Method for establishing high-stability inertial reference of satellite star sensor

Abstract

A method for establishing a high-stability inertial reference of a satellite star sensor relates to the technical field of spacecraft control, and comprises the following steps: integrally mounting the star sensor and a load accelerometer; stray light suppression is carried out on the star sensor in the cabin by adopting a separated two-stage light shield; sectional type thermal control coating is adopted outside the two-stage light shields, and stray light prevention treatment is carried out on the mounting surfaces of the light shields; calculating the inertial attitude of the load accelerometer by using the output of the star sensor, and establishing an inertial attitude model of the load accelerometer; calibrating a star sensor mounting matrix based on the load accelerometer, and establishing a unified inertial measurement reference between the star sensor and the load accelerometer; for a plurality of star sensors, a unified inertial measurement reference is established among the plurality of star sensors.

Description

Technical Field

[0001] The present invention relates to a method for establishing a highly stable inertial reference for a satellite star sensor, and relates to the field of spacecraft control technology. Background Art

[0002] Spacecraft usually use star sensors for attitude determination. Its accuracy not only depends on the measurement accuracy of the star sensor itself, but is also affected by factors such as external heat flux radiation, installation accuracy, and stray light. At the same time, the star sensor also serves as one of the satellite payloads to provide high-precision inertial attitude data for other satellite payloads; therefore, the inertial attitude data of the star sensor not only affects the attitude determination of the satellite, but is also one of the factors affecting satellite data products. Especially for the star sensor installed inside the spacecraft, compared with the star sensor installed outside the spacecraft, the factors affecting the attitude determination accuracy are more complex. In addition, there are always periodic temperature changes, relative structural deformations, etc. in the on-orbit spacecraft, resulting in the star sensor being unable to provide a continuous and highly stable inertial attitude reference for the spacecraft platform and its payloads. Especially for spacecraft equipped with inertial acceleration sensor payloads, the prior art lacks a means of unifying the inertial measurement reference and cannot provide a long-term highly stable inertial reference for the inertial acceleration sensor payload. The present invention proposes a method for establishing a highly stable inertial reference for a satellite star sensor, which can provide a long-term highly stable inertial measurement reference for the spacecraft. Summary of the Invention

[0003] The technical problem to be solved by the present invention is: to overcome the deficiencies of the prior art, provide a long-term highly stable inertial measurement reference for the spacecraft, and ensure the high-precision attitude determination and unified inertial measurement reference of the spacecraft with a star sensor installed inside the cabin.

[0004] The object of the present invention is achieved by the following technical solutions:

[0005] A method for establishing a highly stable inertial reference for a satellite star sensor includes:

[0006] Integrally install the star sensor and the payload accelerometer;

[0007] Use a split two-stage light shield to suppress stray light for the star sensor inside the cabin;

[0008] Adopt segmented thermal control coating outside the two-stage light shield, and perform anti-stray light treatment on the installation surface of the light shield;

[0009] Calculate the inertial attitude of the payload accelerometer using the output of the star sensor, and establish an inertial attitude model of the payload accelerometer; calibrate the star sensor installation matrix based on the payload accelerometer, and establish a unified inertial measurement reference between the star sensor and the payload accelerometer;

[0010] For multiple star sensors, a unified inertial measurement reference is established among the multiple star sensors.

[0011] In an embodiment of the present invention, the star sensor and the payload accelerometer are integrally installed on a base, and the installation part of the star sensor and the base is made of the same material; the base and the satellite structure are flexibly connected to achieve decoupling of thermal deformation.

[0012] In an embodiment of the present invention, the two-stage light shield includes a primary light shield and a secondary light shield. The primary light shield is fixedly connected to the star sensor body, and the secondary light shield is installed at the secondary installation surface outside the satellite cabin. The two-stage light shields do not contact each other.

[0013] In an embodiment of the present invention, a thermal control multi-layer with a thickness of N1 is wrapped by shrinking a certain distance from the end face of the light outlet of the primary light shield. The remaining part of the primary light shield is wrapped with a thermal control multi-layer with a thickness of N2 according to the thermal design result, and N2 > N1 + the wall thickness of the secondary light shield; the thermal control multi-layer on the outer surface of the secondary light shield is wrapped according to the thermal design result.

