CN110595486B - High-precision semimajor axis deviation calculation method based on double-star on-orbit telemetry data - Google Patents

High-precision semimajor axis deviation calculation method based on double-star on-orbit telemetry data Download PDF

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CN110595486B
CN110595486B CN201910838243.5A CN201910838243A CN110595486B CN 110595486 B CN110595486 B CN 110595486B CN 201910838243 A CN201910838243 A CN 201910838243A CN 110595486 B CN110595486 B CN 110595486B
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杨盛庆
杜耀珂
王文妍
王禹
刘美师
王嘉轶
陈桦
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Shanghai Aerospace Control Technology Institute
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Abstract

The invention provides a semi-major axis deviation calculation method based on double-star on-orbit telemetering data, which is used for eliminating error data by combining threshold judgment aiming at double-star orbit telemetering data; time synchronization is carried out according to the track telemetering data files acquired by the two stars respectively; processing the double-satellite orbit telemetering data on the basis of time synchronization, establishing a formation coordinate system with a main satellite as an origin, and determining the course deviation of the two satellites; and (4) obtaining the relative position in the tangent plane under the formation coordinate system during dynamic compensation, and carrying out ellipse geometric fitting on the relative position to obtain the semimajor axis deviation. The method can overcome the influence of the track curvature on the determination of the double-star semimajor axis deviation under the condition of long distance at the initial stage of track entry. Compared with the traditional methods such as the method of direct difference of flat roots or estimation of single-point drift variation along the course, the method has the advantage that the determination accuracy of the semi-long axis deviation is obviously improved.

