the calculation is mainly carried out by combining the relative coordinate systems of the earth and the earth surface position, and the pointing positioning calculation of the ground station antenna to the satellite is completed.
The patent "design method of deep space probe antenna pointing" (patent number: CN105184002A) describes a method of pointing a deep space probe antenna to the ground center, which is used to realize the deep space probe antenna to the ground center orientation. The patent is directed to the orientation of the antenna to the geocenter and not to the given position of the earth surface, and the pointing vector of the detector antenna to the geocenter is directly given, and no algorithm for calculating the satellite position through the orbit parameters exists. The invention differs from the method in that a calculation process for calculating the position of the geostationary satellite by the satellite orbit parameters at a given moment is designed.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a target positioning method suitable for pointing a satellite by a ground survey station antenna.
According to the target positioning method suitable for the ground observation station antenna to point to the satellite, the position of the satellite under the earth-fixed system is obtained through the satellite ephemeris data of the given target moment through the correlation calculation of the satellite orbit and the conversion calculation of the correlation coordinate system, and therefore the positioning process of the satellite is completed.
Preferably, the method comprises the following steps:
calculating the position of the inertial system satellite: according to a given target time t1Computing t from ephemeris information of1Position R of time satellite under inertial systemwECI；
Calculating a target time second counting value: according to a given target time t1Calculating a second count value t of epoch J2000.0 to a predetermined target timec；
And a step of calculating the position of the earth-fixed satellite: according to the epoch J2000.0 obtained by calculation to the second counting value t of the given target timecCalculating t1Position R of time satellite under earth's fixationwECF。
Preferably, the calculation t in the inertial system satellite position calculation step1Position R of time satellite under inertial systemwECIThe method comprises the following steps:
inputting a target time t1Includes: the device comprises a track semi-major axis a, a track eccentricity e, a track inclination angle i, a rising intersection declination omega, an argument omega of a near place and an average and near point angle M;
calculating t1Position R of time satellite under inertial systemwECI：
RwECI＝Q*rp
Wherein the rotation matrix Q is described in a 3-1-3 rotation order:
vector rp：
Wherein M is1True proximal angle:
preferably, the epoch J2000.0 is calculated in the target time second counting value calculating step to give the second counting value t of the target timecThe method comprises the following steps:
input t1Year of the moment, month, day, hour, min, sec, calculate julian day JD:
wherein, floor () is a round-down operation;
calculating a second count value t from epoch J2000.0 to a given target time based on the julian day JDc：
tc＝(JD-2455197.5)×86400+315547200
Preferably, the calculation t in the geostationary satellite position calculation step1Position R of time satellite under earth's fixationwECFThe method comprises the following steps:
calculating an earth rotation matrix ER, a nutation matrix NR and a precision matrix PR;
calculating a conversion matrix M from an inertia system to a ground-fixed system according to the calculated earth rotation matrix ER, nutation matrix NR and time difference matrix PRECI2ECF；
T is obtained according to the calculation1Position R of time satellite under inertial systemwECICalculating t1Position R of time satellite under earth's fixationwECF。
Preferably, the method for calculating the earth rotation matrix ER, the nutation matrix NR and the time offset matrix PR is as follows:
epoch J2000.0 (second counting value t to given target time) obtained from the calculationcCalculating a yellow meridian nutation delta psi, a yellow-red intersection angle and an intersection angle nutation delta:
wherein, T2kRelative epoch J2000.0, julian century:
the earth rotation matrix ER calculation method comprises the following steps:
calculating the declination nutation delta mu:
Δμ＝Δψ*cos
calculating Greenwich mean time
Calculating greenwich mean time SG：
Calculating an earth rotation matrix ER:
the nutation matrix NR calculation method comprises the following steps:
NR＝RX(--Δ)RZ(-Δψ)RX()
wherein the content of the first and second substances,
the calculation method of the age matrix PR comprises the following steps:
calculating the age constant ζA、θA、ZA：
Calculating a time offset matrix PR:
PR＝RZ(-ZA)RY(θA)RZ(-ζA)
wherein the content of the first and second substances,
preferably, the conversion matrix M from the inertia system to the earth-fixed system is calculated according to the calculated earth rotation matrix ER, the nutation matrix NR and the time difference matrix PRECI2ECFThe calculation method of (2) is as follows:
MECI2ECF＝ER*NR*PR
preferably, said t is calculated from1Position R of time satellite under inertial systemwECICalculating t1Position R of time satellite under earth's fixationwECFThe calculation method of (2) is as follows:
RwECF＝MECI2ECF*RwECI
compared with the prior art, the invention has the following beneficial effects:
the invention takes the actual running condition of the satellite into consideration to calculate the position relation between the ground station antenna and the satellite, does not depend on simulation software or excessive assumed contents, effectively solves the problem of satellite positioning for the ground station receiving antenna to point to the satellite, and achieves higher positioning precision.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
fig. 1 is a schematic flow chart of a method for positioning a target pointed by a satellite by a ground observation station antenna according to the present invention.