[0014] In an embodiment of the present invention, the installation surface of the secondary light shield is blackened.

[0015] In an embodiment of the present invention, the inertial attitude model of the payload accelerometer is:

[0016] q acc =C b ·C s ·q m

[0017] Where q acc is the measurement value of the payload accelerometer relative to the inertial system, C b is the relationship matrix between the optical reference axis of the star sensor and the coordinate system of the payload accelerometer, C s is the relationship matrix between the measurement axis of the star sensor and its own optical reference axis, q m is the measurement value of the star sensor measurement system relative to the inertial system; denote C bs =C b ·C s to represent the initial installation matrix of the star sensor relative to the payload accelerometer, and C s 、C b are obtained through ground precise measurement.

[0018] In an embodiment of the present invention, based on the calibration of the star sensor installation matrix by the payload accelerometer, establishing a unified inertial measurement reference between the star sensor and the payload accelerometer includes:

[0019] Using the spacecraft actuator, perform an attitude maneuver with a fixed angular acceleration in a certain order, and calculate the installation deviation between the star sensor and the payload accelerometer under a constant acceleration: the measurement quaternion of the star sensor measurement system relative to the inertial system is expressed as q m =[q 1t ,q 2t ,q 3t ,q 4t , where q 1t ,q 2t ,q 3t is the vector part of the measurement value, q 4t is the scalar part, and the subscript t represents the sampling time; perform two difference calculations on two adjacent q m to obtain the angular acceleration of the star sensor relative to the inertial system:

[0020] aqdw=[aqdw x aqdw y aqdw z

[0021] where aqdw x , aqdw y , aqdw z are the components of the angular acceleration of the star sensor relative to the inertial system in the star sensor measurement system;

[0022] At the same time, collect the three-axis angular acceleration Accdw = [Accdw x Accdw y Accdw z output by the payload accelerometer; where Accdw x , Accdw y , Accdw z are the three-axis components of Accdw, and Accdw is used as the theoretical satellite angular velocity;

[0023] Establish the star sensor installation matrix calibration formula as follows:

[0024] Accdw = C as ·aqdw

[0025] where C as is the true installation matrix of the star sensor relative to the payload accelerometer coordinate system to be estimated. During the above attitude maneuver, collect the measurement data of the star sensor and the payload accelerometer in real time and perform data processing. Store the processed data in the sliding window sequence. Define the length of the sliding serial port sequence as 9 according to the number of parameters to be estimated. Whenever the sliding window sequence data is updated once, use the least squares method to perform a real-time online recursive estimation of C as ; ​

[0026] Based on the recursive estimation results, use C as to assign and update the initial installation matrix C bs , thereby establishing a unified inertial measurement reference between the star sensor and the payload accelerometer.

[0027] In an embodiment of the present invention, a unified inertial measurement reference is established between multiple star sensors. Taking two star sensors as an example, it includes: First, eliminate the errors caused by in-orbit vibration and thermal alternation between the two star sensors. Taking one of the star sensors as a reference, use the measurement quaternion to perform real-time calibration and correction of the installation matrix between the two star sensors, and the calibration reference is dynamically adjusted according to the relationship between the solar vector direction and the angle between the star sensor optical axis.

[0028] In an embodiment of the present invention, using the measurement quaternion to perform real-time calibration and correction of the installation matrix between two star sensors includes:

[0029] Define the output quaternion of star sensor 1 as q Am1 , q Am2 , q Am3 , q Am4 , and the output quaternion of star sensor 2 as q Bm1 , q Bm2 , q Bm3 , q Bm4 . The initial installation matrix of star sensor 1 relative to the satellite is C Abs = [X Abs Y Abs Z Abs , where X Abs , Y Abs , Z Abs are the direction cosines of the measurement coordinate system of star sensor 1 in the satellite body system respectively. The initial installation matrix of star sensor 2 relative to the satellite is C Bbs = [X Bbs Y Bbs Z Bbs , where X Bbs , Y Bbs , Z Bbs are the direction cosines of the measurement coordinate system of star sensor 2 in the satellite body system respectively; use the quaternions output by the star sensors to establish the real-time attitude of the star sensors relative to the J2000 inertial coordinate system, and obtain the three-axis inertial space attitude matrices C Abm and C Bbm of star sensor 1 and star sensor 2:

[0030]

[0031] Taking the star sensor 1 as the calibration reference, constructing the reference attitude of the integrated installation frame coordinate system relative to the inertial space, and then calculating the actual on-orbit installation matrix C of the star sensor 2 by using the reference attitude and the attitude measured by the star sensor in real time nBbs :

[0032]

[0033] Then calculate the error matrix C between the ground initial installation matrix and the actual on-orbit installation matrix of the star sensor 2 ssd , and the formula is as follows:

[0034] C ssd = C nBbs ·C Bbs

[0035] Calculate the installation deviation angles of the three axes of the star sensor according to the X-Y-Z rotation sequence ψ ssd :

[0036]

[0037] θ ssd = asin(C ssd [2][0])

[0038]

[0039] For θ ssd , ψ ssd After filtering, we get Construct the installation correction matrix C Δds , specifically as follows:

[0040]

[0041] Perform the correction and update of the initial installation matrix C Bbs of the star sensor 2, specifically as follows:

[0042] C Bbs = C Δssd ·C Bbs .

[0043] In an embodiment of the present invention, the reference dynamic adjustment of the star sensor includes: when the star sensor 1 is used as the calibration reference, calculate the included angle Z1S between the optical axis ZAbs of the star sensor 1 and the solar vector Si, Z1S = ZAbs·Si. If Z1S is greater than SLA, then the star sensor 2 is calibrated with the star sensor 1 as the reference, where SLA is the stray light suppression angle of the star sensor; otherwise, if the included angle Z2S between the optical axis of the star sensor 2 and the solar vector is greater than SLA, then the calibration reference is set to the star sensor 2, and the star sensor 1 is calibrated with the star sensor 2 as the reference.

[0044] The present invention has the following beneficial effects compared with the prior art:

[0045] (1) The method of the present invention can be applied to a spacecraft system with integrated installation of a type of star sensor and a payload accelerometer inertial sensor in the cabin. By adopting the integrated structure installation design, thermal stability and stray light protection design, calibration of the payload accelerometer and the star sensor, calibration between star sensors, and calibration reference dynamic adjustment method, the star sensors participating in attitude determination always maintain a unified inertial reference, comprehensively realizing the establishment of a long-term high-stability inertial measurement reference for the satellite, and meeting the usage requirements of the satellite platform and the payload accelerometer inertial sensors.

[0046] (2) The present invention adopts the "same quality and same material" integrated installation of the star sensor and the payload accelerometer inertial sensor, and the flexible structure installation of the payload accelerometer installation base and the whole satellite structure, effectively reducing the installation deviation caused by structural deformation.

[0047] (3) The present invention adopts a two-stage separated light-shielding design method, which meets the requirements of stray light suppression under the complex conditions of the star sensor installed inside the spacecraft, and at the same time can isolate the influence of external heat flux radiation on the optical system structure of the star sensor, solving the problems of stray light suppression and isolation of heat source disturbance for the star sensor installed in the spacecraft cabin.

[0048] (4) The present invention uses the spacecraft attitude maneuver as the calibration excitation, utilizes the outputs of the payload accelerometer and the star sensor during the maneuver to establish a calibration equation, and takes the payload accelerometer as the reference to perform on-line real-time calibration of the installation deviation of the star sensor relative to the payload accelerometer, realizing the unification with the inertial measurement reference of the payload accelerometer. Description of the Drawings

[0049] Figure 1 It is a schematic diagram of the installation of the two-stage light-shielding cover of the present invention.

