Single satellite or multi-satellite heavy orbit control method based on effective vertical baseline

Abstract

The application provides a single-satellite or multi-satellite re-track interference orbit control method based on an effective vertical baseline, and belongs to the technical field of space-borne SAR re-track interference. The single-satellite or multi-satellite re-track interference orbit control method is based on an eccentricity vector and a relative inclination, a reference-observation virtual formation model of re-track interference is established; based on the geometric relationship between the target detection area latitude and the subsatellite point latitude and the orbit root number, the relationship between the target detection area latitude and the effective vertical baseline is obtained, and the virtual formation effective vertical baseline of the reference-observation satellite is determined; and the in-plane / out-of-plane orbit control strategy suitable for single-satellite or multi-satellite re-track interference is obtained by taking the effective vertical baseline as an evaluation index. The application solves the problem that the orbit control strategy research of single-satellite re-track interference is relatively insufficient at home and abroad, can provide a new method and a new direction for establishing a remote sensing system of China, and can be applied to space-based situation awareness to support the improvement of the space-based attack and defense system of China.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of spaceborne SAR repeat-pass interferometry, in particular to a single-satellite or multi-satellite repeat-pass interferometry orbit control method based on effective vertical baseline. BACKGROUND

[0002] The emergence of spaceborne synthetic aperture radar interferometry (InSAR) overcomes the difficulty of deformation monitoring in earth observation, has the advantages of high spatial resolution, short repeat-pass cycle, large observation area, etc., and can have an imaging observation width of more than several tens of kilometers in strip mode. At the same time, the InSAR system can realize imaging and information acquisition all-weather, all-day, and large-scale, and is not affected by the shielding and influence of clouds and rain targets, compared with infrared and visible light imaging technology, it has obvious advantages. InSAR not only contains the backscattering amplitude information, but also contains the interferometric phase information. Therefore, this technology can provide high-resolution and high-precision data for global topographic mapping, tree height inversion, soil moisture inversion, global forestry research, etc.

[0003] And spaceborne repeat-pass interferometry (referred to as repeat-pass interferometry) InSAR is a further application on the basis of spaceborne InSAR imaging technology, and currently plays a key role in global elevation mapping and deformation measurement. Combined with multi-pass data to jointly invert the three-dimensional deformation information of the ground or to monitor large-scale deformation, repeat-pass interferometry InSAR uses data obtained from radar detectors in closely spaced repeat orbits to distinguish clutter and estimate surface / underground slope. This method uses cross-orbit signal migration to distinguish non-minimum point clutter and underground signal returns received at the minimum point, and uses the repeated configuration of the same sensor at different times to obtain stereo image information.

[0004] Repeat-pass interferometry requires the satellite to make at least two repeat observations of the same imaging area, and has strict requirements for the observation angle and baseline length. However, due to the demand for global observation, the revisit period of the satellite is usually between several dozen days to several tens of days. In this process, the satellite is affected by various perturbation factors, such as atmospheric resistance, non-spherical gravity of the earth, and gravity of the sun and the moon, etc., causing its flight orbit to deviate from the predetermined orbit, thereby causing the offset of the revisit orbit. This directly affects the configuration of the interference baseline, including the length and direction of the baseline. The existing control of the satellite operating state is based on the flight state of a single satellite, which can ensure the general requirements of time and spatial position of the revisit. Due to the relative lack of single-satellite repeat-pass interferometry orbit control strategy research at home and abroad, and even less multi-satellite repeat-pass interferometry orbit control method, a repeat-pass interferometry orbit control strategy with simple calculation, no complex geometric analysis, and practical use is needed. SUMMARY

[0005] The technical problem solved by the present application is that there is relatively lack of single-satellite re-orbiting interference orbit control strategy research at home and abroad at present.

[0006] The present application establishes a virtual formation system taking the adjacent two flight trajectories of a satellite as a "reference satellite" and an "observation satellite", and introduces the theory in the formation dynamics field into the re-orbiting interference orbit control problem. The single-satellite or multi-satellite re-orbiting interference orbit control is realized with the effective vertical baseline as the evaluation index, the calculation amount is reduced, and it is expected to be applied to the orbit control of actual re-orbiting interference satellites.

[0007] To solve the above problems, the technical scheme of the present application is as follows:

[0008] The single-satellite or multi-satellite re-orbiting interference orbit control method based on the effective vertical baseline comprises the following steps:

[0009] S1, a reference-observation virtual formation model of re-orbiting interference is established based on the eccentricity vector and the relative inclination, and the motion relationship of the slave satellite relative to the master satellite is determined;

[0010] S2, the relationship between the target detection region latitude and the effective vertical baseline is obtained based on the geometric relationship between the target detection region latitude and the subsatellite point latitude and the orbit elements, and the virtual formation effective vertical baseline of the reference-observation satellite is determined;

[0011] S3, the in-plane / out-of-plane orbit control strategy suitable for single-satellite or multi-satellite re-orbiting interference is obtained with the effective vertical baseline as the evaluation index.

[0012] As one aspect of the present application, in step S1, the method for establishing the reference-observation virtual formation model of re-orbiting interference is:

[0013] S1-1, a reference-observation satellite virtual formation is introduced, the reference-observation satellite virtual formation is composed of an observation satellite and a reference satellite, the observation satellite is the satellite position of the current orbit, the reference satellite is the satellite position of the previous orbit, the vertical baseline configuration is obtained and maintained through relative orbit control means, and the reference satellite is taken as the master satellite and the observation satellite is taken as the slave satellite;

[0014] S1-2, the reference-observation virtual formation model is described in the orbit plane with the eccentricity vector, and in the vertical orbit plane with the relative inclination; the reference-observation virtual formation model of re-orbiting interference is obtained;

[0015] The method for determining the motion relationship of the slave satellite relative to the master satellite is:

[0016] On the basis of the reference-observation virtual formation model, an RTN coordinate system is established to analyze the motion state of the satellite from three directions of radial direction R, normal direction T and trace direction N, wherein the radial direction R is from the center of the earth to the satellite, the normal direction T is the direction of the angular momentum of the satellite, the trace direction N is the direction of the motion of the satellite, and the relationship among the radial direction R, the normal direction T and the trace direction N is T=N×R,

[0017] The deviation of the actual position of the observation satellite from the reference position of the last round orbit is set, that is, the expression formula of the motion relationship of the satellite relative to the primary satellite is:

[0018] Δr=r Obser -r Refer =r R e R +Δr N e N +Δr T e T ,

[0019] Wherein, e R is the unit vector of the radial direction R in the RTN coordinate system, e T is the unit vector of the normal direction T in the RTN coordinate system, e N is the unit vector of the radial direction R in the RTN coordinate system, Δr is the relative position vector, r Obser is the reference satellite position vector, r Refer is the observation satellite position vector, Δr R is the R-direction component of the relative position vector, Δr N is the N-direction component of the relative position vector, and Δr T is the T-direction component of the relative position vector.

[0020] Description: Step S1-2 establishes the relative eccentricity / inclination vector to clearly define the design variable of the dynamic model formula, that is, to describe the right side of the dynamic model formula; and step S1-3 establishes the RTN coordinate system to project the baseline vector to three component directions, that is, to describe the left side of the dynamic model formula.

