A method for inter-satellite link based satellite attitude cooperative calibration of internet constellation

Abstract

The application provides a satellite attitude cooperative calibration method based on inter-satellite link for internet constellation, comprising: acquiring node satellite and adjacent satellite data; calculating link inertia vector in ideal state through orbit vector; calculating attitude estimation value through the relationship between link inertia vector in ideal state and actual link vector; when satellite attitude sensor works normally, correcting estimated attitude based on error attitude correction amount; completing link alignment or alignment optimization according to satellite attitude of attitude estimation or correction completion. The application utilizes the relationship between theoretical inertia vector and actual load vector of inter-satellite link, calculates attitude estimation matrix through pseudo-inverse solution and orthogonal optimization, utilizes multi-vector joint attitude calibration of inter-satellite link of adjacent satellite, corrects attitude error through link vector when attitude sensor is effective, improves attitude calibration precision, and provides reliable attitude estimation and maintenance function through link assistance when attitude sensor is invalid.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of satellite attitude estimation, and in particular to a satellite attitude cooperative calibration method based on inter-satellite links for an Internet constellation. BACKGROUND

[0002] In recent years, low-orbit Internet satellite constellations have gradually become a hot topic in the global communication field. Low-orbit satellite constellations provide low-latency, high-bandwidth communication services for global users by deploying a large number of low-orbit satellites. In order to achieve stable communication between satellites, especially laser links between low-orbit satellites (inter-satellite links), are widely used. Accurate control of the satellite attitude is crucial for the alignment and communication quality of the inter-satellite link, because the high-speed motion and orbital changes of the satellite will cause attitude errors, which in turn affect the stability and communication efficiency of the link. Therefore, the satellite constellation needs to be regularly calibrated and corrected with high precision to ensure the alignment of the laser link.

[0003] Existing satellite attitude correction methods mainly rely on satellite attitude sensors (such as star sensors, gyroscopes, etc.) and ground control instructions. However, these methods have certain limitations when faced with a large number of low-orbit satellites and complex operating environments. The existing satellite attitude calibration technology mainly has the following shortcomings:

[0004] 1) Limitations of on-board sensors

[0005] Single-satellite attitude determination methods usually rely on on-board sensors (such as star sensors, gyroscopes, etc.), but these sensors have measurement errors and drift problems. When the star sensor fails, it is difficult to maintain high precision for single-satellite attitude, which will directly affect the alignment accuracy of the laser link.

[0006] 2) High dependence on ground control

[0007] Traditional satellite attitude calibration methods usually rely on ground station control instructions to correct the satellite attitude. However, due to the large number of low-orbit constellation satellites and short orbital periods, it is difficult for ground stations to efficiently cover and perform real-time attitude calibration, and there are problems of delay and poor timeliness. This method is particularly unsuitable for low-orbit constellations and cannot meet the real-time and efficient attitude correction requirements.

[0008] 3) Lack of multi-satellite cooperative calibration mechanism

[0009] Neither the data based on single-satellite attitude sensors nor the ground-led calibration methods make full use of the attitude data of other satellites within the constellation for cooperative calibration. In the existing technology, inter-satellite links are mainly used for data transmission, and the attitude information of multiple satellites within the constellation is not effectively utilized for cooperative work, resulting in a lack of efficient attitude error correction mechanism. SUMMARY

[0010] In view of the deficiencies of the prior art, the present application provides a satellite attitude cooperative calibration method based on inter-satellite links for an Internet constellation, which calibrates the attitude by using the multi-vector inter-satellite links of adjacent satellites, and improves the attitude calibration accuracy by correcting the attitude error through the link vector when the attitude sensor is effective; and provides reliable attitude estimation and maintenance functions through link assistance when the attitude sensor is ineffective.

[0011] The technical scheme of the present application is as follows: a satellite attitude cooperative calibration method based on inter-satellite links for an Internet constellation, comprising the following steps:

[0012] S1), acquiring node satellite and adjacent satellite data;

[0013] S2), calculating the link inertial vector in an ideal state through an orbit vector;

[0014] S3), calculating an attitude estimation value through the relationship between the link inertial vector in an ideal state and an actual link vector;

[0015] S4), when the satellite attitude sensor is working normally, correcting the estimated attitude based on an error attitude correction amount;

[0016] S5), completing the alignment or alignment optimization of the link according to the satellite attitude of the attitude estimation or correction.

[0017] Preferably, in step S1), the links are built in a cross configuration, and the acquired node satellite and adjacent satellite data include satellite orbit position, link load optical axis vector, and satellite own attitude data.

