Method for determining reference of control coordinate system of active rendezvous and docking spacecraft

Abstract

The invention relates to the technical field of aerospace, in particular to an active rendezvous and docking spacecraft control coordinate system reference determination method. The method comprises the following steps: for each star sensor, determining a first angle deviation between a first mounting matrix obtained by actual measurement and a theoretical mounting matrix in a state that a columnar spacecraft cabin is not provided with any equipment and the axis is in a vertical direction; when the star sensor is mounted on the spacecraft and the axis of the cabin of the spacecraft is in the horizontal direction, determining a second angle deviation between a second mounting matrix obtained by actual measurement and a theoretical mounting matrix; and taking the star sensor with the minimum difference between the first angle deviation and the second angle deviation as a reference of a control coordinate system. The active rendezvous and docking spacecraft control coordinate system benchmark determination method provided by the embodiment of the invention can determine a suitable star sensor as a reference benchmark, so that the deviation between a control coordinate system and a mechanical coordinate system is minimum.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and particularly to a method for determining the reference of the control coordinate system of an active rendezvous and docking spacecraft. Background Art

[0002] In related technologies, the selection of the control coordinate system generally uses a certain absolute attitude measurement sensor (generally a star sensor) of the GNC subsystem as a reference benchmark. This is reasonable for general small spacecraft. However, due to the different deformation amounts of different absolute measurement sensors relative to the mechanical coordinate system in orbit, for large active rendezvous and docking spacecraft, the selection of this benchmark must be carefully handled, and at the same time, the deviation estimation between the benchmark and the mechanical coordinate system must be reliable to evaluate whether the uncertainty between the control coordinate system and the mechanical coordinate system meets the design requirements. Currently, the precise measurement of large spacecraft still uses traditional methods and defaults that there is no deviation between the control coordinate system and the mechanical coordinate system.

[0003] Large active rendezvous and docking spacecraft have large structural dimensions. During their ground precise measurement, they are all placed horizontally with lateral supports and no pressurization treatment is carried out. The influence of gravity and atmospheric pressure will inevitably cause the deformation of the cabin body, resulting in the fact that the ground precise measurement results cannot directly reflect the deviation between the control coordinate system and the mechanical coordinate system after the spacecraft is in orbit. Therefore, this deviation cannot be eliminated by traditional precise measurement means. At the same time, the large structure will cause more obvious vacuum deformation, thermal deformation and weightlessness deformation after the spacecraft is in orbit, further increasing the uncertainty between the control coordinate system and the mechanical coordinate system. Summary of the Invention

[0004] An embodiment of the present invention provides a method for determining the reference of the control coordinate system of an active rendezvous and docking spacecraft, which can determine a suitable star sensor as a reference benchmark to minimize the deviation between the control coordinate system and the mechanical coordinate system.

[0005] An embodiment of the present invention provides a method for determining the reference of the control coordinate system of an active rendezvous and docking spacecraft, including:

[0006] For each star sensor, the following operations are performed:

[0007] In the state where no equipment is installed on the columnar spacecraft cabin body and the axis is in the vertical direction, determine the first angular deviation between the actually measured first installation matrix and the theoretical installation matrix;

[0008] When the star sensor is installed on the spacecraft and the axis of the spacecraft cabin body is in the horizontal direction, determine the second angular deviation between the actually measured second installation matrix and the theoretical installation matrix;

[0009] Take the star sensor with the smallest difference between the first angular deviation and the second angular deviation as the control coordinate system benchmark.

[0010] In a possible design, the first installation matrix is determined as follows:

[0011] In a state where the axis of the spacecraft cabin is vertical and no equipment is installed, measure the machining accuracy of the mounting base of the star sensor;

[0012] Establish the first installation matrix according to the machining accuracy of the mounting base of the star sensor.