[0021] As one aspect of the present application, in step S1-2, the method for describing the reference-observation virtual formation model by using the eccentricity vector in the orbit plane is:

[0022] In the orbit plane, the eccentricity vector is used to describe the in-orbit plane model, wherein Δe is the eccentricity vector, Δe X is the X component of the relative eccentricity, Δe Y is the Y component of the relative eccentricity, δe is the relative eccentricity, is the sub-satellite point latitude,

[0023] The method for describing the reference-observation virtual formation model in the relative inclination in the vertical orbital plane is:

[0024] The relative inclination vector Δi is described in the vertical orbital plane, The vertical orbital plane model is described, wherein Δi is the relative inclination vector Δi x is the X component of the relative inclination, Δi y is the Y component of the relative inclination, θ is the phase of the relative inclination vector, ΔΩ is the relative right ascension of the ascending node, δi is the relative orbital inclination, and i is the orbital inclination.

[0025] Description: The heavy rail interference geometrically includes ascending rail-ascending rail, ascending rail-descending rail, descending rail-descending rail, and descending rail-ascending rail. If the four combinations of ascending and descending rails of multiple satellites are considered, the calculation amount of the track control of the multiple satellite heavy rail interference will be further increased. The relative orbit position of the reference-virtual satellite is only needed to be discussed when the heavy rail interference is discussed by using the reference-observation satellite virtual formation, which is convenient in geometry. Meanwhile, the relative eccentricity / inclination vector is used to describe the reference-observation virtual formation dynamics model, which can effectively avoid the singularity of the near-circular orbit.

[0026] As an aspect of the present application, based on the geometric relationship between the target detection area latitude and the subsatellite point latitude and the orbit elements, the relationship between the target detection area latitude and the effective vertical baseline is obtained, and the effective vertical baseline of the reference-observation satellite virtual formation is determined, including the following steps:

[0027] S2-1, a dynamics model of the internal relative motion of the virtual formation is established;

[0028] S2-2, based on the geometric relationship between the subsatellite point latitude and the orbit elements, the relationship between the target detection area latitude and the effective vertical baseline is obtained, so as to obtain the design parameters related to the effective vertical baseline, the design parameters including the relative semi-major axis, the relative eccentricity and the relative orbital inclination, and the effective vertical baseline of the reference-observation satellite virtual formation is determined based on the influence law of the design parameters on the spatial latitude distribution of the effective vertical baseline.

[0029] Description: The relative position described by the relative eccentricity / inclination vector established in step S1-3 is decomposed into the RTN direction, which is convenient for designing the track control strategy for the effective vertical baseline design parameters.

[0030] As an aspect of the present application, the model formula of the dynamics model is:

[0031]

[0032] In the above formula, Δr Ris the R-component of the position vector of the satellite relative to the position vector of the main satellite, a is the semi-major axis of the orbit, Δa is the relative semi-major axis, δe is the relative eccentricity, δi is the relative inclination, u is the latitude amplitude, is the sub-satellite latitude, Δr T is the T-component of the position vector of the satellite relative to the position vector of the main satellite, Δr N is the N-component of the position vector of the satellite relative to the position vector of the main satellite.

[0033] Description: The relative position described by the relative eccentricity / inclination vector is decomposed into the RTN directions, which is beneficial to design the orbit control strategy aiming at the effective vertical baseline design parameters.

[0034] As an aspect of the present application, in step S2-1, the geometric relationship between the sub-satellite latitude and the orbital elements is:

[0035]

[0036] In the above formula, is the sub-satellite latitude, u is the latitude amplitude, and i is the orbital inclination.

[0037] Description: According to the spherical triangle formed by the satellite, the ascending node of the orbit, and the intersection of the meridian and the equator, the above geometric relationship exists.

[0038] As an aspect of the present application, the relationship is:

[0039]

[0040] In the above formula, r γ is the effective vertical baseline, r R is the R-component of the baseline, γ1 is the side-looking angle of the satellite-borne SAR, r N is the N-component of the baseline, Δa is the relative semi-major axis, a is the semi-major axis of the orbit, δe is the relative eccentricity, δi is the relative inclination, u is the latitude amplitude, is the sub-satellite latitude.

[0041] Description: The baseline of the virtual formation is decomposed into the RTN coordinate system, and the relationship between the design parameters and the effective vertical baseline is established.

[0042] As an aspect of the present application, the influence of the design parameters on the spatial latitude distribution of the effective vertical baseline is: the distribution is centrally symmetric in the ascending and descending directions, forming a spindle-shaped spatial distribution image; the relative semi-major axis causes the distribution of the effective vertical baseline to drift, but the distribution pattern remains unchanged; the relative eccentricity causes the distribution of the effective vertical baseline to be longitudinally deformed, but the center of the distribution remains unchanged; the relative inclination causes the distribution of the effective vertical baseline to be transversely deformed, but the center of the distribution remains unchanged.

[0043] Description: The design parameters relative semi-major axis, relative eccentricity, relative orbital inclination respectively produce the influence of central drift, longitudinal or transverse deformation on the spatial distribution of effective vertical baseline, and the subsequent orbit control strategy can be designed according to the influence.

[0044] As an aspect of the present application, the in-plane / out-of-plane orbit control strategy suitable for single satellite, multi-satellite heavy orbit interference is obtained based on the effective vertical baseline as the evaluation index, including the following steps:

[0045] S3-1, after considering the heavy orbit interference into the reference-observation virtual formation model, 2n satellite heavy orbit / formation constellation is formed by n satellite heavy orbit interference, and baseline, baseline includes baseline, n baseline heavy orbit interference is formed by the same satellite heavy orbit at different times and spaces; baseline, n baseline heavy orbit interference is formed by the same satellite heavy orbit at different times and spaces;

[0046] S3-2, the orbit control impulse of n satellite heavy orbit interference is calculated.

[0047] Description: Through the cyclic algorithm, the in-plane and out-of-plane n satellite orbit control strategy for different design parameters is designed with the effective vertical baseline as the design target.

[0048] As an aspect of the present application, the orbit control impulse of n satellite heavy orbit interference includes the following cyclic steps:

[0049] S3-2-1, baseline determination: determine whether the baseline of n satellite heavy orbit interference meets the effective vertical baseline length requirement, and the effective vertical baseline length requirement is whether the baseline length is the maximum or minimum value, if it meets, directly output the orbit control strategy formed in the last round, wherein, if it is the initial cycle, the orbit control strategy is not to apply the orbit control, otherwise enter the next step;

[0050] S3-2-2, select the baseline with the maximum and minimum length, determine the baseline to be changed and the satellite to be applied with orbit control;

[0051] S3-2-3, calculate the orbit control impulse: derive the in-plane / out-of-plane orbit control impulse calculation formula of the orbit plane according to the design parameters, and calculate the orbit control impulse of this round through the in-plane / out-of-plane orbit control impulse calculation formula, wherein,

[0052] The in-plane / out-of-plane orbit control impulse calculation formula based on changing the relative semi-major axis is:

[0053]

[0054] The in-plane / out-of-plane orbit control pulse calculation formula based on changing the relative eccentricity is:

[0055]

