High-precision interferometric constellation error evolution method based on high-order characteristics

Abstract

The application discloses a high-precision interferometric measurement constellation error evolution method based on high-order characteristics and belongs to the field of space technology. The application realizes the method as follows: setting the orbit states of three spacecrafts in an inertial system, setting the orbit dynamics model and initial state of each spacecraft, selecting the characteristic direction of the initial orbit error of the spacecraft, establishing a high-order model of the characteristic direction, establishing a high-order inverse model of the characteristic direction, setting the initial value of the orbit error evolution semi-analytical expression based on the high-order characteristics, calculating the orbit error evolution semi-analytical expression based on the high-order characteristics, constructing the orbit error evolution semi-analytical expression of the three spacecrafts, constructing a configuration stability index model, establishing a high-precision interferometric measurement constellation error evolution model based on the characteristic direction, giving an initial deviation, calculating the configuration stability index deviation at a given moment, and realizing high-precision interferometric measurement constellation error evolution analysis based on the calculated configuration stability index deviation at the given moment.

Description

TECHNICAL FIELD

[0001] The application belongs to the field of space technology, and relates to a high-precision space interferometric constellation configuration error evolution method based on high-order characteristic directions. BACKGROUND

[0002] Interferometric measurement technology has been one of the important technologies in the field of precision measurement since the nineteenth century, and has a wide range of applications in precision industrial production and processing and basic scientific measurement. In recent decades, with the development of space science application needs, such as space gravitational wave detection, high-precision inter-satellite laser ranging and interferometric measurement technology has been widely valued.

[0003] High-precision inter-satellite laser ranging is to use two or more laser beams between satellites for interference, and to obtain inter-satellite distance change information by reading the phase information of the interference signal. This technology is based on high-precision satellite laser frequency stabilization, precise phase measurement, weak light phase-locked loop technology, inter-satellite laser pointing control and other technologies, and can realize picometer-level inter-satellite displacement measurement, which has important value and application prospect in the fields of space science and technology, basic physics experiment and the like. The existing space high-precision interferometric constellation includes the “Taiji” plan proposed by the Chinese Academy of Sciences, the “Tianqin” plan proposed by the Tianqin Center of Sun Yat-sen University, and the “LISA” plan proposed by the European Space Agency (ESA).

[0004] In the research of space high-precision interferometric constellation, the range of interferometric measurement is affected by the arm length between the measurement devices. Compared with ground-based interferometric measurement, by deploying interferometric measurement spacecraft in space to form a space high-precision interferometric constellation, the measurement range can be significantly improved to realize the measurement of low-frequency signals.

[0005] The space high-precision interferometric constellation usually includes three spacecraft forming an approximate equilateral triangle to form a detection configuration with three laser measurement arms. The space high-precision interferometric constellation requires that the spacecraft cannot be actively controlled during the implementation of the task, because active control will affect the performance of interferometric measurement and cause false measurement. At the same time, the space high-precision interferometric constellation has a very high requirement for its configuration stability, which requires that the three spacecraft form an approximate equilateral triangle in a long-term task process. Although the nominal orbit of the space high-precision interferometric constellation usually meets the configuration stability requirement through optimization design, the spacecraft in the space high-precision interferometric constellation will deviate from the nominal orbit set in advance due to the influence of the in-orbit deviation, and thus the configuration stability of the space high-precision interferometric constellation is destroyed. Therefore, it is necessary to predict the error evolution of the configuration stability in the task design stage, analyze the influence of the in-orbit error on the configuration stability, and thus provide support for the task orbit design.

[0006] It is very important to study the configuration error evolution method of space high-precision interferometric measurement constellation. The existing configuration error evolution method of space high-precision interferometric measurement constellation has low calculation accuracy and large calculation amount, and cannot meet the task requirements, so it is necessary to study a high-precision and fast calculation speed configuration error evolution method of space high-precision interferometric measurement constellation. SUMMARY

[0007] In order to solve the problem of inaccurate prediction in the prior art, the main purpose of the present application is to provide a high-precision interferometric measurement constellation error evolution method based on high-order characteristics, which can quickly and accurately calculate the error evolution of the space high-precision interferometric measurement constellation, calculate the configuration stability index deviation at a given time, and realize the deviation prediction of the configuration stability based on the calculated configuration stability index deviation at the given time. The present application has the advantages of high accuracy and fast speed in constellation error evolution prediction.

