Pipeline deviation full-orbit fitting method of strict regression orbit satellite

Abstract

A pipeline deviation full-orbit fitting method of a strict regression orbit satellite can determine single-point pipeline deviation on the basis of alignment of a reference trajectory coordinate system and an orbit cutting plane according to the characteristics of a strict regression orbit, and can also perform calculation and curve fitting on data of at least one orbit. And the fitted maximum deviation is used as the pipeline deviation of the satellite, so that comprehensive analysis of strict regression precision of the satellite is carried out, limitation of single-point pipeline deviation calculation is effectively overcome, and meanwhile, the influence of single-point error data on pipeline deviation calculation can be inhibited by adopting a curve fitting method, so that the accuracy of pipeline deviation calculation is improved. And the calculation result is more accurate and reliable.

Description

Technical Field

[0001] The invention relates to a pipeline deviation full-orbit fitting method for a strictly regressive orbit satellite, and belongs to the technical field of satellite high-precision orbit determination and control. Background Art

[0002] With the further increase in the demand for on-orbit applications of formation satellites, the previous orbit design and control can no longer meet the higher requirements of satellite orbit control accuracy and autonomy level for highly complex applications. Strict regression orbit is a new orbit design and control concept proposed for satellite re-orbit earth observation missions in recent years. The design feature of the strict regression orbit is that within a regression cycle, the satellite's space trajectory is completely consistent in the initial state and the final state, thus achieving strict regression. When the satellite crosses the strict regression reference orbit tangent plane, the deviation from the origin of the trajectory coordinate system is the pipeline deviation of the current reference point. The previous method of using six orbit numbers to determine the satellite orbit is no longer applicable to the analysis of pipeline deviations, and an accurate and reliable strict regression orbit pipeline deviation calculation method is urgently needed. Summary of the invention

[0003] The technical problem solved by the present invention is: in view of the problem that the method of determining satellite orbits using six orbital numbers in the current prior art is no longer applicable to pipeline deviation analysis, a pipeline deviation full-orbit fitting method for strictly regressive orbit satellites is proposed.

[0004] The present invention solves the above technical problems by the following technical solutions:

[0005] A pipeline deviation full-orbit fitting method for a strictly regressive orbit satellite, comprising:

[0006] Collect satellite measured orbit state sequence information and reference orbit state sequence information;

[0007] The reference orbit sequence time of the current reference orbit point is recursively pushed forward in a regression cycle to find the reference orbit point closest to the pre-selected time point of the satellite's measured orbit state sequence information;

[0008] With the reference trajectory point as the origin, a formation coordinate system is established, and the relative position vector of the satellite in the formation coordinate system at the pre-selected time point is calculated;

[0009] The recursive operation time required for the satellite to move from the position at the pre-selected time point to the reference orbit point under the measured orbit is calculated based on the relative position vector, and the orbit state sequence information corresponding to the orbit point after the recursive operation is used to calculate the relative position vector of the orbit point after the recursive operation in the trajectory coordinate system with the reference trajectory point as the origin;

[0010] Traversing the pre-selected time points in the measured orbit state sequence information of all satellites under the measured orbit, obtaining the relative position vector of the orbit point after recursive operation corresponding to each pre-selected time point in the trajectory coordinate system with the reference trajectory point as the origin;

[0011] The double coordinate information is obtained from the relative position vectors of all orbital points after recursion in the trajectory coordinate system with the reference trajectory point as the origin to form a two-dimensional vector sequence. After ellipse fitting, the minimum circumscribed circle radius is obtained as the pipeline deviation of the satellite.

[0012] The satellite measured orbit state sequence information includes time t i , three-dimensional position [x i ,y i z i ], three-dimensional velocity [vx i ,vy i ,vz i ]; the reference orbit state sequence information includes time t ci , three-dimensional position [x ci ,y ci ,z ci ], three-dimensional velocity [vx ci ,vy ci ,vz ci ]; i = 1, 2, 3…N; the satellite measured orbit state sequence information and the reference orbit state sequence information are both parameter information within a regression cycle;

[0013] When the pre-selected time point is determined, the reference orbit point P closest to the pre-selected time point is selected on the satellite measured orbit c , determine the reference orbit point P c The corresponding reference orbit time t on the reference orbit cj .