[0056] The in-plane / out-of-plane orbit control pulse calculation formula based on changing the relative orbital inclination is:

[0057]

[0058] In the above formula, is the trace orbit control double pulse, u is the latitude amplitude angle, u1 is the corresponding first orbit control position latitude amplitude angle, u2 is the corresponding second orbit control position latitude amplitude angle, Δr γ is the effective vertical baseline deviation, αa is the relative semi-major axis, a is the orbit semi-major axis, γ is the spaceborne SAR side-looking angle, μ is the earth gravity constant, Δe X is the X component of the relative eccentricity, Δe Y is the Y component of the relative eccentricity, δe is the relative eccentricity, is the subsatellite point latitude, Δv N is the normal orbit control pulse, δi is the relative orbital inclination;

[0059] S3-2-4, the in-plane / out-of-plane orbit control pulse quantity calculated in S3-2-3 is applied to the satellite and baseline determined in step S3-2-2, the determined orbit control pulse quantity is applied, a new round of baseline is obtained, and in the case of a single satellite, the orbit control strategy corresponding to the design parameters in the in-plane / out-of-plane case is obtained; in the case of multiple satellites, a mathematical induction relationship between the number of re-orbit satellites and the number of formation baselines is established, and the orbit control strategy applicable to multiple satellites is obtained; the in-plane / out-of-plane orbit control strategy of single / multiple satellite re-orbit interference is established; return to step S3-2-1.

[0060] Explanation: The above algorithm takes the effective vertical baseline as the design target, and can simplify the algorithm process; the above algorithm gives the in-plane and out-of-plane orbit control strategy for different design parameters, and can give different orbit control strategies according to the actual on-orbit task.

[0061] The beneficial effects of the present application are:

[0062] (1) In the aspect of model establishment, the re-orbit interference virtual formation orbit control method proposed in the present application converts the complex low / high orbit repeat orbit geometry problem into a perfect formation dynamics and control problem, simplifies the establishment of the model, reduces the calculation amount, and accelerates the orbit control calculation process;

[0063] (2) In the aspect of application scope, the orbit control calculation method proposed in the present application discusses the baseline law starting from n satellites, is applicable to single / multiple satellites, re-orbit and formation mixed multiple re-orbit interference, and has a wide application scope;

[0064] (3) In terms of orbit control consumption, the orbit control design method proposed in the application has a small order of magnitude of orbit control impulse, and the heavy orbit interference configuration can be continuously maintained, so the consumed orbit control fuel is less;

[0065] (4) The application realizes heavy orbit interference baseline design, single satellite or multi-satellite heavy orbit interference orbit control strategy design and other contents, can be applied to the optimization design of a new generation of microwave imaging satellite constellation formation configuration, the development of orbit control implementation level, can further support the implementation of space-borne synthetic aperture radar interference technology, realize all-weather, all-day, large-scale imaging and information acquisition, contribute to the field of global terrain mapping, tree height inversion, soil moisture inversion, global forestry research, can provide new methods and new directions for establishing China's remote sensing system, and can also be applied to space-based situation awareness, support the improvement of China's space-based attack and defense system. BRIEF DESCRIPTION OF DRAWINGS

[0066] Figure 1 is a single satellite or multi-satellite heavy orbit interference orbit control method flowchart based on the effective vertical baseline of the embodiment of the application;

[0067] Figure 2 is a thermal map of the effective vertical baseline length of the heavy orbit interference reference-observation virtual satellite constellation with respect to deltae-deltai in the embodiment of the application;

[0068] Figure 3 is a reference satellite-observation satellite virtual constellation configuration of heavy orbit interference in the embodiment of the application;

[0069] Figure 4 is a graph of the relationship between the effective vertical baseline length of the heavy orbit interference virtual constellation and the subsatellite point latitude in the embodiment of the application;

[0070] Figure 5 is a graph of the time variation relationship between the effective vertical baseline length of the heavy orbit interference virtual constellation and the effective vertical baseline length in the embodiment of the application;

[0071] Figure 6 is a graph of the influence of changing the relative semi-major axis on the baseline spatial distribution in the embodiment of the application;

[0072] Figure 7 is a graph of the influence of changing the relative eccentricity on the baseline spatial distribution in the embodiment of the application;

[0073] Figure 8 is a graph of the influence of changing the relative orbit inclination on the baseline spatial distribution in the embodiment of the application;

[0074] Figure 9 is an algorithm logic flowchart of step S3-2 in the embodiment of the application;

[0075] Figure 10is the configuration change diagram of the single-satellite heavy orbit interference relative semi-major axis orbit control strategy in the application example 1 of the present application;

[0076] Figure 11 is the effective vertical baseline space change diagram of the single-satellite heavy orbit interference relative semi-major axis orbit control strategy in the application example 1 of the present application;

[0077] Figure 12 is the effective vertical baseline time change diagram of the single-satellite heavy orbit interference relative semi-major axis orbit control strategy in the application example 1 of the present application;

[0078] Figure 13 is the configuration change diagram of the single-satellite heavy orbit interference relative eccentricity orbit control strategy in the application example 1 of the present application;

[0079] Figure 14 is the effective vertical baseline space change diagram of the single-satellite heavy orbit interference relative eccentricity orbit control strategy in the application example 1 of the present application;

[0080] Figure 15 is the effective vertical baseline time change diagram of the single-satellite heavy orbit interference relative eccentricity orbit control strategy in the application example 1 of the present application;

[0081] Figure 16 is the multi-satellite heavy orbit interference reference-observation virtual formation baseline schematic diagram in the application example 2 of the present application, n = 2;

[0082] Figure 17 is the multi-satellite heavy orbit interference reference-observation virtual formation baseline schematic diagram in the application example 2 of the present application, n = 3;

[0083] Figure 18 is the multi-satellite heavy orbit interference reference-observation virtual formation baseline schematic diagram in the application example 2 of the present application, n = 4;

[0084] Figure 19 is the baseline-latitude space change diagram of the multi-satellite heavy orbit interference effective vertical baseline in the application example 2 of the present application;

[0085] Figure 20 is the time change diagram of the multi-satellite heavy orbit interference effective vertical baseline in the application example 2 of the present application. DETAILED DESCRIPTION

[0086] In order to make the objects, technical solutions and advantages of the present application clearer, the present application will be further described in detail below with reference to the drawings. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by a person of ordinary skill in the art without creative work fall within the protection scope of the present application.

[0087] The terminology used in the description of the application herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the application. As used in the description of the application and the appended claims, the singular forms "a", "an" and "the" are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be understood that the terms "comprises" and / or "comprising," when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and / or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or groups thereof.

[0088] It is to be understood that, although the terms first, second, third, etc. can be used herein to describe various..., these... should not be limited to these terms. These terms are only used to distinguish one... from another. For example, a first... could be termed a second..., and, similarly, a second... could be termed a first..., without departing from the scope of the present application.

[0089] The existing control of the satellite operating state is based on the flight state of a single satellite, and can guarantee the general requirements of the time and spatial position of revisit. However, the research on the orbit control strategy of single-satellite repeat-track interference is relatively insufficient at home and abroad at present.