[0008] To achieve the above-mentioned purpose, the technical scheme adopted by the present application is as follows:

[0009] The high-precision interferometric measurement constellation error evolution method based on high-order characteristics disclosed by the present application sets the orbit state of three spacecrafts in the space high-precision interferometric measurement constellation in the inertial system, sets the orbit dynamics model and initial state of each spacecraft in the high-precision interferometric measurement constellation, selects the characteristic direction of the initial orbit error of the spacecraft according to the initial state of the spacecraft, establishes a high-order model of the characteristic direction and a high-order inverse model of the characteristic direction, sets the initial value of the orbit error evolution semi-analytical expression based on the high-order characteristics according to the high-order inverse model of the characteristic direction, calculates the orbit error evolution semi-analytical expression based on the high-order characteristics, constructs the orbit error evolution semi-analytical expression of the three spacecrafts in the space high-precision interferometric measurement constellation, constructs the configuration stability index model of the space high-precision interferometric measurement constellation, establishes the high-precision interferometric measurement constellation error evolution model based on the characteristic direction, calculates the configuration stability index deviation at a given time given the initial deviation, and further analyzes the configuration stability based on the calculated configuration stability index deviation at the given time to obtain the configuration stability analysis result. Based on the deviation of the obtained configuration stability and the configuration stability analysis result, the accuracy of the task orbit optimization and orbit reconstruction is improved. The present application has the advantages of high calculation accuracy and fast calculation speed, can be calculated on orbit, improves the implementation efficiency of the space high-precision interferometric measurement constellation task, and reduces the interferometric measurement error.

[0010] The high-precision interferometric measurement constellation error evolution method based on high-order characteristics disclosed by the present application comprises the following steps:

[0011] Step 1: Set the orbit states of three spacecrafts in the high-precision space interferometric constellation in the inertial system as x1(t) = [r1(t); v1(t)], x2(t) = [r2(t); v2(t)], and x3(t) = [r3(t); v3(t)], and set the orbit dynamics model f[x(t), t] and the initial state x0 of each spacecraft in the high-precision space interferometric constellation.

[0012] Set the orbit states of three spacecrafts in the high-precision space interferometric constellation in the inertial system as x1(t) = [r1(t); v1(t)], x2(t) = [r2(t); v2(t)], and x3(t) = [r3(t); v3(t)], where the subscripts 1, 2, and 3 represent spacecraft 1, spacecraft 2, and spacecraft 3. The initial states of the spacecrafts at the initial time t0 are represented as x 1,0 = x1(t0), x 2,0 = x2(t0), and x 3,0 = x3(t0).

[0013] Set the orbit dynamics model f[x(t), t] and the initial state x0 = [r0; v0] of each spacecraft in the high-precision space interferometric constellation as follows:

[0014]

[0015] where f(x, t) is the spacecraft dynamics model, x0 = [r0; v0] is the initial state, r0 is the initial position vector, and v0 is the initial velocity vector. In formula (1), the subscripts 1, 2, and 3 used to represent spacecraft 1, spacecraft 2, and spacecraft 3 are omitted for convenience.

[0016] Step 2: According to the initial state x0 = [r0; v0] of the spacecraft, select the characteristic direction of the initial orbit error of the spacecraft, establish a high-order model of the characteristic direction and calculate its coefficients and According to the established high-order model of the characteristic direction, establish a high-order inverse model of the characteristic direction and calculate the corresponding coefficients and

[0017] According to the initial state x0 = [r0; v0] of the spacecraft, select the characteristic direction of the initial orbit error of the spacecraft as the modulus of the position vector and the velocity vector:

[0018] y0 = [||r0||, ||v0||] T (2)

[0019] where y0represents the characteristic direction. A high-order model of the characteristic direction is further established as follows:

[0020]

[0021] where represents the i-th element of the characteristic direction y0, represents the k1-th element of the initial state x0, represents the k2-th element of the initial state x0, and the symbol δ represents a deviation, and is a tensor, defined as:

[0022]

[0023] A high-order inverse model of the characteristic direction is established in an inverse tensor manner as follows:

[0024]

[0025] where is an inverse tensor of is an inverse tensor of

[0026] Step 3: Setting the initial value of the semi-analytical expression of the orbit error evolution based on the high-order characteristic according to the high-order inverse model of the characteristic direction established in Step 2 and the corresponding coefficients and and