Fig. 2 is a schematic diagram of the positions of the ground station antenna and the satellite according to the present invention.
Fig. 3 is a schematic diagram of a variation curve of a satellite position under an inertial system according to the present invention.
Fig. 4 is a schematic diagram of a change curve of a satellite position under a geostationary system according to the present invention.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
According to the target positioning method suitable for the ground observation station antenna to point to the satellite, the position of the satellite under the earth-fixed system is obtained through the satellite ephemeris data of the given target moment through the correlation calculation of the satellite orbit and the conversion calculation of the correlation coordinate system, and therefore the positioning process of the satellite is completed.
Specifically, the method comprises the following steps:
calculating the position of the inertial system satellite: according to a given target time t1Computing t from ephemeris information of1Position R of time satellite under inertial systemwECI；
Calculating a target time second counting value: according to a given target time t1Calculating a second count value t of epoch J2000.0 to a predetermined target timec；
And a step of calculating the position of the earth-fixed satellite: according to the epoch J2000.0 obtained by calculation to the second counting value t of the given target timecCalculating t1Position R of time satellite under earth's fixationwECF。
Specifically, the calculation t in the inertial system satellite position calculation step1Position R of time satellite under inertial systemwECIThe method comprises the following steps:
inputting a target time t1Includes: the device comprises a track semi-major axis a, a track eccentricity e, a track inclination angle i, a rising intersection declination omega, an argument omega of a near place and an average and near point angle M;
calculating t1Position R of time satellite under inertial systemwECI：
RwECI＝Q*rp
Wherein the rotation matrix Q is described in a 3-1-3 rotation order:
vector rp：
Wherein M is1True proximal angle:
specifically, the epoch J2000.0 is calculated in the target time second counting value calculation step to give the second counting value t of the target timecThe method comprises the following steps:
input t1Year of the moment, month, day, hour, min, sec, calculate julian day JD:
wherein, floor () is a round-down operation;
calculating a second count value t from epoch J2000.0 to a given target time based on the julian day JDc：
tc＝(JD-2455197.5)×86400+315547200
Specifically, the calculation t in the geostationary satellite position calculation step1Position R of time satellite under earth's fixationwECFThe method comprises the following steps:
calculating an earth rotation matrix ER, a nutation matrix NR and a precision matrix PR;
calculating a conversion matrix M from an inertia system to a ground-fixed system according to the calculated earth rotation matrix ER, nutation matrix NR and time difference matrix PRECI2ECF；
T is obtained according to the calculation1Position R of time satellite under inertial systemwECICalculating t1Time guardPosition R of the star under the Earth's systemwECF。
Specifically, the method for calculating the earth rotation matrix ER, the nutation matrix NR and the time matrix PR is as follows:
epoch J2000.0 (second counting value t to given target time) obtained from the calculationcCalculating a yellow meridian nutation delta psi, a yellow-red intersection angle and an intersection angle nutation delta:
wherein, T2kRelative epoch J2000.0, julian century:
the earth rotation matrix ER calculation method comprises the following steps:
calculating the declination nutation delta mu:
Δμ＝Δψ*cos
calculating Greenwich mean time
Calculating greenwich mean time SG：
Calculating an earth rotation matrix ER:
the nutation matrix NR calculation method comprises the following steps:
NR＝RX(--Δ)RZ(-Δψ)RX()
wherein the content of the first and second substances,
the calculation method of the age matrix PR comprises the following steps:
calculating the age constant ζA、θA、ZA：
Calculating a time offset matrix PR:
PR＝RZ(-ZA)RY(θA)RZ(-ζA)
wherein the content of the first and second substances,
specifically, the earth rotation matrix ER, the nutation matrix NR and the time difference matrix PR are calculated according to the calculated earth rotation matrix ER, the nutation matrix NR and the time difference matrix PRCalculating transformation matrix M from inertia system to ground fixation systemECI2ECFThe calculation method of (2) is as follows:
MECI2ECF＝ER*NR*PR
in particular, said t is obtained from a calculation1Position R of time satellite under inertial systemwECICalculating t1Position R of time satellite under earth's fixationwECFThe calculation method of (2) is as follows:
RwECF＝MECI2ECF*RwECI
the present invention will be described more specifically below with reference to preferred examples.