[0013] In a possible design, determining the first angular deviation between the actually measured first installation matrix and the theoretical installation matrix includes:

[0014] Left-multiply the first installation matrix by the theoretical installation matrix and perform normalization to obtain a first comparison matrix;

[0015] The first comparison matrix determines the first angular deviation through a deviation formula.

[0016] In a possible design, the second installation matrix is determined as follows:

[0017] When the star sensor is installed on the spacecraft and the axis of the spacecraft cabin is horizontal, perform precise measurement of the star sensor every 90° rotation to obtain four groups of second installation matrices.

[0018] In a possible design, determining the second angular deviation between the actually measured second installation matrix and the theoretical installation matrix includes:

[0019] After left-multiplying each of the four groups of second installation matrices by the theoretical installation matrix, accumulate and normalize to obtain a second comparison matrix;

[0020] The second comparison matrix determines the second angular deviation through a deviation formula.

[0021] In a possible design, the deviation formula is:

[0022] δ = arccos[tr(K) / 3 - 1]

[0023] where δ is the angular deviation and K is the comparison matrix for determining the angular deviation to be determined.

[0024] In a possible design, the initial rotation position of the star sensor is such that the angle between the line connecting it to the axis of the spacecraft cabin and the horizontal plane is 45°.

[0025] In a possible design, after using the star sensor with the smallest difference between the first angular deviation and the second angular deviation as the control coordinate system reference, it further includes:

[0026] The first angular deviation of the star sensor serving as the reference is used as the error estimation matrix between the control coordinate system reference and the mechanical coordinate system reference.

[0027] Compared with the prior art, the present invention has at least the following beneficial effects:

[0028] In this embodiment, after the spacecraft cabin is processed and in the vertical state without any equipment installed, the deformation of the cabin is the smallest, and this state is closest to the state after the spacecraft is in orbit. Therefore, by comparing the first installation matrix established from the measurement data in this state with the theoretical installation matrix, the first angular deviation can be obtained, and this first angular deviation can be basically consistent with the actual situation in the in-orbit state. When the spacecraft is horizontal, the second installation matrix is measured, and the second installation matrix can reflect the actual installation situation on the ground. By comparing it with the theoretical installation matrix, the second angular deviation is obtained. By comparing the first angular deviation and the second angular deviation, when the deviation between the two is the smallest, it proves that the ground state and the in-orbit state are the closest. Therefore, the star sensor corresponding to the smallest deviation is selected as the reference of the control coordinate system. Description of the Drawings

[0029] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for the description of the embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0030] Figure 1 It is a schematic diagram of different positions for collecting data in the horizontal state of the spacecraft cabin provided by the embodiment of the present invention. Detailed Embodiments

[0031] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0032] In the description of the embodiments of the present invention, unless otherwise clearly specified and defined, the terms "first" and "second" are only used for descriptive purposes and cannot be construed as indicating or implying relative importance; unless otherwise specified or stated, the term "plurality" means two or more; the terms "connection", "fixation", etc. should be understood in a broad sense. For example, "connection" can be a fixed connection, a detachable connection, an integral connection, or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0033] In the description of this specification, it should be understood that the orientation terms such as "upper" and "lower" described in the embodiments of the present invention are described from the angles shown in the drawings and should not be construed as limiting the embodiments of the present invention. In addition, in the context, it should also be understood that when it is mentioned that an element is connected "above" or "below" another element, it can not only be directly connected "above" or "below" another element, but also be indirectly connected "above" or "below" another element through an intermediate element.

[0034] The embodiments of the present invention provide a method for determining the reference of the control coordinate system of an active rendezvous and docking spacecraft, including:

[0035] For each star sensor, the following operations are performed:

[0036] In the state where no equipment is installed on the columnar spacecraft cabin and the axis is in the vertical direction, determine the first angular deviation between the actually measured first installation matrix and the theoretical installation matrix;

[0037] When the star sensor is installed on the spacecraft and the axis of the spacecraft cabin is in the horizontal direction, determine the second angular deviation between the actually measured second installation matrix and the theoretical installation matrix;

[0038] Use the star sensor with the smallest difference between the first angular deviation and the second angular deviation as the reference of the control coordinate system.