[0014] According to the reference orbit point P c The reference orbit time t cj , pre-select time point t i The comparison results of the reference orbit point P in the case of insufficient time c The reference orbit time t cj , according to the whole regression period forward recursion, determine the number of regression periods N that need to be recursed forward:

[0015] N=fix((t i -t cj ) / T)

[0016] Where T is the regression period, N is the number of regression periods that need to be recursively calculated, and the fix() operation indicates rounding the calculated numerical result.

[0017] After determining the number of regression cycles N that need to be recursively pushed forward, select the reference orbit point to be targeted in the reference orbit point sequence. The method is:

[0018] M=fix(((t i -t cj )-N*T) / Ts)+1

[0019] Where Ts is the time interval between adjacent strict regression orbit reference points, and M represents the Mth point in the reference orbit point sequence, that is, the reference orbit point that needs to be targeted.

[0020] The formation coordinate system OXYZ takes the vector direction from the center of the earth E to the origin as the OX axis, sets the OY axis in the orbital plane of the main satellite, points to the flight direction as the OX axis, and the OZ axis, OX axis and OY axis satisfy the right-hand rule. The preselected time point t i At the moment, the relative position vector of the satellite in the formation coordinate system is recorded as R 1 (R Hx1 , R Hy1 , R Hz1 ).

[0021] The relative position vector R 1 The calculation method is:

[0022]

[0023] In the formula, the reference orbit point P c The position vector of a point in the inertial coordinate system, P c The velocity vector of a point in the inertial coordinate system, P s The position vector of a point in the inertial coordinate system, P s The velocity vector of the point in the inertial coordinate system, A i2b is the transformation matrix from the inertial system to the formation coordinate system.

[0024] The track point P after running s The corresponding orbital state sequence information includes time t k , three-dimensional position [x k ,y k ,z k ], three-dimensional velocity [vx k ,vy k ,vz k ]; track point P after running s The relative position vector in the trajectory coordinate system with Pc as the origin is R i (R gjxi , R gjyi, R gjzi );

[0025] The calculation method for the running time △t required for the satellite to run from the position at the pre-selected time point to the reference orbit point under the measured orbit is:

[0026]

[0027] In the formula, the molecule is the projection R of the relative position in the Hy direction Hy1 , the denominator is the magnitude of the velocity vector

[0028] According to the track point P after running s The time t i +△t time satellite orbital elements [a,e,i,Ω,ω,M+n*Δt], converted to obtain the orbit point P after operation s The orbital state sequence information [x si ,y si ,z si ] and [vx si ,vy si ,vz si ],calculate In is the relative position vector of the trajectory coordinate system of the origin The method is:

[0029]

[0030] In the formula, P c The position vector of the point in the Earth-fixed coordinate system, P c The velocity vector of a point in the Earth-fixed coordinate system is, P s The position vector of the point in the Earth-fixed coordinate system, P s Velocity vector of a point in the Earth-fixed coordinate system.

[0031] The calculation method of the relative position vector of the orbital point after running recursion corresponding to the pre-selected time point in the measured orbital state sequence information of all satellites is the same in the trajectory coordinate system with the reference trajectory point as the origin. The reference orbit time interval is considered when selecting each pre-selected time point. In the absence of orbit control, the measured orbit is the same as the reference orbit. In the case of orbit control, the measured orbit is smaller than the reference orbit.

[0032] The N relative position vectors X gj and Z gjThe coordinates form a two-dimensional vector sequence, and this two-dimensional sequence is fitted with an ellipse. The general equation of the ellipse is obtained by numerical fitting using the method of undetermined coefficients:

[0033] ax 2 +bxy+cy 2 +dx+ey+f=0

[0034] Let α=[a,b,c,d,e,f] T , x=[x 2 ,xy,y 2 ,x,y,1] T , using the elliptic least squares fitting method, solve α = [a, b, c, d, e, f] T ;

[0035] After solving, draw the elliptic curve and gj O gj Z gj In the coordinate system, with the origin as the center, draw the circumscribed circle of the elliptical curve, and use the minimum circumscribed circle radius as the satellite's pipeline deviation.