[0050] Figure 2 It is a schematic diagram of the heat and stray light protection of the two-stage light-shielding cover of the present invention. Detailed Embodiment

[0051] To make the objectives, technical solutions and advantages of the present invention clearer, the following will further describe the embodiments of the present invention in detail with reference to the drawings.

[0052] A method for establishing a high-stability inertial reference for a satellite star sensor, including:

[0053] (1) Integrated installation of star sensor and payload accelerometer: The star sensor and the payload accelerometer are integrally installed through a base with extremely high thermal stability. The main structural material of the star sensor optical system and the base bracket both adopt aluminum matrix silicon carbide with a small coefficient of thermal expansion. The same material ensures good thermal sensitivity consistency. In addition, the base is flexibly connected to the whole satellite structure by flexible angle bars made of titanium alloy, realizing the thermal deformation decoupling between the base and the whole satellite structure, and further reducing the structural thermal sensitivity.

[0054] (2) Separated two-stage light shield: To meet the requirements of stray light suppression for the star sensor inside the cabin, a two-stage light shield is designed according to the structural characteristics. The first-stage light shield is fixedly connected to the star sensor body, and its form is similar to the traditional star sensor light shield. The second-stage light shield is in an independent state and is installed at the second-stage installation surface outside the satellite cabin. In the satellite layout installation, the two-stage light shield adopts a non-contact installation method, which can avoid the influence of external heat flux on the optical system of the star sensor body and isolate the heat transfer path. The installation schematic diagram of the two-stage light shield is as Figure 1 shown.

[0055] (3) Segmented thermal control coating and stray light protection: To prevent stray light in the spacecraft cabin from entering the star sensor through slits and to ensure good temperature gradient effects for the star sensor optical system and its main structure, a circular thermal control multi-layer with a certain unit thickness is coated on the outer surface of the two-stage light shield. A thermal control multi-layer with a thickness of N1 is coated at a distance of 2 mm inward from the end face of the light exit of the first-stage light shield. The remaining part of the first-stage light shield is coated with a thermal control multi-layer with a thickness of N2 according to the thermal design situation, where it is necessary to ensure that N2 > N1 + the wall thickness of the second-stage light shield. The thermal control multi-layer on the outer surface of the second-stage light shield is coated according to the specific thermal design, and the installation surface of the second-stage light shield is blackened to prevent stray light interference caused by multiple reflections. The schematic diagram of the heat flux and stray light protection of the two-stage light shield is as Figure 2 shown ( Figure 2 the thermal control multi-layer on the outer surface of the second-stage light shield is not shown in the figure).

[0056] (4) Calculate the inertial attitude of the payload accelerometer using the output of the star sensor and establish the inertial attitude model q acc = C b ·C s ·q m , where q acc is the measured value of the payload accelerometer relative to the inertial system, C b is the relationship matrix between the optical reference axis of the star sensor and the coordinate system of the payload accelerometer, C s is the relationship matrix between the measurement axis of the star sensor and its own optical reference axis, and q m is the measured value of the star sensor measurement system relative to the inertial system. Denote C bs = C b ·C sRepresents the initial installation matrix of the star sensor relative to the payload accelerometer, C s 、C b Obtained through ground precision measurement.

[0057] (5) Calibrate the star sensor installation matrix based on the payload accelerometer: Use the spacecraft actuator to perform attitude maneuvers with a fixed angular acceleration in a certain order, and calculate the installation deviation between the star sensor and the payload accelerometer using the outputs of the star sensor and the payload accelerometer under constant acceleration. The quaternion representing the measured value of the star sensor measurement system relative to the inertial system is q m =[q 1t ,q 2t ,q 3t ,q 4t , where q 1t ,q 2t ,q 3t is the vector part of the measured value, q 4t is the scalar part, and the subscript t represents the sampling moment; perform two differential calculations on two adjacent q m to obtain the angular acceleration of the star sensor relative to the inertial system:

[0058] aqdw=[aqdw x aqdw y aqdw z , where aqdw x 、aqdw y 、aqdw z are the components of the angular acceleration of the star sensor relative to the inertial system in the star sensor measurement system; at the same time, the control computer communicates to collect the three-axis angular acceleration Accdw output by the payload accelerometer Accdw=[Accdw x Accdw y Accdw z ; where Accdw x 、Accdw y 、Accdw z are the three-axis components of Accdw, and Accdw is used as the theoretical satellite angular velocity.