[0039] In this embodiment, after the spacecraft cabin is processed, in the vertical state without installing any equipment, the deformation of the cabin is the smallest, and this state is closest to the state of the spacecraft after it enters orbit. Therefore, by comparing the first installation matrix established from the measurement data in this state with the theoretical installation matrix, the first angular deviation can be obtained, and this first angular deviation can basically be consistent with the actual situation of the in-orbit state. When the spacecraft is horizontal, the second installation matrix is measured. The second installation matrix can reflect the actual installation situation on the ground. By comparing it with the theoretical installation matrix, the second angular deviation is obtained. By comparing the first angular deviation and the second angular deviation, when the deviation between the two is the smallest, it proves that the ground state and the on-orbit state are the closest. Therefore, the star sensor corresponding to the smallest deviation is selected as the benchmark for the control coordinate system.

[0040] In some embodiments of the present invention, the first installation matrix is determined in the following manner:

[0041] Measure the processing accuracy of the star sensor mounting base;

[0042] Establish the first installation matrix according to the processing accuracy of the star sensor mounting base.

[0043] In this embodiment, after the spacecraft cabin is processed, in the vertical state without installing any equipment, measure the processing accuracy of the star sensor mounting base, and calculate the first installation matrix H of each star sensor based on this sts_i_0 .

[0044] In some embodiments of the present invention, determining the first angular deviation between the actually measured first installation matrix and the theoretical installation matrix includes:

[0045] Left-multiply the first installation matrix by the theoretical installation matrix and perform normalization processing to obtain the first comparison matrix;

[0046] The first comparison matrix determines the first angular deviation through the deviation formula.

[0047] The left multiplication of this first installation matrix by the inverse of the theoretical installation matrix of the star sensor (i.e., the transpose of the transformation matrix T of the mechanical coordinate system sts_i ) and perform normalization processing to obtain the first comparison matrix K sts_i_0 (i.e., ), and obtain the first angular deviation δ according to the deviation formula i_0 .

[0048] That is, δ i_0 = arccos[tr(K sts_i_0 ) / 3 - 1], where i is the number of the star sensor.

[0049] For example, Figure 1 , in some embodiments of the present invention, the second installation matrix is determined in the following manner:

[0050] When the star sensor is installed on the spacecraft and the axis of the spacecraft cabin is horizontal, the star sensor performs precise measurements every 90° rotation, and four groups of second installation matrices are obtained.

[0051] In this embodiment, the large dispersion of the precise measurement results of the star sensor caused by the flexible deformation of the spacecraft in different states is overcome. Therefore, four groups of data in different attitudes are selected for comparison.

[0052] In some embodiments of the present invention, determining the second angular deviation between the actually measured second installation matrix and the theoretical installation matrix includes:

[0053] After left-multiplying the four groups of second installation matrices by the theoretical installation matrix, they are accumulated and normalized to obtain a second comparison matrix;

[0054] The second comparison matrix determines the second angular deviation through the deviation formula.

[0055] Specifically, for the precise measurement of each star sensor in four states, 4 groups of precise measurement results H sts_i_1 、H sts_i_2 、H sts_i_3 、H sts_i_4 are obtained for each star sensor, as follows:

[0056] Table 1 4 groups of attitude measurement data of star sensor i

[0057]

[0058]

[0059] Note: In the table, X sts_i , Y sts_i , Z sts_i are the X, Y, and Z axes of the star sensor measurement coordinate system respectively; X J , Y J , Z J are the X, Y, and Z axes of the mechanical coordinate.