Description

High-precision semimajor axis deviation calculation method based on double-star on-orbit telemetry data
Technical Field
The invention belongs to the field of spacecraft engineering technology application, and relates to a high-precision semimajor axis deviation calculation method based on double-star on-orbit telemetry data.
Background
The dynamics and navigation filtering algorithm of the relative navigation method for on-orbit long-term autonomous formation is mainly based on the relative state measurement technology of the differential GNSS. The technology requires to establish an inter-satellite link, and for the condition that the inter-satellite link is not established by double satellites at the initial stage of the orbit entering, a large satellite-ground loop is required to carry out formation initialization control. The formation initialization mainly controls the phase relation in the double-star orbit plane, and the semimajor axis deviation of the double stars is an important input parameter of the control method. The double-star semimajor axis deviation numerical value of the large satellite-ground loop is quantified and can only be determined by orbit telemetering data downloaded by two stars respectively. At present, for a double-star semimajor axis deviation high-precision determination technology in the working state, no published patent or paper and other research results exist.
Disclosure of Invention
The invention provides a high-precision semimajor axis deviation calculation method based on double-star on-orbit telemetry data. Aiming at the working state that inter-satellite links and relative measurement are not established in the initial stage of orbit entering, high-precision semimajor axis deviation calculation is carried out by using respective orbit telemetering data of double satellites.
The invention provides a high-precision semimajor axis deviation calculation method based on double-star on-orbit telemetry data. According to the double-satellite track telemetering data (GNSS measured data and track number determined by absolute navigation), the threshold value judgment is combined, and the error data are eliminated; time synchronization is carried out on track telemetering data files acquired by the two stars respectively; processing the double-satellite orbit telemetering data on the basis of time synchronization, establishing a formation coordinate system with a main satellite as an origin, and determining the course deviation of the two satellites; performing dynamic compensation on the basis of the course deviation of the two stars along the direction, and obtaining the relative position in the tangent plane under the formation coordinate system after the compensation along the course deviation is obtained; and carrying out ellipse geometric fitting on the basis of the relative position in the tangent plane under the formation coordinate system to obtain the high-precision semimajor axis deviation.
The method can overcome the influence of the track curvature on the determination of the double-star semimajor axis deviation under the condition of long distance at the initial stage of track entry. Compared with the traditional methods such as the method of direct difference of flat roots or estimation of single-point drift variation along the course, the method has the advantage that the determination accuracy of the semi-long axis deviation is obviously improved.
Drawings
Fig. 1 is a schematic flow chart of the semimajor axis deviation calculation method of the present invention.
FIG. 2 is a schematic diagram of ellipse fitting parameter definition of the present invention.
FIG. 3 is the relative position in the tangent plane after dynamic time compensation provided by the present invention.
FIG. 4 is a timing diagram of the two-star semimajor axis deviation determined by elliptical geometry fitting as provided by the present invention.
Detailed Description
The invention provides a high-precision semimajor axis deviation calculation method based on double-star on-orbit telemetry data. Aiming at the condition that an inter-satellite link and relative measurement are not established in the initial stage of orbit entering, supplement along the course distance between two satellites is constructed through satellite-ground large loop GNSS track telemetering data of the two satellites, and further quantification of semimajor axis deviation is achieved through a geometric fitting method.
As shown in fig. 1, the method for calculating the high-precision semimajor axis deviation based on the two-star on-orbit telemetry data of the invention comprises the following steps:
step 1: according to the double-satellite track telemetering data (GNSS measured data and track number determined by absolute navigation), the threshold value judgment is combined, and the error data are eliminated;
step 2: time synchronization is carried out on track telemetering data files acquired by the two stars respectively;
and step 3: processing the double-satellite orbit telemetering data on the basis of time synchronization, establishing a formation coordinate system with a main satellite as an origin, and determining the course deviation of the two satellites;
and 4, step 4: performing dynamic compensation on the basis of the course deviation of the two stars along the direction, and obtaining the relative position in the tangent plane under the formation coordinate system after the compensation along the course deviation is obtained;
and 5: and carrying out ellipse geometric fitting on the basis of the relative position in the tangent plane under the formation coordinate system to obtain the high-precision semimajor axis deviation.
Step 1, when preprocessing orbit data, setting a threshold judgment formula according to the dynamics characteristics of the orbit:
6.79×106m≤||r||≤7.05×106m
7.46×103m/s≤||v||≤7.75×103m/s
Figure GDA0002799427700000021
Figure GDA0002799427700000022
wherein x, y, z are the three-axis position data under the earth fixed system measured by the GNSS receiver, vx,vy,vzTriaxial velocity data under a geostationary system measured for a GNSS receiver. And the measurement input in the out-of-tolerance threshold range is taken as error data to be eliminated, and the calculation of navigation filtering is not introduced any more.
And 2, carrying out time synchronization on the track telemetering data files acquired by the two stars respectively when carrying out time synchronization. In the file after the synchronous processing, the two-star orbit data compared in the same row are in the same epoch moment.
Step 3, when determining the deviation along the course, further comprising:
a) according to the epoch parameters, the GNSS measurement data of two stars are converted from a ground-fixed system to an inertial system, and the coordinate conversion process comprises time conversion and calculation of the time difference RPNutating RNPolar motion RMAnd a rotation matrix RSThe conversion formula is as follows
xJ2000＝(RM·RS·RN·RP)-1xWGS84
b) Using the principal star as the origin of the formation coordinate system, and using the track root determined by absolute navigation to establish the formation coordinate system, wherein the conversion formula is as follows
Figure GDA0002799427700000031
Wherein L isiHThe transformation matrix from the inertial system to the formation coordinate system is shown, u, i and omega are latitude argument, orbit inclination angle and rising point right ascension under the inertial system, and subscript is the orbit transient root of the main star used for calculation, which is different from the orbit flat root.
c) And the Y-axis value in the formation coordinate system is the deviation along the course.
When the dynamic compensation is carried out in the step 4,
the time-compensating duration delta t is determined by the deviation l along the course and the flying speed v along the course:
Figure GDA0002799427700000032
the dynamic time compensation is based on delta t and combines a dynamic equation
Figure GDA0002799427700000033
Wherein
Figure GDA0002799427700000041
And fourth order Longge Kutta method
Figure GDA0002799427700000042
Figure GDA0002799427700000043
Figure GDA0002799427700000044
Figure GDA0002799427700000045
Figure GDA0002799427700000046
In the formula, r is a lower position module value of the satellite inertial coordinate system, and the unit is m;
Figure GDA0002799427700000047
the unit is m/s, which is the lower position velocity of the inertial coordinate system of the satellite J2000.0;
μ denotes the earth's gravitational constant, 3.986005 × 1014m3/s2；
aJ2、aJ3、aJ4Respectively represent J2, J3 and J4 perturbation accelerations with the unit of m/s2；
J2Is 1.082636X 10-3；J3Is-2.5356 x 10-6；J4Is-1.62336 x 10-6；
REThe earth radius is 6378140 m.
After the dynamic time compensation, the relative position in the tangent plane is obtained, as shown in fig. 2.
An ellipse geometric fit is performed at step 5, for the general equation of the ellipse:
ax2+bxy+cy2+dx+ey+f＝0
defining the intermediate variables:
p1＝a(cosη)2-bcosηsinη+c(sinη)2
p2＝2(a-c)cosηsinη+b((cosη)2-(sinη)2)
p3＝a(sinη)2+bcosηsinη+c(cosη)2
p4＝dcosη-esinη
p5＝dsinη+ecosη
p6＝f
Figure GDA0002799427700000048
ellipse geometric parameter value:
Figure GDA0002799427700000051
Figure GDA0002799427700000052
Figure GDA0002799427700000053
Figure GDA0002799427700000054
Figure GDA0002799427700000055
wherein (x)c,zc) Is the center of the ellipse, aFFLength of the major half-axis of the ellipse, bFFIs the length of the short semi-axis of the ellipse, and eta is the included angle between the long semi-axis of the ellipse and the horizontal axis. Value z of the center position of the ellipsecI.e. the semimajor axis deviation of the two stars. FIG. 4 is a timing diagram of the two-star semimajor axis deviation determined by ellipse geometric fitting, illustrating that the method can effectively determine the two-star semimajor axis deviation, and is effective and stable for a long period of time.
The method can overcome the influence of the track curvature on the determination of the double-star semi-major axis deviation under the condition of long distance at the initial stage of track entering, and obviously improve the determination precision of the semi-major axis deviation.
While the present invention has been described in detail with reference to the preferred embodiments, it should be understood that the above description should not be taken as limiting the invention. Various modifications and alterations to this invention will become apparent to those skilled in the art upon reading the foregoing description. Accordingly, the scope of the invention should be determined from the following claims.