[0059] Establish the star sensor installation matrix calibration formula as follows:

[0060] Accdw=C as ·aqdw

[0061] where C asis the true installation matrix of the star sensor relative to the load accelerometer coordinate system. During the above-mentioned attitude maneuvers, the measurement data of the star sensor and the load accelerometer are collected in real time and processed. The processed data is stored in a sliding window sequence. The length of the sliding serial port sequence is defined as 9 according to the number of parameters to be estimated. Whenever the data in the sliding window sequence is updated once, the least squares method is used to calculate C as for a real-time online recursive estimation.

[0062] According to the recursive estimation results, use C as to assign and update C in step (4) bs The initial installation matrix, thus establishing a unified inertial measurement reference between the star sensor and the load accelerometer.

[0063] (6) On the basis of the calibration in step (5), a unified inertial measurement reference is established for multiple star sensors and load accelerometers. The dynamic adjustment of the mutual calibration of multiple star sensors includes: First, establish a unified inertial reference. Eliminate the errors caused by vibration and thermal alternation during the on-orbit section between two star sensors. Taking one of the star sensors as the reference, use the measured quaternion to perform real-time calibration and correction of the installation matrix between the two star sensors. The calibration reference is dynamically adjusted according to the relationship between the direction of the sun vector and the angle between the optical axis of the star sensor. Define the output quaternion of star sensor 1 as q Am1 、q Am2 、q Am3 、q Am4 (scalar), the output quaternion of star sensor 2 is q Bm1 、q Bm2 、q Bm3 、q Bm4 (scalar), the initial installation matrix of star sensor 1 relative to the satellite is C Abs =[X Abs Y Abs Z Abs , where X Abs 、Y Abs 、Z Abs are the direction cosines of the measurement coordinate system of star sensor 1 in the satellite body coordinate system respectively. The initial installation matrix of star sensor 2 relative to the satellite is C Bbs =[X Bbs Y Bbs Z Bbs , where X Bbs 、Y Bbs 、Z Bbs are the direction cosines of the measurement coordinate system of star sensor 2 in the satellite body coordinate system respectively. Using the quaternions output by the star sensors to establish the real-time attitude of the star sensors relative to the J2000 inertial coordinate system, the three-axis inertial space attitude matrices C Abm and C Bbm of star sensors 1 and 2 can be obtained:

[0064]

[0065] Taking the star sensor 1 as the calibration reference, construct the reference attitude of the integrated installation frame coordinate system relative to the inertial space, and then calculate the actual on-orbit installation matrix C of the star sensor 2 by using the reference attitude and the attitude measured by the star sensor in real time nBbs ,

[0066]

[0067] Then calculate the error matrix C between the ground initial installation matrix and the actual on-orbit installation matrix of the star sensor 2 ssd , the formula is as follows:

[0068] C ssd = C nBbs ·C Bbs

[0069] Generally, the error angles of the actual installation of the star sensor relative to the initial installation are all small change amounts, and calculate the installation deviation angles of the three axes of the star sensor according to the X-Y-Z rotation sequence θ ssd 、ψ ssd :

[0070]

[0071] θ ssd = asin(C ssd [2][0])

[0072]

[0073] For θ ssd 、ψ ssd After filtering, get Construct the installation correction matrix C Δds , specifically as follows:

[0074]

[0075] Perform the correction and update of the initial installation matrix C Bbs of the star sensor 2, specifically as follows:

[0076] C Bbs = C Δssd |C Bbs

[0077] Star sensor reference dynamic adjustment steps: When star sensor 1 is the calibration reference, calculate the angle Z1S between the optical axis ZAbs of star sensor 1 and the solar vector Si, Z1S = ZAbs · Si. If Z1S is greater than SLA (where SLA is the stray light suppression angle of the star sensor), then other star sensors are calibrated with star sensor 1 as the reference; otherwise, if the angle Z2S between the optical axis of star sensor 2 and the solar vector is greater than SLA, then the calibration reference is set to star sensor 2, and other star sensors are calibrated with star sensor 2 as the reference; and so on for multiple star sensors. The dynamic adjustment method can always unify the measurement inertial reference of the star sensors participating in attitude determination.