[0090] In order to solve the above problems, the present embodiment proposes a single-satellite or multi-satellite repeat-track interference orbit control method based on an effective vertical baseline, as shown in the following figure, which comprises the following steps: Figure 1

[0091] S1, based on the eccentricity vector and the relative inclination, a reference-observation virtual formation model of repeat-track interference is established, and the motion relationship of the slave satellite relative to the master satellite is determined;

[0092] It can be understood that the repeat-track interference forms a configuration in at least two revisits, and the satellite position has a higher requirement in relative imaging. The control method based on a single satellite is difficult to meet the accuracy of the baseline and the requirement of the configuration, and therefore the "virtual relative configuration" of the reference satellite-observation satellite needs to be used to complete the design. Under the existing control condition of the repeat-track observation satellite, the satellite can already achieve close-range motion relative to the reference orbit. The close-range relative motion of the virtual formation is described from the direction of the orbit plane and the direction perpendicular to the orbit plane.

[0093] It can be understood that in step S1, the method for establishing the reference-observation virtual formation model of repeat-track interference is as follows:

[0094] S1-1, in order to realize the repeat-track interference orbit control, a reference-observation satellite virtual formation is introduced. The reference-observation satellite virtual formation is composed of an observation satellite and a reference satellite. The observation satellite is the satellite position of the current orbit, and the reference satellite is the satellite position of the previous orbit. The vertical baseline configuration is obtained and maintained through relative orbit control means, and the reference satellite is taken as the master satellite, and the observation satellite is taken as the slave satellite.

[0095] ​S1-2, describing the reference-observation virtual formation model in the orbit plane with eccentricity vector, and describing the reference-observation virtual formation model in the vertical orbit plane with relative inclination; obtaining the reference-observation virtual formation model of the heavy orbit interference;

[0096] The method for determining the motion relationship of the satellite relative to the main satellite is:

[0097] On the basis of the reference-observation virtual formation model, an RTN coordinate system is established to analyze the motion state of the satellite from the radial direction R of the satellite flight, the normal direction T of the satellite angular momentum, and the trace direction N of the satellite motion direction, wherein the radial direction R is directed from the center of the earth to the satellite, the normal direction T is the direction of the satellite angular momentum, and the trace direction N is the direction of the satellite motion, and the relationship among the radial direction R, the normal direction T, and the trace direction N is T=N×R,

[0098] The actual position of the observation satellite is set to deviate from the reference position of the last round of orbit, that is, the expression formula of the motion relationship of the satellite relative to the main satellite is:

[0099] Δr=r Obser -r Refer =r R e R +Δr N e N +Δr T e T ,

[0100] wherein, wherein, e R is the unit vector of the radial direction R in the RTN coordinate system, e T is the unit vector of the normal direction T in the RTN coordinate system, e N is the unit vector of the radial direction R in the RTN coordinate system, Δr is the relative position vector, r Obser is the reference satellite position vector, r Refer is the observation satellite position vector, Δr R is the relative position vector R component, Δr N is the relative position vector N component, and Δr T is the relative position vector T component;

[0101] It can be understood that in step S1-2, the method for describing the reference-observation virtual formation model in the orbit plane with eccentricity vector is:

[0102] In the orbit plane, the eccentricity vector is used to describe the orbit plane model, wherein Δe is the relative eccentricity, Δe X is the X component of the relative eccentricity, Δe Y is the Y component of the relative eccentricity, δe is the relative eccentricity, is the substar latitude,

[0103] The method for describing the reference-observation virtual formation model in the relative inclination angle in the vertical orbital plane is:

[0104] The relative inclination angle vector Δi is described in the vertical orbital plane, wherein Δi is the relative inclination angle vector, and Δi The vertical orbital plane model is described, wherein Δi is the relative inclination angle vector Δi x is the X component of the relative inclination angle, Δi y is the Y component of the relative inclination angle, θ is the phase of the relative inclination angle vector, ΔΩ is the relative right ascension of the ascending node, δi is the relative orbital inclination, and i is the orbital inclination;

[0105] It can be understood that, the step S1-2 establishes the relative eccentricity / inclination vector in order to determine the design variable of the dynamic model formula, that is, to describe the right side of the dynamic model formula; and the step S1-3 establishes the RTN coordinate system in order to project the baseline vector into three component directions, that is, to describe the left side of the dynamic model formula;

[0106] S2, based on the geometric relationship between the target detection area latitude and the subspace point latitude and the orbital elements, obtaining the relationship between the target detection area latitude and the effective vertical baseline, determining the virtual formation effective vertical baseline of the reference-observation satellite;

[0107] It can be understood that, based on the geometric relationship between the target detection area latitude and the subspace point latitude and the orbital elements, obtaining the relationship between the target detection area latitude and the effective vertical baseline, determining the virtual formation effective vertical baseline of the reference-observation satellite includes the following steps:

[0108] S2-1, establishing a dynamic model of the relative motion inside the virtual formation;

[0109] S2-2, based on the geometric relationship between the subspace point latitude and the orbital elements, obtaining the relationship between the target detection area latitude and the effective vertical baseline, thereby obtaining the design parameters related to the effective vertical baseline, the design parameters including the relative semi-major axis, the relative eccentricity and the relative orbital inclination, and determining the virtual formation effective vertical baseline of the reference-observation satellite based on the influence law of the design parameters on the spatial latitude distribution of the effective vertical baseline;

[0110] It can be understood that, based on the above in-plane / out-of-plane virtual formation model, the influence of the orbital element change on the virtual baseline in the RTN coordinate system is obtained from the linear motion equation of the virtual formation satellite

[0111]

[0112] The in-plane / out-of-plane model of the orbital plane described by the relative eccentricity / inclination vector is deformed, and the plane relative latitude amplitude angle Δu is set to 0, to obtain a simplified relative motion equation, that is, the model formula of the dynamic model, and the model formula of the dynamic model is:

[0113]

[0114] In the above formula, ΔrR is the R-direction component of the relative position vector of the satellite to the main satellite, a is the semi-major axis of the orbit, Δa is the relative semi-major axis, δe is the relative eccentricity, δi is the relative orbit inclination, u is the latitude amplitude angle, is the subsatellite latitude, Δr T is the T-direction component of the relative position vector of the satellite to the main satellite, Δr N is the N-direction component of the relative position vector of the satellite to the main satellite.

[0115] That is, the expression of the three-axis components (Δr R , Δr T , Δr N ) of the formation baseline with respect to Δa, δe, δi is obtained, and the three parameters are taken as the baseline design parameters in the embodiment; in the case where the relative semi-major axis Δa of the virtual formation is 0, the relative orbit of the satellite to the main satellite is a three-dimensional ellipse, and the minimum distance between the two-satellite formation is given by min (aδe, aδi); the projection of the relative orbit on the orbit plane is an ellipse, the semi-major axis length along the orbit direction is 2aδe, and the semi-minor axis length along the radial direction is aδe; the projection on the plane perpendicular to the flight direction is an ellipse, the semi-major axis length along the radial direction is max (aδe, aδi), and the semi-minor axis length along the cross-orbit direction is min (aδe, aδi); based on the model formula of the dynamic model, the baseline length variation of the virtual formation can be analyzed.