Claims (7)

1. A semimajor axis deviation calculation method based on double-star on-orbit telemetry data is characterized by comprising the following steps:
step 1: according to the double-star orbit telemetering data, combining threshold judgment and rejecting error data;
step 2: time synchronization is carried out according to the track telemetering data files acquired by the two stars respectively;
and step 3: processing the double-satellite orbit telemetering data on the basis of time synchronization, establishing a formation coordinate system with a main satellite as an origin, and determining the course deviation of the two satellites;
and 4, step 4: obtaining the relative position in the cutting plane under the formation coordinate system during dynamic compensation;
and 5: carrying out ellipse geometric fitting on the relative position in the tangent plane under the formation coordinate system to obtain semimajor axis deviation; wherein, for an ellipse the general equation:
ax2+bxy+cy2+dx+ey+f＝0
defining the intermediate variables:
p1＝a(cosη)2-bcosηsinη+c(sinη)2
p2＝2(a-c)cosηsinη+b((cosη)2-(sinη)2)
p3＝a(sinη)2+bcosηsinη+c(cosη)2
p4＝dcosη-esinη
p5＝dsinη+ecosη
p6＝f
Figure FDA0002799427690000011
ellipse geometric parameter value:
Figure FDA0002799427690000012
the value z of the center position of the ellipsecAs a bi-star semi-major axis deviation.
2. The method for calculating the semi-major axis deviation based on the two-star on-orbit telemetry data as claimed in claim 1, wherein in step 1, the two-star orbit telemetry data comprises GNSS measurement data and orbit number determined by absolute navigation.
3. The method for calculating the semimajor axis deviation based on the two-star on-orbit telemetry data as claimed in claim 1, wherein in step 1, a threshold judgment formula is set according to the dynamics characteristics of the orbit:
6.79×106m≤||r||≤7.05×106m
7.46×103m/s≤||v||≤7.75×103m/s
Figure FDA0002799427690000021
Figure FDA0002799427690000022
wherein x, y, z are the three-axis position data under the earth fixed system measured by the GNSS receiver, vx,vy,vzMeasuring triaxial speed data under a ground fixation system for a GNSS receiver; and the measurement input in the out-of-tolerance threshold range is taken as error data to be eliminated, and the calculation of navigation filtering is not introduced any more.
4. The method for calculating the semimajor axis deviation based on the two-star on-orbit telemetry data as claimed in claim 1, wherein in the step 2, the two-star orbit data compared in the same row in the orbit telemetry data file after the synchronization processing are in the same epoch time.
5. The method for calculating the semimajor axis deviation based on the two-star on-orbit telemetry data as claimed in claim 1, wherein the step 3 further comprises:
a) converting the GNSS measurement data of the two satellites into an inertial system from a geostationary system according to the epoch parameters;
b) using a main star as an origin of a formation coordinate system, and establishing the formation coordinate system by using the track root determined by absolute navigation;
c) and taking the Y-axis value in the formation coordinate system as the deviation along the course.
6. The method for calculating the semimajor axis deviation based on the two-star on-orbit telemetry data as claimed in claim 1, wherein in step 4, the time-compensating duration Δ t is determined by the deviation l along the course and the flying speed v along the course:
Figure FDA0002799427690000023
the dynamic time compensation is based on delta t and combines a dynamic equation
Figure FDA0002799427690000031
Wherein
Figure FDA0002799427690000032
And fourth order Longge Kutta method
Figure FDA0002799427690000033
Figure FDA0002799427690000034
Figure FDA0002799427690000035
Figure FDA0002799427690000036
Figure FDA0002799427690000037
In the formula, r is a lower position module value of the satellite inertial coordinate system, and the unit is m;
Figure FDA0002799427690000038
the unit is m/s, which is the lower position velocity of the inertial coordinate system of the satellite J2000.0;
μ denotes the earth's gravitational constant, 3.986005 × 1014m3/s2；
aJ2、aJ3、aJ4Respectively represent J2, J3 and J4 perturbation accelerations with the unit of m/s2；
J2Is 1.082636X 10-3；J3Is-2.5356 x 10-6；J4Is-1.62336 x 10-6；
RE6378140m radius of the earth;
and obtaining the relative position in the tangent plane after the time compensation of dynamics.
7. The method for calculating the semimajor axis deviation based on the two-star on-orbit telemetry data as claimed in claim 1, wherein the semimajor axis deviation calculation method is suitable for the working state that the inter-satellite link and the relative measurement are not established at the initial stage of the orbit entering.
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CN112327261B (en) * 2020-10-22 2022-10-25 上海卫星工程研究所 Distributed InSAR satellite time synchronization on-orbit testing method and system
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Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2113786A1 (en) * 2008-04-30 2009-11-04 GMV Aerospace and Defence S.A. Method for autonomous determination of protection levels for GNSS positioning based on navigation residuals and an isotropic confidence ratio
CN102322862A (en) * 2011-06-29 2012-01-18 航天东方红卫星有限公司 Method for determining absolute orbit and relative orbit of formation flight satellite
CN103645485A (en) * 2013-10-28 2014-03-19 中国科学院国家授时中心 Pseudorange differential method based on dual-satellite time difference and frequency difference passive positioning
CN103940431A (en) * 2014-04-11 2014-07-23 北京空间飞行器总体设计部 Tangential low-thrust in-orbit circular orbit calibration method based on (Global Navigation Satellite System) GNSS precise orbit determination
CN109062243A (en) * 2018-10-31 2018-12-21 长光卫星技术有限公司 A kind of autonomous collision avoidance method of spacecraft energetic optimum under multiple constraint
CN109189102A (en) * 2018-11-23 2019-01-11 上海航天控制技术研究所 A kind of method of high precision computation double star semi-major axis deviation on star