[0078] The content not described in detail in the specification of the present invention belongs to the well-known technology in the art.

[0079] Although the present invention has been disclosed above with preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solution of the present invention using the methods and technical content disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modification, equivalent change, and modification made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention all belong to the protection scope of the technical solution of the present invention.

Claims

1. A method for establishing a high-stability inertial reference for a satellite star sensor, characterized in that: include: The star sensor and load accelerometer are installed in an integrated manner; A separate two-stage sunshade is used to suppress stray light from the star sensor in the cabin; A segmented thermal control coating is used outside the two-stage sunshade, and the installation surface of the sunshade is treated to prevent stray light; The inertial attitude of the load accelerometer is calculated using the output of the star sensor, and the inertial attitude model of the load accelerometer is established; Based on the load accelerometer, the star sensor installation matrix is ​​calibrated to establish a unified inertial measurement benchmark between the star sensor and the load accelerometer. For multiple star sensors, a unified inertial measurement reference is established among the multiple star sensors.

2. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 1, wherein: The star sensor and the load accelerometer are integrated with the base, and the installation parts of the star sensor and the base are made of the same material; The base and the entire satellite structure are flexibly connected to achieve thermal deformation decoupling.

3. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 1, wherein: The two-stage sunshade comprises a primary sunshade and a secondary sunshade, the primary sunshade is fixedly connected to the star sensor body, and the secondary sunshade is installed at the secondary installation surface outside the satellite cabin, and the two-stage sunshades do not touch each other.

4. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 3, characterized in that: A thermal control multilayer with a thickness of N1 is coated on the light outlet end face of the first-level light shield by shrinking a certain distance inward, and the rest of the first-level light shield is coated with a thermal control multilayer with a thickness of N2 according to the thermal design results, and N2>N1+the wall thickness of the second-level light shield; The thermal control multiple layers on the outer surface of the secondary sunshade are coated according to the thermal design results.

5. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 3, characterized in that: The mounting surface of the secondary sunshade is blackened.

6. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 1, characterized in that: The inertial attitude model of the load accelerometer is: q acc =C b ·C s ·q m where q acc is the measurement value of the load accelerometer relative to the inertial system, C b is the relationship matrix between the optical reference axis of the star sensor and the load accelerometer coordinate system, C s is the relationship matrix between the star sensor measurement axis and its own optical reference axis, q m is the measurement value of the star sensor measurement system relative to the inertial system; Remember C bs =C b ·C s represents the initial installation matrix of the star sensor relative to the load accelerometer, C s , C b Obtained through ground precision measurement.

7. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 1, characterized in that: Based on the calibration of the star sensor installation matrix using the load accelerometer, a unified inertial measurement benchmark between the star sensor and the load accelerometer is established, including: Using the spacecraft actuator, perform attitude maneuvers with fixed angular acceleration in a certain order, and use the outputs of the star sensor and payload accelerometer under constant acceleration to calculate the installation deviation between the two: the quaternion of the measurement value of the star sensor measurement system relative to the inertial system is expressed as q m =[q 1t ,q 2t ,q 3t ,q 4t ], where q 1t ,q 2t ,q 3t is the measured value vector part, q 4t is the scalar part, the subscript t represents the sampling time; for two consecutive q m Perform two differential calculations to obtain the angular acceleration of the star sensor relative to the inertial system: aqdw=[aqdw x aqdw y aqdw z ] where aqdw x 、aqdw y 、aqdw z is the component of the angular acceleration of the star sensor relative to the inertial system in the star sensor measurement system; At the same time, the three-axis angular acceleration Accdw output by the load accelerometer is collected. x dx y dx z ]; Accdw x 、Accdw y 、Accdw z are the three-axis components of Accdw, and Accdw is taken as the theoretical satellite angular velocity; The calibration formula for establishing the star sensor installation matrix is ​​as follows: <h2 style=";text-align:left;direction:ltr">Accdw=C<h2 style=";text-align:left;direction:ltr"> as <h2 style=";text-align:left;direction:ltr"> aqdw Among them C as is the real installation matrix of the star sensor to be estimated relative to the load accelerometer coordinate system. During the above-mentioned attitude maneuver, the measurement data of the star sensor and the load accelerometer are collected in real time and processed. The processed data are stored in the sliding window sequence. The length of the sliding serial port sequence is defined as 9 according to the number of parameters to be estimated. Whenever the sliding window sequence data is updated, the least square method is used to calculate C as Perform a real-time online recursive estimation; According to the recursive estimation results, using C as Assignment update initial installation matrix C bs , thereby establishing a unified inertial measurement reference between the star sensor and the payload accelerometer.

8. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 7, characterized in that: A unified inertial measurement reference is established between multiple star sensors. Taking two star sensors as an example, the method includes: firstly, eliminating the errors caused by vibration during the orbital stage and thermal exchange between the two star sensors, taking one of the star sensors as the reference, and using the measurement quaternion to perform real-time calibration and correction of the installation matrix between the two star sensors. The calibration reference is dynamically adjusted according to the relationship between the solar vector direction and the angle between the optical axis of the star sensor.

9. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 8, characterized in that: Real-time calibration and correction of the installation matrix between two star sensors using measured quaternions include: Define the output quaternion of star sensor 1 as q Am1 ,q Am2 ,q Am3 ,q Am4 , the output quaternion of star sensor 2 is q Bm1 ,q Bm2 ,q Bm3 ,q Bm4 , the initial installation matrix of star sensor 1 relative to the satellite is C Abs =[X Abs Y Abs Z Abs ], where X Abs , Y Abs , Z Abs are the direction cosines of the measurement coordinate system of star sensor 1 in the satellite system, and the initial installation matrix of star sensor 2 relative to the satellite is C Bbs =[X Bbs Y Bbs Z Bbs ], where X Bbs , Y Bbs , Z Bbs are the direction cosines of the measurement coordinate system of star sensor 2 in the satellite system; the quaternion output by the star sensor is used to establish the attitude of the real-time star sensor relative to the J2000 inertial coordinate system, and the three-axis inertial space attitude matrix C of the measurement system of star sensor 1 and star sensor 2 is obtained. Abm and C Bbm : Star sensor 1 is used as the calibration reference to construct the reference attitude of the integrated installation frame coordinate system relative to the inertial space. The actual on-orbit installation matrix C of star sensor 2 is calculated using the reference attitude and the real-time measured attitude of the star sensor. nBbs : Then calculate the error matrix C between the initial ground installation matrix of star sensor 2 and the actual on-orbit installation matrix ssd , the formula is as follows: C ssd =C nBbs ·C Bbs Calculate the three-axis installation deviation angle of the star sensor according to the XYZ conversion sequence θ ssd , ssd : θ ssd =and(C ssd [2][0]) right θ ssd , ssd After filtering, we get Build and install the correction matrix C Δds , as follows: Perform initial installation of star sensor 2 matrix C Bbs The revisions and updates are as follows: C Bbs =C Δssd ·C Bbs 。 10. The method for establishing a high-stability inertial reference for a satellite star sensor according to claim 9, characterized in that: The dynamic adjustment of the star sensor reference includes: when the star sensor 1 is used as the calibration reference, the angle Z1S between the optical axis ZAbs of the star sensor 1 and the solar vector Si is calculated, Z1S=ZAbs·Si, if Z1S is greater than SLA, the star sensor 2 is calibrated with the star sensor 1 as the reference, wherein SLA is the stray light suppression angle of the star sensor; otherwise, if the angle Z2S between the optical axis of the star sensor 2 and the solar vector is greater than SLA, the calibration reference is set to the star sensor 2, and the star sensor 1 is calibrated with the star sensor 2 as the reference.

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