[0036] The advantages of the present invention compared with the prior art are:

[0037] (1) The present invention provides a pipeline deviation full-track fitting method for a strictly regressed orbit satellite, which can determine the single-point pipeline deviation based on the characteristics of the strictly regressed orbit and the alignment of the reference trajectory coordinate system and the tangent plane. It can also calculate and curve fit the data of at least one track, and use the maximum deviation after fitting as the pipeline deviation of the satellite, so as to conduct a comprehensive analysis of the satellite's strict regression accuracy. This method effectively overcomes the limitations of single-point pipeline deviation calculation. At the same time, the use of a curve fitting method can suppress the influence of single-point error data on pipeline deviation calculation, making the calculation result more accurate and reliable;

[0038] (2) The process of the present invention is clear and has strong engineering feasibility: combined with the actual situation of orbit changes, the orbit recursive calculation amount in the orbit cutting and alignment process is avoided, and it has good engineering feasibility. At the same time, the method has strong anti-interference ability and the determination result is accurate and reliable: the curve fitting process avoids the errors and data disturbances that may exist in single-shot or partial orbit measurement data, and can process orbit data of different scales as needed, which can be realized both onboard and on the ground. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1 A flow chart for determining the full track pipeline deviation of a strict regression track provided for the invention;

[0040] Figure 2 A schematic diagram of a pipeline deviation curve in a trajectory coordinate system within one orbital period of a satellite provided by the invention; DETAILED DESCRIPTION

[0041] A pipeline deviation full-orbit fitting method for a strict regression orbit satellite can determine the single-point pipeline deviation based on the alignment of the reference trajectory coordinate system and the tangent plane according to the characteristics of the strict regression orbit, and can also calculate and curve fit the data of at least one orbit, and use the maximum deviation after fitting as the pipeline deviation of the satellite, so as to conduct a comprehensive analysis of the satellite strict regression accuracy and effectively overcome the limitations of the single-point pipeline deviation calculation. At the same time, the use of the curve fitting method can suppress the influence of single-point error data on the pipeline deviation calculation, making the calculation result more accurate and reliable.

[0042] The pipeline deviation full orbit fitting method for strictly regressive orbit satellites has the following steps:

[0043] Collect satellite measured orbit state sequence information and reference orbit state sequence information;

[0044] The reference orbit sequence time of the current reference orbit point is recursively pushed forward in a regression cycle to find the reference orbit point closest to the pre-selected time point of the satellite's measured orbit state sequence information;

[0045] The specific meaning of regression period recursion is that for a specific set of strict regression orbit parameters, the regression period is determined (such as 3 days or 8 days, and there can only be one starting time). Therefore, for actual tasks, it is necessary to correspond the starting time of the reference orbit to the vicinity of the actual mission time before it can be used.

[0046] With the reference trajectory point as the origin, a formation coordinate system is established, and the relative position vector of the satellite in the formation coordinate system at the pre-selected time point is calculated;

[0047] The recursive operation time required for the satellite to move from the position at the pre-selected time point to the reference orbit point under the measured orbit is calculated based on the relative position vector, and the orbit state sequence information corresponding to the orbit point after the recursive operation is used to calculate the relative position vector of the orbit point after the recursive operation in the trajectory coordinate system with the reference trajectory point as the origin;

[0048] Traversing the pre-selected time points in the measured orbit state sequence information of all satellites under the measured orbit, obtaining the relative position vector of the orbit point after recursive operation corresponding to each pre-selected time point in the trajectory coordinate system with the reference trajectory point as the origin;

[0049] The double coordinate information is obtained from the relative position vectors of all orbital points after recursion in the trajectory coordinate system with the reference trajectory point as the origin to form a two-dimensional vector sequence. After ellipse fitting, the minimum circumscribed circle radius is obtained as the pipeline deviation of the satellite.