[0018] Preferably, in step S2), the calculation formula of the link inertial vector in an ideal state is as follows:

[0019]

[0020] In the formula, l i-1，i eci is the link inertial vector of S i-1，j pointing to S i，j , l i+1，i eci is the link inertial vector of S i+1，j pointing to S i，j , l j-1，j eci is the link inertial vector of S i，j-1 pointing to S i，j , l j+1，j eci is the link inertial vector of S i，j+1 pointing to S i，j ; and l i，j, i is the index of the orbital plane, j is the index of the satellite within the orbital plane; S i-1，j , S i+1，j , S i，j-1 , S i，j+1 is the adjacent satellite of the satellite S i，j ; r i，j is the orbital position of the satellite S i，j ; r i-1，j , r i+1，j , r i，j-1 , r i，j+1 are the orbital positions of the satellites S i-1，j , S i+1，j , S i，j-1 , S i，j+1 , respectively; the superscript b represents the body coordinate system; the superscript eci represents the inertial coordinate system.

[0021] As preferred, in step S3), the attitude estimation value is calculated by the relationship between the link inertial vector in ideal state and the actual link vector, and specifically includes the following steps:

[0022] S31), the relationship between the link inertial vector in ideal state and the actual link vector is represented as:

[0023] L b = A kj L eci (1) ;

[0024] In the formula, L b is the actual link vector matrix in the body coordinate system; L eci is the link inertial vector matrix in ideal state; A kj is the attitude estimation matrix; the subscript kj represents the index of the attitude estimation matrix;

[0025] S32), the pseudo-inverse solution formula (1) is constructed by minimizing the sum of squares of errors to obtain the inverse solution matrix A * :

[0026] A * = L b (L eci ) T (L eci (L eci ) T ) -1 (2) ;

[0027] In the formula, A * represents the inverse solution matrix; (L eci ) T represents the transpose matrix of the link inertial vector matrix L eci in ideal state; T represents the transpose operation;

[0028] S33), the inverse solution matrix A * Orthogonalization is performed to meet the orthogonal constraint condition of the rotation matrix, and the final attitude estimation matrix A is obtained by optimization kj ;

[0029] S34), the final attitude estimation matrix A kj is converted into an estimated attitude quaternion

[0030] As preferred, in step S4), the estimated attitude is corrected based on the error attitude correction quantity, specifically including the following steps:

[0031] S41), the laser load error and the attitude determination error are combined into a correction quantity Q e ;

[0032] S42), the estimated attitude quaternion e is corrected based on the error attitude correction quantity Q to obtain a corrected satellite attitude quaternion

[0033] As preferred, in step S5), when the attitude sensor fails, the estimated attitude quaternion is used to maintain the alignment of the laser link.

[0034] As preferred, in step S5), when the attitude sensor is valid, the corrected satellite attitude quaternion is input into the control system to further adjust the attitude and realize link alignment optimization.

[0035] The beneficial effects of the present application are:

[0036] 1. The present application utilizes the relationship between the theoretical inertial vector of the inter-satellite link and the actual load vector, constructs a high-precision attitude estimation matrix through pseudo-inverse solution and orthogonal optimization, and further converts it into a quaternion for attitude correction;

[0037] 2. The present application fully utilizes the inter-satellite link vector information to provide a reliable auxiliary mechanism for attitude estimation, which can still ensure high attitude precision even in complex environments or under increased sensor error;

[0038] 3. The present application estimates the attitude of the target satellite through the orbit vector and link load information, combined with the attitude and orbit data of adjacent satellites, to ensure the alignment of the laser link and the stability of communication in the short term;

[0039] 4. The present application constructs a distributed attitude calibration architecture through the cooperative action of multiple satellites, corrects the sensor error through link attitude estimation when the star sensor is valid, and maintains the attitude reference of the target satellite through multi-satellite cooperation when the sensor fails.

[0040] 5、The application adopts the optimization mechanism of minimizing the square sum of link errors in the distributed architecture, and effectively avoids the situation that some links are better aligned while others are worse in the system calibration process through comprehensive balancing of link performance. The balancing optimization feature ensures the overall accuracy of the target satellite attitude and the consistency of link performance, significantly improving the overall stability of the link.