[0027] Step 3: Setting the initial value of the semi-analytical expression of the orbit error evolution based on the high-order characteristic according to the high-order inverse model of the characteristic direction established in Step 2 and the corresponding coefficients and and

[0028]

[0029] Step 4: Calculating the semi-analytical expression of the orbit error evolution based on the high-order characteristic according to the orbit dynamics model f[x(t), t] of the spacecraft in the high-precision interferometric constellation set in Step 1, the initial state x0, and the initial value of the semi-analytical expression of the orbit error evolution based on the high-order characteristic set in Step 3 and and

[0030] ​​​​​The orbit dynamics model f [x(t), t] of the space vehicles in the high-precision interferometric constellation according to step 1, and the initial state x0, the orbit error evolution semi-analytical expression based on high-order characteristics according to step 3 and The orbit error evolution semi-analytical expression based on high-order characteristics is calculated by numerical integration and The derivatives of the expressions and are:

[0031]

[0032] Wherein:

[0033]

[0034] Where f i [x(t), t] represents the i-th element of the orbit dynamics model f [x(t), t].

[0035] Step 5: Based on the orbit error evolution semi-analytical expression based on high-order characteristics calculated in step 4 and The orbit error evolution semi-analytical expressions of three space vehicles in the high-precision interferometric constellation are constructed as follows:

[0036] Based on the orbit error evolution semi-analytical expression based on high-order characteristics calculated in step 4 and The orbit error evolution semi-analytical expressions of three space vehicles in the high-precision interferometric constellation are constructed as follows:

[0037]

[0038] Wherein y 1,0 represents the characteristic direction vector of the first space vehicle, y 2,0 represents the characteristic direction vector of the second space vehicle, y 3,0 represents the characteristic direction vector of the third space vehicle, and are coefficients, defined as:

[0039]

[0040] Step 6: Based on the orbital states of the three spacecraft in the space high-precision interferometry constellation in the inertial frame set in Step 1, namely x1(t)=[r1(t); v1(t)], x2(t)=[r2(t); v2(t)] and x3(t)=[r3(t); v3(t)], construct the configuration stability index model z(t)=h[X(t)].

[0041] The configurational stability parameters considered include arm length l ij (t), breathing angle θ i (t) and the rate of change of arm length The corresponding function expression is as follows:

[0042]

[0043] in

[0044] r ij (t)=r i (t)-r j (t) (18)

[0045] v ij (t)=v i (t)-v j (t) (19)

[0046] make The descriptive vector representing the configuration stability index, let Let the state vectors of the three spacecraft represent the following nonlinear mapping relationship:

[0047]

[0048] Where z(t) = h[X(t)] represents the model of configuration stability index.

[0049] Step 7: Based on the configuration stability index model z(t)=h[X(t)] of the high-precision interferometry constellation constructed in Step 6, and the model from Step 5... and Establish a high-precision interferometric constellation error evolution model based on characteristic directions.

[0050] Based on the configuration stability index model z(t)=h[X(t)] of the high-precision interferometric constellation constructed in step 6, the error evolution model of the high-precision interferometric constellation based on the characteristic direction is established as follows:

[0051]

[0052] in:

[0053]

[0054] Step 8: Based on the high-precision interferometric measurement constellation error evolution model based on the characteristic direction established in step 7, given the initial deviation, calculate the configuration stability index deviation at a given time, and further based on the calculated configuration stability index deviation at a given time, i.e. the deviation of the predicted configuration stability, analyze the configuration stability to obtain the configuration stability analysis result.

[0055] Further comprising step 9: based on the configuration stability deviation and configuration stability analysis result obtained in step 8, improve the accuracy of mission orbit optimization and orbit reconstruction.

[0056] Advantages:

[0057] 1. The high-precision interferometric measurement constellation error evolution method based on high-order characteristics disclosed in the present application sets the orbit state of three spacecraft in the space high-precision interferometric measurement constellation in the inertial system, sets the orbit dynamics model and initial state of each spacecraft in the high-precision interferometric measurement constellation; according to the initial state of the spacecraft, the characteristic direction of the initial orbit insertion error of the spacecraft is selected, the high-order model of the characteristic direction is established, and the high-order inverse model of the characteristic direction is established; according to the high-order inverse model of the characteristic direction, the initial value of the orbit error evolution semi-analytical expression based on the high-order characteristics is set; the orbit error evolution semi-analytical expression based on the high-order characteristics is calculated; the high-order characteristic direction is used to capture the high-order information of the initial error, which effectively improves the calculation accuracy.