Preferred example 1:
the coordinate system required by the invention is as follows: the inertial system is a J2000.0 inertial coordinate system, and the earth fixation system is a WGS-84 coordinate system.
The calculation process of the present invention is detailed below:
the simulation of this algorithm was verified using MAT L AB, with the earth-related parameters and the station center set as described above, and the input simulation parameters for UTC time 2018, 12, 3, 5, and 30 were as follows:
and (3) carrying out simulation positioning calculation on the satellite from 12, 3 and 5 hours and 30 minutes in the UTC time of 2018, wherein the simulation step length is 1s, the simulation is continuously carried out for 30 minutes, and satellite ephemeris data of every 1s are obtained by STK simulation.
(1) According to a given target time (UTC time) t1Computing t from ephemeris information of1Position R of time satellite under inertial systemwECIThe specific calculation process is as follows:
according to t1Includes: the track comprises a semi-major axis a of the track, an eccentricity e of the track, a track inclination angle i, a rising intersection declination omega, an amplitude angle omega at a near place and an average angle M at a near point. Calculating t1Time of dayPosition R of satellite under inertial systemwECIThe obtained position variation curve of the satellite under the inertial system is shown in fig. 3:
RwECI＝Q*rp
wherein the rotation matrix Q is described in a 3-1-3 rotation order:
vector rp：
Wherein M is1True proximal angle:
(2) according to a given target time t1Calculating a second counting value t from epoch J2000.0 (1/12/2000) to a predetermined target timecThe specific calculation process is as follows:
input t1Year (year), month (month), day (day), hour (hour), minute (min), and second (sec) of time (UTC time), julian day JD is calculated:
wherein floor () is a round-down operation.
Calculating a second counting value t from epoch J2000.0 (1 month, 1 day, 12 hours in 2000) to a given target time according to the julian day JDc：
tc＝(JD-2455197.5)×86400+315547200
(3) A second counting value t from epoch J2000.0 (1 month, 1 day, 12 days 2000) to a given target time calculated according to the step (2)cCalculating t1Position R of time satellite under earth's fixationwECFThe specific calculation process is as follows:
calculating an earth rotation matrix ER, a nutation matrix NR and a precession matrix PR, and firstly calculating a yellow meridian nutation delta psi, a yellow-red crossing angle and a crossing angle nutation delta:
wherein, T2kRelative epoch J2000.0 (1 month, 1 day, 12 of 2000):
the earth rotation matrix ER calculation method comprises the following steps:
calculating the declination nutation delta mu:
Δμ＝Δψ*cos
calculating Greenwich mean time
Calculating greenwich mean time SG：
Calculating an earth rotation matrix ER:
the nutation matrix NR calculation method comprises the following steps:
NR＝RX(--Δ)RZ(-Δψ)RX()
wherein the content of the first and second substances,
the calculation method of the age matrix PR comprises the following steps:
calculating the age constant ζA、θA、ZA：
Calculating a time offset matrix PR:
PR＝RZ(-ZA)RY(θA)RZ(-ζA)
wherein the content of the first and second substances,
calculating a conversion matrix M from an inertia system to a ground-fixed system according to the earth rotation matrix ER, the nutation matrix NR and the precision matrix PRECI2ECF：
MECI2ECF＝ER*NR*PR
And calculating t according to the step (1)1Position R of time satellite under inertial systemwECICalculating t1Position R of time satellite under earth's fixationwECFThe obtained satellite position variation curve under the earth-fixed system is shown in fig. 4:
RwECF＝MECI2ECF*RwECI
preferred example 2:
the technical problem to be solved by the invention is as follows: the method is characterized in that satellite ephemeris data at a given target moment and geographical position information of an antenna of the ground survey station are converted into the position of a satellite in a ground fixed system through satellite orbit correlation calculation and conversion calculation of a plurality of correlation coordinate systems, and the satellite positioning process is completed.