[0050] The satellite measured orbit state sequence information includes time t i, three-dimensional position [x i ,y i z i ], three-dimensional velocity [vx i ,vy i ,vz i ]; the reference orbit state sequence information includes time t ci , three-dimensional position [x ci ,y ci ,z ci ], three-dimensional velocity [vx ci ,vy ci ,vz ci ]; i = 1, 2, 3…N; the satellite measured orbit state sequence information and the reference orbit state sequence information are both parameter information within a regression cycle;

[0051] When the pre-selected time point is determined, the reference orbit point P closest to the pre-selected time point is selected on the satellite measured orbit c , determine the reference orbit point P c The corresponding reference orbit time t on the reference orbit cj .

[0052] According to the reference orbit point P c The reference orbit time t cj , pre-select time point t i The comparison results of the reference orbit point P in the case of insufficient time c The reference orbit time t cj , according to the whole regression period forward recursion, determine the number of regression periods N that need to be recursed forward:

[0053] N=fix((t i -t cj ) / T)

[0054] Where T is the regression period, N is the number of regression periods that need to be recursively calculated, and the fix() operation indicates rounding the calculated numerical result.

[0055] After determining the number of regression cycles N that need to be recursively pushed forward, select the reference orbit point to be targeted in the reference orbit point sequence. The method is:

[0056] M=fix(((t i -t cj )-N*T) / Ts)+1

[0057] Where Ts is the time interval between adjacent strict regression orbit reference points, and M represents the Mth point in the reference orbit point sequence, that is, the reference orbit point that needs to be targeted.

[0058] The formation coordinate system OXYZ, with the vector direction from the center of the earth E to the origin as the OX axis, the OY axis is set in the orbital plane of the main satellite, and the direction of flight is the OX axis. The OZ axis, OX axis, and OY axis satisfy the right-hand rule. The pre-selected time point t i At the moment, the relative position vector of the satellite in the formation coordinate system is recorded as R 1 (R Hx1 , R Hy1 , R Hz1 ).

[0059] The formation coordinate system OXYZ, with the vector direction from the center of the earth E to the origin as the OX axis, the OY axis is set in the orbital plane of the main satellite, and the direction of flight is the OX axis. The OZ axis, OX axis, and OY axis satisfy the right-hand rule. The pre-selected time point t i At the moment, the relative position vector of the satellite in the formation coordinate system is recorded as R 1 (R Hx1 , R Hy1 , R Hz1 ).

[0060] Track point P after running s The corresponding orbital state sequence information includes time t k , three-dimensional position [x k ,y k ,z k ], three-dimensional velocity [vx k ,vy k ,vz k ]; track point P after running s The relative position vector in the trajectory coordinate system with Pc as the origin is R i (R gjxi , R gjyi , R gjzi );

[0061] The calculation method for the running time △t required for the satellite to run from the position at the pre-selected time point to the reference orbit point under the measured orbit is:

[0062]

[0063] In the formula, the molecule is the projection R of the relative position in the Hy direction Hy1 , the denominator is the magnitude of the velocity vector

[0064] According to the track point P after running s The time t i +△t time satellite orbital elements [a,e,i,Ω,ω,M+n*Δt], converted to obtain the orbit point P after operation s The orbital state sequence information [x si ,y si ,zsi ] and [vx si ,vy si ,vz si ],calculate In is the relative position vector of the trajectory coordinate system of the origin The method is:

[0065]

[0066] In the formula, P c The position vector of the point in the Earth-fixed coordinate system, P c The velocity vector of a point in the Earth-fixed coordinate system is, P s The position vector of the point in the Earth-fixed coordinate system, P s Velocity vector of a point in the Earth-fixed coordinate system.

[0067] The calculation method of the relative position vector of the orbit point after recursion corresponding to the pre-selected time point in the satellite's measured orbit state sequence information is the same in the trajectory coordinate system with the reference trajectory point as the origin. The reference orbit time interval is considered when selecting each pre-selected time point. In the absence of orbit control, the measured orbit is the same as the reference orbit. In the case of orbit control, the measured orbit is smaller than the reference orbit.

[0068] The X of N relative position vectors gj and Z gj The coordinates form a two-dimensional vector sequence, and this two-dimensional sequence is fitted with an ellipse. The general equation of the ellipse is obtained by numerical fitting using the method of undetermined coefficients:

[0069] ax 2 +bxy+cy 2 +dx+ey+f=0

[0070] Let α=[a,b,c,d,e,f] T , x=[x 2 ,xy,y 2 ,x,y,1] T , using the elliptic least squares fitting method, solve α = [a, b, c, d, e, f] T ;

[0071] After solving, draw the elliptic curve and gj O gj Z gj In the coordinate system, with the origin as the center, draw the circumscribed circle of the elliptical curve, and use the minimum circumscribed circle radius as the satellite's pipeline deviation.