[0058] 2. The high-precision interferometric measurement constellation error evolution method based on high-order characteristics disclosed in the present application only needs numerical integration when calculating the orbit error evolution semi-analytical expression based on high-order characteristics, so only one numerical integration is needed, and it has the advantage of high calculation efficiency.

[0059] 3. The high-precision interferometric measurement constellation error evolution method based on high-order characteristics disclosed in the present application uses the modulus of the initial position and velocity vector as the characteristic direction to establish the high-precision interferometric measurement constellation error evolution model based on the characteristic direction; the characteristic direction selection method is simple, and a large amount of calculation is not needed to obtain the characteristic direction, which further improves the prediction efficiency of the high-precision interferometric measurement constellation error evolution. BRIEF DESCRIPTION OF DRAWINGS

[0060] Figure 1 It is a flowchart of the high-precision interferometric measurement constellation error evolution method based on high-order characteristics in the present application. DETAILED DESCRIPTION

[0061] The embodiments of the present application will be described in detail below with reference to the accompanying drawings, but those skilled in the art will understand that the following examples are only for illustration of the present application and should not be regarded as limiting the scope of the present application.

[0062] Referring to Figure 1 The embodiment discloses a high-precision interferometric constellation error evolution method based on high-order characteristics, and the specific implementation steps are as follows:

[0063] Step 1: Set the orbit states of three spacecrafts in the high-precision interferometric constellation in the inertial system as x1(t)=[r1(t);v1(t)]、x2(t)=[r2(t);v2(t)] and x3(t)=[r3(t);v3(t)], and set the orbit dynamics model f[x(t),t] of the spacecrafts in the high-precision interferometric constellation and the initial state x0.

[0064] Set the orbit states of three spacecrafts in the high-precision interferometric constellation in the inertial system as x1(t)=[r1(t);v1(t)]、x2(t)=[r2(t);v2(t)] and x3(t)=[r3(t);v3(t)], wherein the subscripts 1, 2 and 3 represent spacecraft 1, spacecraft 2 and spacecraft 3. The initial states of the spacecrafts at the initial time t0 are represented as x 1,0 =x1(t0)、x 2,0 =x2(t0) and x 3,0 =x3(t0), and the initial states of the spacecrafts at the initial time t0 are shown in Table 1.

[0065] Table 1 Initial states of spacecrafts at initial time (the initial time is January 1, 2015)

[0066]

[0067] Set the orbit dynamics model f[x(t),t] of the spacecrafts in the high-precision interferometric constellation and the initial state x0=[r0;v0] as follows:

[0068]

[0069] Wherein f(x,t) is the spacecraft dynamics model, x0=[r0;v0] is the initial state, r0 is the initial position vector, and v0 is the initial velocity vector. In formula (1), in order to facilitate representation, the subscripts 1, 2 and 3 used to represent spacecraft 1, spacecraft 2 and spacecraft 3 are removed. The spacecraft dynamics model mainly includes the sun center gravity and the earth third body perturbation force, and the specific dynamics model expression is:

[0070]

[0071] Wherein μ s And μ e Respectively represent the gravitational constant of the sun and the earth, r eThe position of the Earth in the heliocentric inertial system. The ephemeris DE432 is used.

[0072] Step 2: According to the initial state of the spacecraft x0=[r0;v0], select the characteristic direction of the initial orbit error of the spacecraft, and establish a high-order model of the characteristic direction With According to the established high-order model of the characteristic direction, a high-order inverse model of the characteristic direction is established With

[0073] Step 3: According to the high-order inverse model of the characteristic direction established in step 2 With Set the initial value of the high-order characteristic-based orbit error evolution semi-analytical expression And

[0074] Step 4: According to the high-precision interferometric constellation orbit dynamics model f[x(t),t] of the spacecraft set in step 1 and the initial state x0, the high-order characteristic-based orbit error evolution semi-analytical expression initial value set in step 3 And Calculate the high-order characteristic-based orbit error evolution semi-analytical expression And

[0075] Step 5: Based on the high-order characteristic-based orbit error evolution semi-analytical expression calculated in step 4 And Construct the orbit error evolution semi-analytical expression of three spacecrafts in the space high-precision interferometric constellation.