[0041] 6、The distributed architecture of the application is more robust in calibration accuracy and link alignment effect, and is particularly suitable for low-orbit constellation in high dynamic environment. BRIEF DESCRIPTION OF DRAWINGS

[0042] Figure 1 is a flowchart of the method of the application;

[0043] Figure 2 is a schematic diagram of the satellite cross-configuration inter-satellite link of the embodiment of the application. DETAILED DESCRIPTION

[0044] The specific embodiments of the application will be further described below in conjunction with the accompanying drawings:

[0045] As shown in the figure, the embodiment provides a satellite attitude cooperative calibration method based on inter-satellite link for internet constellation, which includes the following steps: Figure 1

[0046] S1), acquiring node satellite and adjacent satellite data;

[0047] The embodiment utilizes the multi-vector joint attitude determination of inter-satellite link of adjacent satellites, as shown in the figure, when the chain is built in the cross-configuration, the adjacent satellites of satellite S i，j are S i-1，j , S i+1，j , S i，j-1 , and S i，j+1 . Figure 2

[0048] The satellite data acquired by the embodiment includes: orbit position, link load optical axis vector, and satellite own attitude data.

[0049] The node satellite acquires the orbit vector r of the satellite in the inertial system according to its own orbit, and acquires the orbit vector of the adjacent satellite through the inter-satellite link, and acquires the load optical axis vector matrix in the satellite system, that is, the actual link vector matrix L b :

[0050] L b = [p i，i-1 b p j，j-1 b p i，i+1 ​​b p j，j+1 b ];

[0051] wherein,

[0052]

[0053]

[0054] wherein, Q = [q0, q1, q2, q3] T represents satellite attitude inertial quaternion; superscript T represents transpose operator; superscript b represents body coordinate system; superscript eci represents inertial coordinate system; q0, q1, q2, q3 represent each element of quaternion respectively; o represents rotation operation of quaternion and vector; subscript i is orbit plane number, subscript j is satellite number in orbit plane;

[0055] When the satellite attitude sensor (such as star sensor) works normally, the satellite own attitude data is acquired. S2), the link inertial vector in ideal state is calculated through orbit vector; the calculation formula is:

[0056]

[0057] wherein, l i-1，i eci is the link inertial vector of S i-1，j pointing to S i，j , l i+1，i eci is the link inertial vector of S i+1，j pointing to S i，j , l j-1，j eci is the link inertial vector of S i，j-1 pointing to S i，j , l j+1，j eci is the link inertial vector of S i，j+1 pointing to S i，j ; for the satellite S i，j on the constellation, i is orbit plane number, j is satellite number in orbit plane; S i-1，j , S i+1，j , S i，j-1 , S i，j+1 are adjacent satellites of satellite S i，j ; r i，j is the orbit position of satellite S i，j ; r i-1，j , r i+1，j , r i，j-1 , r i，j+1 are respectively the orbit positions of satellite S i-1，j , S i+1，j , S i，j-1, S i，j+1 The superscript b represents the body coordinate system; the superscript eci represents the inertial coordinate system.

[0058] S3), calculating the attitude estimation value through the relationship between the ideal state link inertial vector and the actual link vector; specifically comprising the following steps:

[0059] S31), the relationship between the ideal state link inertial vector and the actual link vector is represented as:

[0060] L b = A kj L eci (1);

[0061] In the formula, L b is the actual link vector matrix in the body coordinate system; L eci is the ideal state link inertial vector matrix; A kj is the attitude estimation matrix; the subscript kj represents the index of the attitude estimation matrix.

[0062] Wherein,

[0063] L b = [p i，i-1 b p j,j-1 b p i,i+1 b p j，j+1 b ];

[0064] L eci = [l i-1，i eci l j-1，j eci l i+1，i eci l j+1，j eci ];

[0065] S32), constructing the pseudo-inverse solution formula (1) by minimizing the error sum of squares to obtain the inverse solution matrix A * :

[0066] A * = L b (L eci ) T (L eci (L eci ) T ) -1 (2);

[0067] In the formula, A * represents the inverse solution matrix; (L eciT L is a link inertia vector matrix in an ideal state eci , and L

[0068] S33), orthogonalizing the inverse solution matrix A * to satisfy the orthogonal constraint condition of the rotation matrix, and optimizing to obtain the final attitude estimation matrix A kj

[0069]

[0070] where a 11 ~ a 33 respectively represent each element of the 3x3 matrix A kj ; I represents a 3x3 unit matrix

[0071] S34), converting the final attitude estimation matrix A kj into an estimated attitude quaternion

[0072]

[0073] The quaternion solution formula is as follows:

[0074]

[0075] Based on the first row of formula (7)-(10), q0, q1, q2, q3 are calculated respectively, the absolute values of q0, q1, q2, q3 are compared, and the formula corresponding to the maximum absolute value is selected to calculate the estimated attitude

[0076] S4), when the satellite attitude sensor works normally, the estimated attitude is corrected based on the error attitude correction quantity; specifically including the following steps:

[0077] S41), combining the laser load error and the attitude determination error into an error attitude correction quantity Q e ; the expression of the error attitude correction quantity Q e is:

[0078]

[0079] wherein, respectively represent each element of the quaternion ; Q s is the attitude quaternion determined by the satellite attitude sensor; is the estimated attitude;

[0080] S42), based on the error attitude correction quantity Q e , the estimated attitude quaternion ​​The modified satellite attitude quaternion is obtained by correction That is,

[0081]

[0082] Q e is an error attitude correction quantity; Q s is an attitude quaternion determined by a satellite attitude sensor.