[0072] The following is further described in conjunction with the accompanying drawings and preferred embodiments of the specification:

[0073] In the current embodiment, if Figure 1 As shown, the pipeline deviation full orbit fitting determination algorithm of the strictly regressive orbit satellite includes the following steps:

[0074] Step 1: Prepare the orbital state sequence (including time t i , three-dimensional position [x i ,y i z i ], three-dimensional velocity [vx i ,vy i ,vz i ]) and the state sequence of the reference orbit within a regression period (including time t c0 , three-dimensional position [x ci ,y ci ,z ci ], three-dimensional velocity [vx ci ,vy ci ,vz ci ]), where i = 1, 2, 3 ... N;

[0075] Step 2: Find the time distance t i The nearest reference orbit point P c (The corresponding time is t cj ), when the reference orbit sequence is insufficient, the whole regression period is recursively forward t cj =N*T+t c0 , where N is the number of regression cycles of forward recursion, and T is the regression period;

[0076] Take N = fix((t i -t cj ) / T), M=fix(((t i -t cj )-N*T) / Ts)+1

[0077] in:

[0078] Ts is the time interval between adjacent strict regression orbit reference points;

[0079] T is the regression period of the strictly regressive orbit;

[0080] N is the number of regression cycles that need to be recursively pushed forward;

[0081] The fix() operation means rounding the calculated numerical result;

[0082] M indicates that the Mth point in the reference track point sequence is the corresponding reference point that needs to be aimed at.

[0083] Step 3: Take reference orbit point P c As the origin, establish the formation coordinate system and calculate the satellite t i The relative position vector R in this coordinate system 1 (R Hx1 , R Hy1 , R Hz1 );

[0084] The formation coordinate system OXYZ is defined as: OX is the vector direction from the center of the earth E to the origin; OY is in the orbital plane of the main satellite, and the flight direction perpendicular to OX is positive; OZ, OX and OY form a right-handed coordinate system.

[0085]

[0086] in: P c The position vector of a point in the inertial coordinate system, P c The velocity vector of a point in the inertial coordinate system, P s The position vector of a point in the inertial coordinate system, P s The velocity vector of the point in the inertial coordinate system, A i2b is the transformation matrix from the inertial system to the formation coordinate system. The specific algorithm is as follows:

[0087] in

[0088] Step 4: According to R 1 The satellite can be quickly calculated from t i Run to P c The time required, △t, can be obtained by orbit recursion i +△t orbital state P s (Time t k , three-dimensional position [x k ,y k ,z k ], three-dimensional velocity [vx k ,vy k ,vz k ]); and calculate the relative position vector R of Ps in the trajectory coordinate system with Pc as the origin i (R gjxi , R gjyi , R gjzi ), the specific algorithm is as follows:

[0089] By [x i ,y i z i ] and [vx i ,vy i ,vz i ] is converted to obtain the orbital elements [a, e, i, Ω, ω, M]. Due to the short recursion time, except for the mean anomaly angle M, which is recursively calculated according to time and the average angular velocity of the orbit, the other orbital elements are slow variables and take the same value as before the recursion. In this way, after a simplified orbit recursion, the orbital elements [a, e, i, Ω, ω, M + n * Δt] after Δt can be quickly obtained, and then converted to obtain the orbital state P s [x si ,y si ,z si ] and [vx si ,vy si ,vz si ].calculate In is the relative position vector of the trajectory coordinate system of the origin The specific process is:

[0090] Calculate the transformation matrix from the ground-fixed system to the trajectory coordinate system:

[0091] A c2g =[x g ,y g ,z g ] T ,in

[0092]

[0093] in: P c The position vector of the point in the Earth-fixed coordinate system, P c The velocity vector of a point in the Earth-fixed coordinate system is, P s The position vector of the point in the Earth-fixed coordinate system, P s Velocity vector of a point in the Earth-fixed coordinate system.