[0076] Step 6: According to the orbit state of the three spacecrafts in the space high-precision interferometric constellation in the inertial system set in step 1 x1(t)=[r1(t);v1(t)], x2(t)=[r2(t);v2(t)] and x3(t)=[r3(t);v3(t)], the configuration stability index model z(t)=h[X(t)] of the space high-precision interferometric constellation is constructed.

[0077] Step 7: According to the configuration stability index model z(t)=h[X(t)] of the space high-precision interferometric constellation constructed in step 6, the high-precision interferometric constellation error evolution model based on the characteristic direction is established.

[0078] Step 8: based on the high-precision interferometric constellation error evolution model based on the characteristic direction established in step 7, the initial deviation is given, the configuration stability index deviation at a given time is calculated, and further based on the calculated configuration stability index deviation at a given time, the configuration stability is analyzed, and support is provided for mission orbit optimization and orbit reconstruction.

[0079] The initial error is set as: position single axis 100km, velocity single axis 1m / s, the average relative error of the predicted stability index after ten years is 0.0851%, and the calculation time is only 21.5070 seconds, which shows that the high-precision interferometric constellation error evolution method based on high-order characteristics has advantages in precision and efficiency.

[0080] The above specific description further details the purpose, technical scheme and beneficial effects of the application, and it should be understood that the above description is only a specific embodiment of the application, used to explain the application, and does not limit the protection scope of the application, and any modification, equivalent replacement, improvement, etc. within the spirit and principles of the application should be included in the protection scope of the application.

Claims

1. A high-precision interferometric constellation error evolution method based on high-order characteristics, characterized by: Comprising the following steps, Step 1: setting the orbit states of three spacecrafts in the high-precision space interferometric constellation in the inertial system as x1(t)=[r1(t);v1(t)], x2(t)=[r2(t);v2(t)] and x3(t)=[r3(t);v3(t)], setting the orbit dynamics model f[x(t),t] of each spacecraft in the high-precision space interferometric constellation and the initial state x0; Step 2: According to the initial state of the spacecraft x0=[r0;v0], select the characteristic direction of the initial orbit error of the spacecraft, and establish a high-order model of the characteristic direction and calculate its coefficients and According to the established high-order model of the characteristic direction, a high-order inverse model of the characteristic direction is established and calculate the corresponding coefficients and Step 3: High order inverse model of the characteristic direction established according to step 2 and the corresponding coefficients and Setting initial value of the semi-analytical expression of the orbit error evolution based on high order characteristics and Step 4: Compute the orbit dynamics model f[x(t),t] of the space vehicle in the high-precision interferometric constellation set in Step 1 and the initial state x0, the orbit error evolution semi-analytical expression initial value based on high-order characteristics set in Step 3 and Compute the orbit error evolution semi-analytical expression based on high-order characteristics and Step 5: Constructing the semi-analytical expression of the orbit error evolution based on the high-order features calculated in Step 4 and Constructing the semi-analytical expression of the orbit error evolution of three space vehicles in a high-precision space interferometric constellation Step 6: constructing the configuration stability index model z(t)=h[X(t)] of the high-precision space interferometric constellation according to the orbit states x1(t)=[r1(t);v1(t)], x2(t)=[r2(t);v2(t)] and x3(t)=[r3(t);v3(t)] of the three spacecrafts in the high-precision space interferometric constellation set in step 1; Step 7: According to the configuration stability index model z(t) = h[X(t)] of the spatial high-precision interferometric constellation constructed in Step 6, and the error evolution model of the high-precision interferometric constellation based on the characteristic direction established in Step 5 and establishing an error evolution model of the high-precision interferometric constellation based on the characteristic direction; Step 8: based on the high-precision space interferometric constellation error evolution model based on the characteristic direction established in step 7, giving the initial deviation, calculating the configuration stability index deviation at a given time, further based on the configuration stability index deviation at a given time calculated, i.e. the deviation of the predicted configuration stability, analyzing the configuration stability to obtain the configuration stability analysis result.