[0060] According to the theoretical installation position of star sensor i, the transformation matrix T sts_i (i.e., the transformation matrix to the mechanical coordinate system) from the star sensor i to the theoretical installation matrix is obtained, and the above 4 precise measurement data H sts_i_1 、H sts_i_2 、H sts_i_3 、H sts_i_4 are transformed into the mechanical coordinate system to obtain

[0061] After accumulating and normalizing the matrix , a new matrix, that is, the second comparison matrix K sts_iThe angular deviation estimation (second angular deviation) between the precise measurement data of the star sensor and the actual data after orbit injection is calculated by the following formula:

[0062] δ i = arccos[tr(K sts_i ) / 3 - 1]

[0063] where δ i is the second angular deviation.

[0064] In some embodiments of the present invention, the deviation formula is:

[0065] δ = arccos[tr(K) / 3 - 1]

[0066] where δ is the angular deviation and K is the comparison matrix for determining the angular deviation.

[0067] In some embodiments of the present invention, the initial rotation position of the star sensor is such that the angle between the line connecting it to the axis of the spacecraft cabin and the horizontal plane is 45° (with an error controlled within ±5°).

[0068] In some embodiments of the present invention, after using the star sensor with the smallest difference between the first angular deviation and the second angular deviation as the control coordinate system reference, it further includes:

[0069] Using the first angular deviation of the star sensor used as the reference as the error estimation matrix between the control coordinate system reference and the mechanical coordinate system reference.

[0070] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that they can still modify the technical solutions described in the foregoing embodiments, or perform equivalent replacements for some of the technical features; and these modifications or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for determining the reference of the control coordinate system of an active rendezvous and docking spacecraft, characterized in that Including: For each star sensor, the following operations are performed: In a state where no equipment is installed on the cylindrical spacecraft cabin and its axis is in the vertical direction, determine the first angular deviation between the actually measured first installation matrix and the theoretical installation matrix; When the star sensor is installed on the spacecraft and the axis of the spacecraft cabin is in the horizontal direction, determine the second angular deviation between the actually measured second installation matrix and the theoretical installation matrix; Take the star sensor with the smallest difference between the first angular deviation and the second angular deviation as the control coordinate system reference.

2. The method according to claim 1, characterized in that The first installation matrix is determined in the following manner: In a state where the axis of the spacecraft cabin is vertical and no equipment is installed, measure the machining accuracy of the star sensor mounting base; Establish the first installation matrix based on the machining accuracy of the star sensor mounting base.

3. The method according to claim 2, wherein The determination of the first angular deviation between the actually measured first installation matrix and the theoretical installation matrix includes: Left-multiply the first installation matrix by the theoretical installation matrix and perform normalization to obtain the first comparison matrix; The first comparison matrix determines the first angular deviation through the deviation formula.

4. The method according to claim 1, characterized in that The second installation matrix is determined in the following manner: When the star sensor is installed on the spacecraft and the axis of the spacecraft cabin is in the horizontal direction, perform precise measurements of the star sensor every 90° rotation to obtain four groups of second installation matrices.

5. The method according to claim 4, characterized in that, The determination of the second angular deviation between the actually measured second installation matrix and the theoretical installation matrix includes: After left-multiplying each of the four groups of second installation matrices by the theoretical installation matrix, accumulate and normalize to obtain the second comparison matrix; The second comparison matrix determines the second angular deviation through the deviation formula.

6. The method according to claim 3 or 5, characterized in that, The deviation formula is: δ = arccos[tr(K) / 3 - 1] where δ is the angular deviation and K is the comparison matrix for which the angular deviation is to be determined.

7. The method according to claim 4, characterized in that The initial rotation position of the star sensor is such that the angle between the line connecting it to the axis of the spacecraft cabin and the horizontal plane is 45°.

8. The method according to claim 1, characterized in that, After taking the star sensor with the smallest difference between the first angular deviation and the second angular deviation as the control coordinate system reference, it further includes: Take the first angular deviation of the star sensor used as the reference as the error estimation matrix between the control coordinate system reference and the mechanical coordinate system reference.

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