[0116] It can be understood that, in order to meet the requirements of SAR interferometric measurement on the baseline length and the latitude of the target area, the two-satellite virtual formation flight of the single-satellite re-orbiting interference needs to have flexibility, that is, different formation configurations need to be changed in the satellite life cycle to meet the imaging requirements on different areas. According to the spherical triangle relationship, the geometric relationship between the subsatellite latitude and the orbit elements is obtained as follows:

[0117]

[0118] In the above formula, is the subsatellite latitude, u is the latitude amplitude angle, and i is the orbit inclination.

[0119] It can be understood that, the baseline component perpendicular to the interferometric imaging direction has a huge impact on the interferometric measurement quality, and the average side-looking angle of the satellite-borne SAR interferometric mode is taken as γ1=(γ max +γ min ) / 2, and the relationship of the effective vertical baseline is as follows:

[0120]

[0121] In the above formula, r γ is the effective vertical baseline, and rR R is the baseline component in the radial direction, γ1 is the side-looking angle of the spaceborne SAR, r N N is the baseline component in the normal direction, Δa is the relative semi-major axis, a is the orbit semi-major axis, δe is the relative eccentricity, δi is the relative orbit inclination, u is the latitude amplitude, is the subspace point latitude;

[0122] The baseline analysis related to the heavy rail interference quality can be given according to the virtual formation baseline design variable in the embodiment;

[0123] In the embodiment, the sun-synchronous circular orbit with an orbit height of 500 km is selected, i.e. the eccentricity e = 0, and the orbit semi-major axis a = 6871 km; further, to meet the sun-synchronous orbit condition, the orbit inclination can be calculated by the following formula:

[0124] cosi sso = -0.09892 (a / R E ) 7 / 2 (1-e 2 ) 2 ,

[0125] The orbit inclination i = 97.4° is obtained; the commonly used morning and evening orbit is selected, i.e. the descending node local time T DN = 18h, and the sun's ecliptic longitude is Λ = 280.461 + 0.9856747*MJD; the ascending node right ascension of the target sun-synchronous orbit formation satellite can be calculated as follows:

[0126] Ω = 15*T DN + Λ = 143.309°,

[0127] The effective baseline range is designed to be 200-1000 m, i.e. when the formation satellite runs to the load beam capable of covering the target area, the baseline requirement is within 200-1000 m; the target area is selected to be the longitude range 75°E-98°E and the latitude range 28°N-37°N; the baseline component perpendicular to the interference imaging direction of the observation satellite at the same orbit height, i.e. the effective vertical baseline, is a bivariate function about δe, δi, and the heat map of the effective vertical baseline about the variables δe, δi in the target area 200m γ <1000m range is shown in Figure 2 ; further, the weighted calculation of the average value design configuration satisfying the condition 200m γ <1000m is carried out, and the three-dimensional diagram of the observation satellite-reference satellite virtual formation configuration for the target area latitude range and the normal, trace and radial three views are obtained as shown in Figure 3As shown; since the latitude φ of the nadir point satisfies a certain geometric relationship with the latitude argument u and the orbital inclination i, and the effective vertical baseline is also related to the design variables, the latitude of the nadir point and the length of the effective vertical baseline at the same moment within the simulation period can be used as two coordinate values ​​to establish a spatial relationship diagram as shown. Figure 4 The map also indicates the baseline distribution range corresponding to the latitude range of the target area, from... Figure 4 It can be seen that the upper and lower observation bands that meet the latitude range of the key area are reached during the ascent and descent of the satellites, respectively. When the satellite formation enters the latitude range of the target key area, the corresponding baseline length meets the design requirement of 200-1000m. In addition, a graph showing the change of the effective vertical baseline over time during the simulation period can be generated, as shown below. Figure 5 As shown;

[0128] Understandably, the influence of design parameters on the spatial latitudinal distribution of the effective vertical baseline follows these patterns: It is centrally symmetrically distributed in both ascending and descending directions, forming a spindle-shaped spatial distribution image; the relative semi-major axis causes a center shift in the distribution of the effective vertical baseline, but the shape of the distribution map remains unchanged; the relative eccentricity causes longitudinal deformation in the distribution of the effective vertical baseline, but the center of the distribution map remains unchanged; the relative orbital inclination causes lateral deformation in the distribution of the effective vertical baseline, but the center of the distribution map remains unchanged.

[0129] S3. Using the effective vertical baseline as the evaluation index, we obtain in-plane / out-of-plane orbit control strategies applicable to single-star and multi-star multiple orbit interferometry.

[0130] Understandably, using the effective vertical baseline as an evaluation metric, the in-plane / out-of-plane orbit control strategies applicable to single-satellite and multi-satellite multiple orbit interferometry include the following steps:

[0131] S3-1. After incorporating heavy orbit interferometry into the reference-observation virtual formation model, a 2n-star heavy orbit / formation constellation is formed through n-star heavy orbit interferometry. The heavy orbit / formation constellation then forms... Baseline, The baselines include A single formation interferometric baseline is formed by multiple satellites in formation at the same time and in different spatial locations; n double orbit interferometric baselines are formed by the same satellite in multiple time and space. The array multiple orbit interferometric baseline is formed by multiple satellite array multiple orbits in different times and spaces;

[0132] S3-2. Calculate the orbit control pulses of the n-star interferometer in heavy orbit;

[0133] Understandable, such as Figure 9 As shown, the calculation of the orbit control pulses for the n-star heavy orbit interferometer includes the following iterative steps:

[0134] S3-2-1, baseline determination: determine whether the baseline of n-star heavy rail interference satisfies the effective vertical baseline length requirement, the effective vertical baseline length requirement is whether the baseline length is the maximum or minimum value, if it is satisfied, directly output the orbit control strategy formed in the last round, wherein, if it is the initial cycle, that is, the condition is met, the orbit control strategy is not to apply the orbit control, otherwise, go to the next step;

[0135] S3-2-2, select the baseline with the maximum and minimum length, determine the baseline to be changed and the satellite to apply the orbit control;

[0136] S3-2-3, calculate the orbit control pulse: according to the design parameters, the in-plane / out-of-plane orbit control pulse calculation formula of the orbit plane is derived respectively, and the orbit control pulse of this round is calculated through the in-plane / out-of-plane orbit control pulse calculation formula, wherein,

[0137] The in-plane / out-of-plane orbit control pulse calculation formula based on changing the relative semi-major axis is:

[0138]

[0139] The in-plane / out-of-plane orbit control pulse calculation formula based on changing the relative eccentricity is:

[0140]

[0141] The in-plane / out-of-plane orbit control pulse calculation formula based on changing the relative orbit inclination is:

[0142]

[0143] In the above formula, is the trace orbit control double pulse, u is the latitude amplitude angle, u1 is the corresponding first orbit control position latitude amplitude angle, u2 is the corresponding second orbit control position latitude amplitude angle, Δr γ is the effective vertical baseline deviation, Δa is the relative semi-major axis, a is the orbit semi-major axis, γ is the side-looking angle of the satellite-borne SAR, μ is the earth gravity constant, Δe X is the X component of the relative eccentricity, Δe Y is the Y component of the relative eccentricity, δe is the relative eccentricity, is the sub-satellite point latitude, Δv N is the normal orbit control pulse, δi is the relative orbit inclination;