2. The method of claim 1, wherein: the high-precision interferometric constellation error evolution is based on high-order features. The implementation method of step 1 is, The orbit states of three spacecrafts in the constellation of high-precision space interferometric measurement are set as x1(t) = [r1(t); v1(t)], x2(t) = [r2(t); v2(t)] and x3(t) = [r3(t); v3(t)] in the inertial system, wherein the subscripts 1, 2 and 3 represent spacecraft 1, spacecraft 2 and spacecraft 3; wherein the initial states of the spacecrafts at the initial time t0 are represented as x 1,0 = x1(t0), x 2,0 = x2(t0) and x 3,0 = x3(t0); The orbit dynamics model f[x(t),t] of each spacecraft in the high-precision space interferometric constellation and the initial state x0=[r0;v0] are as follows: Wherein f(x,t) is the spacecraft dynamics model, x0=[r0;v0] is the initial state, r0 is the initial position vector, and v0 is the initial velocity vector; in formula (1), the subscripts 1, 2 and 3 of spacecraft 1, spacecraft 2 and spacecraft 3 are removed.

3. The method of claim 2, wherein: the high-order features are determined based on a plurality of high-precision interferometric measurements of the target object; and the plurality of high-precision interferometric measurements are obtained by a plurality of interferometric measurement devices. The implementation method of step 2 is, According to the initial state x0=[r0;v0] of the spacecraft, the characteristic direction of the initial orbit error of the spacecraft is selected as the modulus of the position vector and the velocity vector: y0 = [||r0||, ||v0||] T (2) Wherein y0 represents the characteristic direction; further, the high-order model of the characteristic direction is established as follows: wherein denotes the i-th element of the feature direction y0, denotes the k1-th element of the initial state x0, denotes the k2-th element of the initial state x0, and the symbol δ denotes a deviation, and is a tensor defined as: The high-order inverse model of the characteristic direction is established by using the inverse tensor as follows: wherein is the inverse tensor of is the inverse tensor of is the inverse tensor of​ 4. The method of claim 3, wherein: The implementation method of step 3 is, High order inverse model of the characteristic direction established according to step 2 and the corresponding coefficients and Setting initial values for the semi-analytical expression of the orbit error evolution based on high order characteristics and 5. The method of claim 4, wherein: the high-order features are determined based on a plurality of high-precision interferometric measurements of the target object; and the plurality of high-precision interferometric measurements are obtained by a plurality of interferometric measurement devices. The implementation method of step 4 is, the orbit dynamics model f [x(t), t] of the space vehicle in the high-precision interferometric constellation set according to step 1 and the initial state x0, the initial value of the orbit error evolution semi-analytical expression based on high-order characteristics set according to step 3 and the orbit error evolution semi-analytical expression based on high-order characteristics is calculated in a numerical integration manner and wherein the derivatives of the expressions and are Wherein: where f i [x(t),t] denotes the i-th element of the orbit dynamics model f[x(t),t].

6. The method of claim 5, wherein: The implementation method of step 5 is, Based on the high-order characteristic-based orbit error evolution semi-analytical expression calculated in step 4 and The orbit error evolution semi-analytical expression of three space vehicles in a high-precision space interferometric measurement constellation is constructed as follows: wherein y 1,0 denotes a characteristic direction vector of the first spacecraft, y 2,0 denotes a characteristic direction vector of the second spacecraft, y 3,0 denotes a characteristic direction vector of the third spacecraft, and are coefficients, defined as:

7. The method of claim 6, wherein: In step 6, The configuration stability index under consideration includes the arm length l ij (t), the breathing angle θ i (t), and the arm length change rate The corresponding functional expression is as follows: Wherein r ij (t) = r i (t) - r j (t) (18) v ij (t) = v i (t) - v j (t) (19) Let the description vector representing the configuration stability index, let the state vector of the three spacecraft, then there is a nonlinear mapping relationship as follows: Wherein z(t)=h[X(t)] represents the model of the configuration stability index.

8. The method of claim 7, wherein: the high-order features are based on a plurality of high-precision interferometric constellation error evolution measurements; and the high-precision interferometric constellation error evolution measurements are based on a plurality of high-precision interferometric measurements of a plurality of satellites. The implementation method of step 7 is, According to the configuration stability index model z(t)=h[X(t)] of the high-precision space interferometric constellation constructed in step 6, the high-precision space interferometric constellation error evolution model based on the characteristic direction is established as follows: Wherein: The implementation method of step 8 is, According to the configuration stability index model z(t)=h[X(t)] of the high-precision space interferometric constellation constructed in step 6, the high-precision space interferometric constellation error evolution model based on the characteristic direction is established as follows: Wherein:

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