[0094] Step 5 Repeat the above steps to obtain the relative position vector R corresponding to all the measured orbital state sequences in one track. i(i=1,2,3…N), the processing method of each measured point is the same. Taking into account the time interval of the reference orbit, the time interval of the measured points should be equivalent to that of the reference orbit when there is no control; if orbit control is performed in a certain time period, the time interval of the measured points in this time period can be smaller, so as to better reflect the actual situation of the pipeline deviation during orbit control.

[0095] Step 6: X of N relative position vectors gj and Z gj The coordinates form a two-dimensional vector sequence, and an ellipse is fitted to the two-dimensional sequence. The minimum circumscribed circle radius of the obtained curve is the satellite pipeline deviation.

[0096] Obtaining the general equation of an ellipse by numerical fitting using the method of undetermined coefficients

[0097] ax 2 +bxy+cy 2 +dx+ey+f=0

[0098] Let α=[a,b,c,d,e,f] T , x=[x 2 ,xy,y 2 ,x,y,1] T , using the elliptic "least squares fitting method", in solving α = [a, b, c, d, e, f] T Once completed, draw the elliptic curve and gj O gj Z gj In the coordinate system, with the origin as the center, draw the circumscribed circle of the elliptical curve. The radius of the circle is the satellite's full-orbit pipeline deviation. Figure 2 As shown, it is a schematic diagram of the pipeline deviation curve in the trajectory coordinate system within one orbital period of the satellite.

[0099] The method proposed in this embodiment has a clear process and strong engineering feasibility: combined with the actual situation of orbit changes, the orbit recursive calculation amount in the orbit cutting and alignment process is avoided, and it has good engineering feasibility. At the same time, the method has strong anti-interference ability and determines the result accurately and reliably: the curve fitting process avoids the errors and data disturbances that may exist in single-shot or partial orbit measurement data, and orbit data of different scales can be processed as needed, which can be realized both onboard and on the ground.

[0100] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0101] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A pipeline deviation full-orbit fitting method for a strictly regressive orbit satellite, characterized in that include: Collect satellite measured orbit state sequence information and reference orbit state sequence information; The reference orbital sequence time of the current reference orbital point in the reference orbital state sequence information is recursively pushed forward in a regression cycle to determine the reference orbital point closest to the preselected time point of the satellite's measured orbital state sequence information; With the reference trajectory point as the origin, a formation coordinate system is established, and the relative position vector of the satellite in the formation coordinate system at the pre-selected time point is calculated; The recursive operation time required for the satellite to move from the position at the preselected time point to the reference orbit point under the measured orbit is calculated based on the relative position vector and the satellite's measured orbit state sequence information, and the relative position vector of the orbit point after the recursive operation in the trajectory coordinate system with the reference trajectory point as the origin is calculated through the orbit state sequence information corresponding to the orbit point after the recursive operation; Traversing the pre-selected time points in the measured orbit state sequence information of all satellites under the measured orbit, obtaining the relative position vector of the orbit point after recursive operation corresponding to each pre-selected time point in the trajectory coordinate system with the reference trajectory point as the origin; The double coordinate information is obtained from the relative position vectors of all orbital points after recursion in the trajectory coordinate system with the reference trajectory point as the origin to form a two-dimensional vector sequence. After ellipse fitting, the minimum circumscribed circle radius is obtained as the pipeline deviation of the satellite.

2. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 1 is characterized in that: The satellite measured orbit state sequence information includes time t i , three-dimensional position [x i ,y i z i ], three-dimensional velocity [vx i ,vy i ,vz i ]; the reference orbit state sequence information includes time t ci , three-dimensional position [x ci ,y ci ,z ci ], three-dimensional velocity [vx ci ,vy ci ,vz ci ]; i = 1, 2, 3…N; the satellite measured orbit state sequence information and the reference orbit state sequence information are both parameter information within a regression cycle; When the pre-selected time point is determined, the reference orbit point P closest to the pre-selected time point is selected on the satellite measured orbit. c , determine the reference orbit point P c The corresponding reference orbit time t on the reference orbit cj .

3. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 2 is characterized in that: According to the reference orbit point P c The reference orbit time t cj , pre-select time point t i The comparison results of the reference orbit point P in the case of insufficient time c The reference orbit time t cj , according to the whole regression period forward recursion, determine the number of regression periods N that need to be recursed forward: N=fix((t i -t cj ) / T) Where T is the regression period, N is the number of regression periods that need to be recursively calculated, and the fix() operation indicates rounding the calculated numerical result.

4. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 2 is characterized in that: After determining the number of regression cycles N that need to be recursively pushed forward, select the reference orbit point to be targeted in the reference orbit point sequence. The method is: M=fix(((t i -t cj )-N*T) / Ts)+1 Where Ts is the time interval between adjacent strict regression orbit reference points, and M represents the Mth point in the reference orbit point sequence, that is, the reference orbit point that needs to be targeted.

5. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 1 is characterized in that: The formation coordinate system OXYZ takes the vector direction from the center of the earth E to the origin as the OX axis, sets the OY axis in the orbital plane of the main satellite, points to the flight direction as the OX axis, and the OZ axis, OX axis and OY axis satisfy the right-hand rule. The preselected time point t i At this moment, the relative position vector of the satellite in the formation coordinate system is recorded as R1(R Hx1 , R Hy1 , R Hz1 ).

6. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 5 is characterized in that: The relative position vector R1 is calculated as follows: In the formula, the reference orbit point P c The position vector of a point in the inertial coordinate system, P c The velocity vector of a point in the inertial coordinate system, P s The position vector of a point in the inertial coordinate system, P s The velocity vector of the point in the inertial coordinate system, A i2b is the transformation matrix from the inertial system to the formation coordinate system.

7. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 6 is characterized in that: The track point P after running s The corresponding orbital state sequence information includes time t k , three-dimensional position [x k ,y k ,z k ], three-dimensional velocity [vx k ,vy k ,vz k ]; track point P after running s The relative position vector in the trajectory coordinate system with Pc as the origin is R i (R gjxi , R gjyi , R gjzi ); The calculation method for the running time △t required for the satellite to run from the position at the pre-selected time point to the reference orbit point under the measured orbit is: In the formula, the molecule is the projection R of the relative position in the Hy direction Hy1 , the denominator is the magnitude of the velocity vector 8. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 6 is characterized by: According to the track point P after running s The time t i +△t time satellite orbital elements [a,e,i,Ω,ω,M+n*Δt], converted to obtain the orbit point P after operation s The orbital state sequence information [x si ,y si ,z si ] and [vx si ,vy si ,vz si ],calculate In is the relative position vector of the trajectory coordinate system of the origin The method is: In the formula, P c The position vector of the point in the Earth-fixed coordinate system, P c The velocity vector of a point in the Earth-fixed coordinate system is, P s The position vector of the point in the Earth-fixed coordinate system, P s Velocity vector of a point in the Earth-fixed coordinate system.

9. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 1 is characterized in that: The calculation method of the relative position vector of the orbital point after running recursion corresponding to the pre-selected time point in the measured orbital state sequence information of all satellites is the same in the trajectory coordinate system with the reference trajectory point as the origin. The reference orbit time interval is considered when selecting each pre-selected time point. In the absence of orbit control, the measured orbit is the same as the reference orbit. In the case of orbit control, the measured orbit is smaller than the reference orbit.

10. The pipeline deviation full-orbit fitting method of a strictly regressive orbit satellite according to claim 2 is characterized in that: The N relative position vectors X gj and Z gj The coordinates form a two-dimensional vector sequence, and this two-dimensional sequence is fitted with an ellipse. The general equation of the ellipse is obtained by numerical fitting using the method of undetermined coefficients: ax 2 +bxy+cy 2 +dx+ey+f=0 Let α=[a,b,c,d,e,f] T , x=[x 2 ,xy,y 2 ,x,y,1] T , using the elliptic least squares fitting method, solve α = [a, b, c, d, e, f] T ; After solving, draw the elliptic curve and gj O gj Z gj In the coordinate system, with the origin as the center, draw the circumscribed circle of the elliptical curve, and use the minimum circumscribed circle radius as the satellite's pipeline deviation.

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