[0144] S3-2-4, the in-plane / out-of-plane orbit control impulse amount calculated in S3-2-3 is applied to the satellite and the baseline determined in S3-2-2, the determined orbit control impulse amount is applied to obtain a new round of baseline, and in the case of a single satellite, the in-plane / out-of-plane design parameter corresponding orbit control strategy is obtained; in the case of multiple satellites, a mathematical induction relationship between the number of re-orbiting satellites and the number of formation baselines is established to obtain a multi-satellite applicable orbit control strategy; the in-plane / out-of-plane orbit control strategy of single / multi-satellite re-orbiting interference is established according to the baseline law of single / multi-satellite; return to S3-2-1;

[0145] It can be understood that according to the relative motion model of the re-orbiting interference relative configuration and the effective vertical baseline calculation formula, it can be known that Δa, δi, δe are the key parameters affecting the effective vertical baseline, and are also important orbit elements that can be changed by using in-plane or out-of-plane orbit control impulse; based on Figure 4 、 Figure 5 The virtual formation configuration given by the relative orbit elements of the space-time relationship is added with Δa, δi, δe bias amount to obtain the effective vertical baseline-subsatellite point latitude distribution graph with different design parameter bias as shown in Figure 6 、 7 、8; it can be known from Figure 6 、 7 、8 that the relative semi-major axis Δa causes the center of the effective vertical baseline distribution to drift, but the distribution pattern does not change; the relative eccentricity δe causes the longitudinal deformation of the effective vertical baseline distribution, but the center of the distribution graph does not change; the relative inclination δi causes the transverse deformation of the effective vertical baseline distribution, but the center of the distribution graph does not change; further, from the three design variables, the orbit control strategy of single / multi-satellite re-orbiting interference can be designed in subsequent application examples.

[0146] Application Example 1: The application example is an application process of designing a single-satellite re-orbiting interference orbit control strategy based on the effective vertical baseline orbit control method of the embodiment, including:

[0147] Based on the orbit control algorithm flow, the in-plane orbit control strategy of single-satellite re-orbiting interference is designed, in order to realize the relative orbit control of the main star passive and the slave star maneuvering; the classic orbit control method starts from the perturbation equation, linearizes the relative motion model to obtain the approximate circular orbit element change equation set generated by the trace Δv T , radial Δv R and normal Δv N velocity impulse:

[0148]

[0149] The effective vertical baseline is changed by default using a single orbit root number, and the in-plane orbit control consumption orbit control pulse (double pulse) and corresponding latitude amplitude angle are calculated by combining the effective vertical baseline calculation formula and the near-circular orbit relative motion control equation.

[0150] According to the in-plane / out-of-plane orbit control pulse calculation formula, an in-plane orbit change strategy using a reference-observation satellite virtual formation relative semi-major axis is designed, and the specific orbit root number values are still as in the embodiment; Matlab and satellite simulation software are used for joint debugging, and the reference satellite (Sat_Refer) and the observation satellite (Sat_Observe) are given; the orbit root numbers of the two satellites in the initial state are the same, but the reference satellite leads the observation satellite by one orbit period; the observation satellite (Sat_Observe) is respectively subjected to the in-track pulse at the latitude amplitude angle with a phase difference of 180°, and the change of the virtual formation relative orbit position is obtained, as shown in Figure 10 ; wherein the circle represents the reference satellite, which is the main star of the virtual formation; the trajectory line is the reference trajectory of the observation satellite to the reference satellite; the target baseline is taken as 200 m, 500 m, and 1000 m respectively, and three relative trajectories in the figure are obtained, and the consumed orbit control pulse and the effective vertical baseline length before and after the orbit change are shown in Table 1; at the same time, the spatial distribution (changes with the sub-satellite latitude) and the time distribution of the effective vertical baseline are shown in Figure 11 and Figure 12 ;

[0151] Table 1 Orbit control amount required for adjusting the in-plane relative semi-major axis, effective vertical baseline length before and after the orbit change

[0152]

[0153] It can be analyzed from Figure 10 , Figure 11 , Figure 12 that the relative semi-major axis is changed by in-plane double pulse orbit control, the effective vertical baseline is constant (the same as the change of the relative semi-major axis), and the re-rail interference effect is better for a specific latitude area; on the other hand, from the perspective of orbit control efficiency, it can be known from the parameters in the relative motion control equation that the orbit control pulse has the highest change efficiency for the relative semi-major axis, and has better performance in fuel saving;

[0154] An in-plane orbit change strategy using a reference-observation satellite virtual formation relative eccentricity is designed, and Matlab and satellite simulation software are used for joint debugging to simulate the virtual formation composed of the reference and observation satellites;

[0155] The change of the virtual formation relative orbit position, the spatial distribution (changes with the sub-satellite latitude) and the time distribution of the effective vertical baseline are obtained, as shown in Figure 13 , Figure 14 , Figure 15 ; Figure 13The middle circle represents a reference satellite, which is the main star of the virtual formation; the trajectory line is the reference trajectory of the observation satellite to the reference satellite; three relative trajectories are obtained by taking the target baseline as 200 m, 500 m and 1000 m respectively; the consumed orbit control impulse and the effective vertical baseline length before and after orbit transfer are shown in Table 2;

[0156] Table 2 Orbit control amount required for adjusting the in-plane eccentricity, effective vertical baseline length before and after orbit transfer

[0157]

[0158] By Figure 13 、 Figure 14 、 Figure 15 It can be analyzed that by changing the relative eccentricity through in-plane double-pulse orbit control, the virtual formation forms a new effective vertical baseline spatial distribution, and only the target baseline can be achieved near the target latitude area, and the control accuracy is not as good as the orbit control strategy of changing the relative semi-major axis. Under this orbit control strategy, the virtual formation configuration is stable;

[0159] Combined with the effective vertical baseline calculation formula (the relationship between the target detection area latitude and the effective vertical baseline) and the relative motion control equation of the near-circular orbit, the orbit control impulse consumed by the out-of-plane orbit control can be calculated, and the formula for changing the relative orbit inclination is:

[0160]

[0161] Since the change in inclination is a plane change, the required orbit control impulse is larger than that of in-plane orbit control, and the fuel consumption is larger; in the case of having a certain side view angle (i.e. sinγ≠0), the efficiency of in-plane orbit control is much higher than that of out-of-plane orbit control; since the side view angle of the heavy orbit SAR is mostly 20-45°, there is no case of side view angle of 0, therefore, it is not recommended to adopt the out-of-plane orbit control strategy.