[0083] S5), according to the satellite attitude after the attitude estimation or correction is completed, the alignment or alignment optimization of the link is completed.

[0084] In the embodiment, when the attitude sensor is invalid, the estimated attitude quaternion maintains the alignment of the laser link.

[0085] When the attitude sensor is valid, the modified satellite attitude quaternion is input into the control system to further adjust the attitude, so as to realize the alignment optimization of the link.

[0086] The above embodiment and the description in the specification only illustrate the principles and the best mode of the present application, and various changes and improvements can be made without departing from the spirit and the scope of the present application, and these changes and improvements all fall within the scope of the present application.

Claims

1. A method for inter-satellite link based satellite attitude co-calibration for internet constellation, characterized in that, The method comprises the following steps: S1), acquiring node satellite and adjacent satellite data; S2), calculating link inertia vector in ideal state through orbit vector; S3), calculating satellite attitude estimation value through the relationship between link inertia vector in ideal state and actual link vector, and converting the final attitude estimation matrix into estimated attitude quaternion; S4), when the satellite attitude sensor works normally, correcting the satellite estimated attitude based on error attitude correction amount; specifically comprising the following steps: S41), combine the laser load error and the pose error into an error pose correction amount ; The error attitude correction amount is expressed as: wherein each element of the quaternion respectively; is a determined attitude quaternion of a satellite attitude sensor; is an estimated attitude; S42)、based on the error attitude correction amount to the estimated attitude quaternion to obtain a corrected satellite attitude quaternion ​ Corrected satellite attitude quaternion is expressed as: ； wherein is an error attitude correction; is an attitude quaternion determined by a satellite attitude sensor; S5), according to the satellite attitude estimation or correction completed, complete link alignment or alignment optimization, when the attitude sensor is effective, the modified satellite attitude quaternion The input control system further adjusts the attitude to achieve link alignment optimization; Utilizing estimated attitude quaternion when attitude sensor fails Maintaining alignment of the laser link.

2. The method of claim 1, wherein the method is used for inter-satellite link based satellite cooperative attitude determination for an internet constellation. In step S1), the node satellite and adjacent satellite data are acquired according to the cross configuration chain, and the data includes satellite orbit position, link load optical axis vector and satellite own attitude data.

3. The method of claim 1, wherein the method is used for inter-satellite link based satellite cooperative attitude determination for internet constellation. In step S2), the calculation formula of the link inertia vector in ideal state is as follows: In the formula, for point to The link inertia vector, for point to The link inertia vector, for point to The link inertia vector, for point to Link inertial vector; for satellites on the constellation , Number the track surfaces. Number the satellites within the orbital plane; , , , For satellite Neighboring satellites; For satellite The orbital position; , , , Satellites , , , The orbital position; the superscript b indicates the body coordinate system; the superscript This represents an inertial coordinate system.

4. The method of claim 3, wherein the method is characterized by: In step S3), the attitude estimation value is calculated through the relationship between the link inertia vector in ideal state and the actual link vector, specifically comprising the following steps: S31), the relationship between the link inertia vector in ideal state and the actual link vector is expressed as: （1）； wherein is the actual link vector matrix under the system; is the ideal link inertia vector matrix; is the attitude estimation matrix, the subscript denotes the index of the attitude estimation matrix; S32)、by minimizing the error sum of squares to construct pseudo-inverse solution formula (1), get inverse solution matrix : （2）； wherein denotes the inverse solution matrix; denotes the transpose of the link inertial vector matrix under ideal conditions, T denotes the transpose operation; S33), to the inverse solution matrix orthogonalization processing is performed to satisfy the orthogonal constraint condition of the rotation matrix, and a final attitude estimation matrix is obtained by optimization ; S34), converting the final pose estimation matrix into an estimated pose quaternion .

5. The method of claim 4, wherein: In step S33), the final pose estimation matrix is expressed as: wherein respectively represent each element of a 3x3 matrix represents a 3x3 identity matrix.​ 6. The method of claim 5, wherein: In step S34), the estimated attitude quaternion is expressed as: ； The quaternion solution formula is as follows: （7）； （8）； （9）； （10）； Based on the absolute value of the first row of formula (7)-(10) respectively And compare The absolute value of the maximum corresponding to the formula is selected to calculate the estimated attitude .

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