The target positioning method suitable for the ground station antenna to point to the satellite calculates the position of the satellite in real time through the satellite ephemeris information, considers the influence factors of the coordinate system conversion relation comprehensively, has high calculation precision and effectively meets the requirement of satellite positioning. The relationship between the ground station antenna and the satellite position is shown in fig. 2.
As shown in fig. 1, a schematic flow chart of a method for positioning a target pointed by a satellite by a ground survey station antenna provided by the present invention is shown, and the steps of the technical solution of the present invention are as follows:
(1) according to a given target time (UTC time) t1Computing t from ephemeris information of1Position R of time satellite under inertial systemwECIThe process comprises the following steps:
inputting a target time t1Includes: the track comprises a semi-major axis a of the track, an eccentricity e of the track, a track inclination angle i, a rising intersection declination omega, an amplitude angle omega at a near place and an average angle M at a near point. Calculating t1Position R of time satellite under inertial systemwECI：
RwECI＝Q*rp
Wherein the rotation matrix Q is described in a 3-1-3 rotation order:
vector rp：
Wherein M is1True proximal angle:
(2) according to a given target time t1Calculating a second counting value t from epoch J2000.0 (1/12/2000) to a predetermined target timecThe process comprises the following steps:
input t1Year (year), month (month), day (day), hour (hour), minute (min), and second (sec) of time (UTC time), julian day JD is calculated:
wherein floor () is a round-down operation.
Calculating a second counting value t from epoch J2000.0 (1 month, 1 day, 12 hours in 2000) to a given target time according to the julian day JDc：
tc＝(JD-2455197.5)×86400+315547200
(3) A second counting value t from epoch J2000.0 (1 month, 1 day, 12 days 2000) to a given target time calculated according to the step (2)cCalculating t1Position R of time satellite under earth's fixationwECFThe process comprises the following steps:
a second counting value t from epoch J2000.0 (1 month, 1 day, 12 days 2000) to a given target time calculated according to the step (2)cCalculating a yellow meridian nutation delta psi, a yellow-red intersection angle and an intersection angle nutation delta:
wherein, T2kRelative epoch J2000.0 (1 month, 1 day, 12 of 2000):
the earth rotation matrix ER calculation method comprises the following steps:
calculating the declination nutation delta mu:
Δμ＝Δψ*cos
calculating Greenwich mean time
Calculating greenwich mean time SG：
Calculating an earth rotation matrix ER:
the nutation matrix NR calculation method comprises the following steps:
NR＝RX(--Δ)RZ(-Δψ)RX()
wherein the content of the first and second substances,
the calculation method of the age matrix PR comprises the following steps:
calculating the age constant ζA、θA、ZA：
Calculating a time offset matrix PR:
PR＝RZ(-ZA)RY(θA)RZ(-ζA)
wherein the content of the first and second substances,
this term is not considered in the present invention, since polar shifts have little effect on the calculation of the conversion matrix. Calculating a conversion matrix M from an inertia system to a ground-fixed system according to the earth rotation matrix ER, the nutation matrix NR and the time difference matrix PRECI2ECF：
MECI2ECF＝ER*NR*PR
And calculating t according to the step (1)1Position R of time satellite under inertial systemwECICalculating t1Position R of time satellite under earth's fixationwECF：RwECF＝MECI2ECF*RwECI
In the description of the present application, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience in describing the present application and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present application.
Those skilled in the art will appreciate that, in addition to implementing the systems, apparatus, and various modules thereof provided by the present invention in purely computer readable program code, the same procedures can be implemented entirely by logically programming method steps such that the systems, apparatus, and various modules thereof are provided in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system, the device and the modules thereof provided by the present invention can be considered as a hardware component, and the modules included in the system, the device and the modules thereof for implementing various programs can also be considered as structures in the hardware component; modules for performing various functions may also be considered to be both software programs for performing the methods and structures within hardware components.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
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