[0162] Application Example 2: The application example is an application process of designing a multi-satellite heavy orbit interferometric orbit control strategy based on the effective vertical baseline orbit control method of the satellite or the multi-satellite heavy orbit interferometric orbit control method of the embodiment, including:

[0163] Figure 16 、 Figure 17 and Figure 18 The baseline conditions of double-satellite, triple-satellite and four-satellite heavy orbit interference are given. It is worth noting that, Figure 16 、 Figure 17 and Figure 18Only for reference - observation orbit reference figure is also the same as the ascending orbit, there are three cases of ascending orbit - descending orbit, descending orbit - descending orbit, descending orbit - ascending orbit. This also reflects the convenience of discussing heavy orbit interference by adopting reference-observation satellite virtual formation: if considering the four combinations of ascending and descending orbits of multiple heavy orbit interference satellites, the baseline case will be further increased; and the relative orbital position of the virtual formation does not need to consider the geometric configuration of the heavy orbit interference;

[0164] Taking the two-star heavy orbit interference of n = 2 of n-star heavy orbit interference as an example, the baseline and orbit control strategy design of multiple heavy orbit interference are discussed; first, the satellite codes SatRfC, SatRfD, SatObC, SatObD are assigned to the reference master star, the reference slave star, the observation master star, and the observation slave star respectively, the virtual formation interference type, baseline definition, and the satellite that applies the orbit control pulse of the four-star virtual formation interference are shown in Table 3; in order to unify the algorithm before and after, first, the reference and observation master stars are taken as the master stars between the two satellites, and only in the case of reference slave star-observation slave star, the reference slave star is taken as the master star between the two satellites;

[0165] Table 3 Related definitions of reference-observation virtual formation of two-star heavy orbit interference

[0166]

[0167]

[0168] Since the two-star heavy orbit interference forms a four-star virtual formation and six formation baselines, for the effective vertical baseline range of 200m-1000m, we define the following two cases that meet the baseline range:

[0169] (1) The best case of two-star heavy orbit interference: all baselines of two-star heavy orbit interference or formation meet the baseline requirements, all six formation baselines can meet the baseline range of a specific latitude area, described in mathematical language

[0170] 200m<min(BL1,…,BL6)<max(BL1,…,BL6)<1000m

[0171] (2) The limit case of two-star heavy orbit interference: any baseline of two-star heavy orbit interference or formation meets the baseline requirements, only part of the formation baselines meet the baseline range of a specific latitude area, described in mathematical language 200m<max(BL1,…,BL6)<1000m or 200m<min(BL1,…,BL6)<1000m

[0172] Considering that the goal of multiple heavy orbit interference is to increase more available baseline links, this application example adopts case (1) to design the orbit control scheme of multiple heavy orbit interference, and here the formation configuration given in the embodiment is taken as the orbit root number input;

[0173] According to the cycle process of step S3-2 in the embodiment, the orbit control strategy under different integral models is given. Under a given integral model (which can be Earth Point Mass, Earth J2, Earth HPOP Default V10, etc.), the following algorithm steps are executed:

[0174] STEP1 (baseline determination) determines whether the baseline of the multi-satellite orbit interference meets the requirements. If it meets the requirements, the algorithm is ended and the orbit control strategy is output. If it does not meet the requirements, STEP2 is entered.

[0175] STEP2 (orbit control satellite selection) If the baseline requirements are not met, the maximum and minimum values of the baseline are selected, and the satellite that applies the orbit control is determined according to Table 3.5.

[0176] STEP 3 (baseline selection) The maximum and minimum values of the baseline required to achieve the target baseline are calculated, and the smaller value is selected as the baseline that needs to be changed.

[0177] STEP 4 (orbit control pulse calculation) For the selected baseline and satellite, the orbit control pulse determined by (24), (25) or (26) is applied, and a new round of baseline is obtained, which is transferred to STEP1.

[0178] The target baseline is set to 500m, and the baseline under different integral models after applying orbit control is obtained by using the above algorithm as shown in Table 4.

[0179] Table 4 Final baseline of double-satellite orbit interference under different integral models

[0180]

[0181] By changing the size of the target baseline, the pulse size under different conditions is obtained as shown in Table 5.

[0182] Table 5 Final baseline of double-satellite orbit interference under different integral models and different target baselines

[0183]

[0184] In the example given in this application example, the spatial and temporal changes of the baseline of multi-satellite orbit interference under the Earth Point Mass integral model are as follows: Figure 19 、 Figure 20As shown in the figure, the space-time changes of the three baselines of simple formation, simple re-rail and formation re-rail are different, which can be applied to different interference scenarios; thus, the application example forms an in-plane rail control strategy for multi-satellite re-rail interference, and since multi-satellite re-rail interference involves many factors, the application example only gives the case of double-satellite re-rail interference, in-plane rail control and relative semi-major axis orbit change; the application realizes re-rail interference baseline design, single-satellite or multi-satellite re-rail interference orbit control strategy design and the like, and can be applied to the optimization design of a new generation of microwave imaging satellite constellation formation configuration. The development of the orbit control implementation level can further support the implementation and landing of the space-borne synthetic aperture radar interference technology, realize all-weather, all-time and large-range imaging and information acquisition, and contribute to the fields of global terrain mapping, tree height inversion, soil moisture inversion, global forestry research and the like. The application can provide new methods and new directions for establishing a remote sensing system of China, and can also be applied to space-based situation awareness, and support the improvement of the space-based attack and defense system of China.

Claims

1. A single-satellite or multi-satellite re-orbit interferometry orbit control method based on an effective vertical baseline, characterized in that, Includes the following steps: S1. Based on the eccentricity vector and relative inclination angle, a reference-observation virtual formation model for multiple orbit interferometry is established to determine the motion relationship between the slave star and the master star. The method for establishing the reference-observation virtual formation model for multiple orbit interferometry is as follows: S1-1. Introduce a virtual formation of reference-observation satellites. The virtual formation of reference-observation satellites consists of observation satellites and reference satellites. The observation satellites are the satellite positions of the current orbit, and the reference satellites are the satellite positions of the previous orbit. The vertical baseline configuration is obtained and maintained through relative orbit control. The reference satellites are used as the master satellites, and the observation satellites are used as slave satellites. S1-2. Describe the reference-observation virtual formation model in the orbital plane using the eccentricity vector, and describe the reference-observation virtual formation model in the vertical orbital plane using the relative inclination angle; thus obtaining the reference-observation virtual formation model for heavy orbit interferometry. The method for determining the motion relationship of the secondary star relative to the primary star is as follows: Based on the reference-observation virtual formation model, an RTN coordinate system is established to analyze the satellite's motion state from three directions: radial (R), normal (T), and track (N). Radial (R) points from the Earth's center to the satellite, normal (T) points from the direction of the satellite's angular momentum, and track (N) points to the direction of the satellite's motion. The relationship between radial (R), normal (T), and track (N) is T = N × R. The deviation between the actual position of the observed satellite and the reference position of the previous orbit is set, and the formula for expressing the motion relationship between the satellite and the primary satellite is as follows: Δr=r Obser -r Refer =Δr R in R +Δr N in N +Δr T in T , Among them, e R e is the unit vector of radial direction R in the RTN coordinate system. T e is the unit vector normal to T in the RTN coordinate system. N Let r be the unit vector of radial direction R in the RTN coordinate system, and Δr be the relative position vector. Obser For the reference satellite position vector, r Refer To observe the satellite's position vector, Δr R Let Δr be the R-axis component of the relative position vector. N For the N-axis component of the relative position vector, Δr T The relative position vector T-axis component; S2. Based on the geometric relationship between the latitude of the target detection area and the latitude of the nadir point and the orbital elements, the relationship between the latitude of the target detection area and the effective vertical baseline is obtained, and the effective vertical baseline of the virtual formation of the reference-observation satellites is determined; including the following steps: S2-1. Establish a dynamic model of the relative motion within the virtual formation; S2-2. Based on the geometric relationship between the latitude of the sub-satellite point and the orbital elements, the relationship between the latitude of the target detection area and the effective vertical baseline is obtained, thereby obtaining the design parameters related to the effective vertical baseline. The design parameters include the relative semi-major axis, relative eccentricity and relative orbital inclination. Based on the influence of the design parameters on the spatial latitude distribution of the effective vertical baseline, the effective vertical baseline of the virtual formation of the reference-observation satellite is determined. S3. Using the effective vertical baseline as the evaluation index, obtain in-plane / out-of-plane orbit control strategies applicable to single-satellite and multi-satellite re-orbit interferometry; including the following steps: S3-1. After incorporating heavy orbit interferometry into the reference-observation virtual formation model, a 2n-star heavy orbit / formation constellation is formed through n-star heavy orbit interferometry. The heavy orbit / formation constellation then forms... Baseline, The baselines include A single formation interferometric baseline is formed by multiple satellites in formation at the same time and in different spatial locations; n double orbit interferometric baselines are formed by the same satellite in multiple time and space. The array multiple orbit interferometric baseline is formed by multiple satellite array multiple orbits in different times and spaces; S3-2. The orbit control pulses of the n-star interferometer in heavy orbit are calculated.