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103557872B (en) * 2013-11-04 2015-11-25 北京航空航天大学 System ensemble error real-time computing technique in a kind of RNP
US20160097788A1 (en) * 2014-10-07 2016-04-07 Snappafras Corp. Pedestrian direction of motion determination system and method

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2113786A1 (en) * 2008-04-30 2009-11-04 GMV Aerospace and Defence S.A. Method for autonomous determination of protection levels for GNSS positioning based on navigation residuals and an isotropic confidence ratio
CN102322862A (en) * 2011-06-29 2012-01-18 航天东方红卫星有限公司 Method for determining absolute orbit and relative orbit of formation flight satellite
CN103645485A (en) * 2013-10-28 2014-03-19 中国科学院国家授时中心 Pseudorange differential method based on dual-satellite time difference and frequency difference passive positioning
CN103940431A (en) * 2014-04-11 2014-07-23 北京空间飞行器总体设计部 Tangential low-thrust in-orbit circular orbit calibration method based on (Global Navigation Satellite System) GNSS precise orbit determination
CN109062243A (en) * 2018-10-31 2018-12-21 长光卫星技术有限公司 A kind of autonomous collision avoidance method of spacecraft energetic optimum under multiple constraint
CN109189102A (en) * 2018-11-23 2019-01-11 上海航天控制技术研究所 A kind of method of high precision computation double star semi-major axis deviation on star

Non-Patent Citations (5)

* Cited by examiner, † Cited by third party
Title
GEO/IGSO/MEO卫星轨道根数演化分析;毛悦 等;《测绘科学》;20090131;第34卷(第01期);第119-121页 *
Rapid Evaluation Algorithm of Station-keeping Control Results for Three-axis Stabilized GEO Satellites;ZHENG Jun 等;《2013 Fifth International Conference on Intelligent Human-Machine Systems and Cybernetics》;20131231;第1-3页 *
分布式卫星群构形初始化控制策略;王兆魁 等;《宇航学报》;20040531;第25卷(第03期);第334-337页 *
基于约化相对轨道拟平根数的长期稳定高精度卫星编队导航技术;杨盛庆 等;《空间控制技术与应用》;20170228;第43卷(第01期);第30-35页 *
环境与灾害监测预报小卫星星座A/B卫星运行与发展展望;白照广;《航天器工程》;20121031(第05期);第11-18页 *

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