2. The single-satellite or multi-satellite multiple orbit interferometry orbit control method based on an effective vertical baseline as described in claim 1, characterized in that, In steps S1-2, the method for describing the reference-observation virtual formation model in the orbital plane using an eccentricity vector is as follows: In the orbital plane, the eccentricity vector is used. Describe the model within the orbital plane, where Δe is the eccentricity vector. X Let X be the component of the relative eccentricity, and Δe Y Let Y be the relative eccentricity, and δe be the relative eccentricity. The latitude of the sub-satellite point. The method for describing the reference-observation virtual formation model in the vertical orbital plane with relative inclination angles is as follows: In the vertical orbital plane, with relative inclination vector Describe a vertical track plane model, where Δi is the relative tilt angle vector Δi x Let Δi be the X component of the relative tilt angle. y Let Y be the Y component of the relative inclination angle, θ be the phase of the relative inclination angle vector, ΔΩ be the right ascension of the relative ascending node, δi be the relative orbital inclination angle, and i be the orbital inclination angle.

3. The single-satellite or multi-satellite multiple orbit interferometry orbit control method based on an effective vertical baseline as described in claim 1, characterized in that, The model formula for the dynamic model is: In the above formula, Δr R Let denot δe be the R-component of the position vector of the secondary star relative to the primary star, Δa be the semi-major axis of the orbit, δe be the relative eccentricity, δi be the relative orbital inclination, and u be the latitudinal argument. The latitude of the sub-satellite point, Δr T Let Δr be the T-component of the vector of position of the secondary star relative to the primary star. N Let N be the N-axis component of the position vector of the slave star relative to the master star.

4. The single-satellite or multi-satellite multiple orbit interferometry orbit control method based on an effective vertical baseline as described in claim 1, characterized in that, In step S2-1, the geometric relationship between the sub-satellite latitude and the orbital elements is as follows: In the above formula, denoted by 'u', where 'u' is the latitude of the sub-satellite point, 'i' is the latitude argument, and 'i' is the orbital inclination.

5. The single-satellite or multi-satellite re-orbit interferometry orbit control method based on an effective vertical baseline as described in claim 1, characterized in that, The relationship is as follows: In the above formula, r γ For an effective vertical baseline, r R The baseline is represented by the R-axis component, γ1 is the side view of the spaceborne SAR, and r N Let N be the N-axis component of the baseline, Δa be the relative semi-major axis, a be the orbital semi-major axis, δe be the relative eccentricity, δi be the relative orbital inclination, and u be the latitudinal argument. This is the latitude of the point below the star.

6. The single-satellite or multi-satellite multiple orbit interferometry orbit control method based on an effective vertical baseline as described in claim 1, characterized in that, The influence of the design parameters on the spatial latitudinal distribution of the effective vertical baseline is as follows: it is centrally symmetrically distributed in both ascending and descending directions, forming a spindle-shaped spatial distribution image; the relative semi-major axis causes the distribution of the effective vertical baseline to drift in the center, but the shape of the distribution map remains unchanged; the relative eccentricity causes the distribution of the effective vertical baseline to deform longitudinally, but the center of the distribution map remains unchanged. The relative orbital inclination causes a lateral deformation in the distribution of the effective vertical baseline, but the center of the distribution map remains unchanged.

7. The single-satellite or multi-satellite multiple orbit interferometry orbit control method based on an effective vertical baseline as described in claim 1, characterized in that, The calculation of the orbit control pulses for the n-star heavy orbit interferometer includes the following iterative steps: S3-2-1, Baseline Determination: Determine whether the baseline of the n-star heavy orbit interferometry meets the effective vertical baseline length requirement. The effective vertical baseline length requirement is whether the baseline length is the maximum or minimum value. If it meets the requirement, directly output the orbit control strategy formed in the previous round. If the condition is met in the first loop, the orbit control strategy is to not apply orbit control. Otherwise, proceed to the next step. S3-2-2. Select baselines with the maximum and minimum lengths to determine the baselines that need to be changed and the satellites to be subject to orbit control. S3-2-3, Calculate Track Control Pulses: Based on the design parameters, derive the in-plane / out-of-plane track control pulse calculation formulas for the track plane. Calculate the track control pulses for this round using these formulas. The formula for calculating in-plane / out-of-plane orbit control pulses based on changing the relative semi-major axis is as follows: The formula for calculating in-plane / out-of-plane orbital control pulses based on changing relative eccentricity is as follows: The formula for calculating in-plane / out-of-plane control pulses based on changing the relative orbital inclination angle is as follows: In the above formula, This is a track-guided double pulse, where u is the latitude argument, u1 is the latitude argument of the corresponding first track control position, u2 is the latitude argument of the corresponding second track control position, and Δr is... γ For effective vertical baseline deviation, Δa is the relative semi-major axis, a is the orbital semi-major axis, γ is the side view of the spaceborne SAR, μ is the Earth's gravitational constant, and Δe X Let X be the component of the relative eccentricity, and Δe Y Let Y be the relative eccentricity, and δe be the relative eccentricity. The latitude of the sub-satellite point, Δv N The pulse is the normal orbit control pulse, and δi is the relative orbit inclination angle. S3-2-4. Apply the in-plane / out-of-plane orbit control pulses calculated in S3-2-3 to the satellites and baselines determined in step S3-2-2, apply the determined orbit control pulses to obtain a new round of baselines, and in the case of a single satellite, derive the orbit control strategy corresponding to the design parameters in the case of in-plane / out-of-plane; in the case of multiple satellites, establish a mathematical inductive relationship between the number of satellites in heavy orbits and the number of formation baselines to obtain the orbit control strategy applicable to multiple satellites; based on the baseline pattern from single satellite to multiple satellites, establish the in-plane / out-of-plane orbit control strategy for single / multiple satellite heavy orbit interference; return to step S3-2-1.

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