122 results about "Star sensor" patented technology

The invention relates to the technical field of spacecraft orbit control and space safety, and discloses a space debris orbit prediction and avoidance method, which comprises the following steps that: a ground computing center screens candidate targets and generates enhanced data packets for uploading; the spaceborne computer combines real-time navigation data to carry out geometric feature screening, and an instant high-risk target list is generated; controlling the star sensor to execute optical observation, and resolving an optical measurement relative orbit state by utilizing displacement superposition and Kalman filtering; selecting a high confidence coefficient or a conservative probability threshold according to the optical measurement state acquisition condition, and generating an avoidance instruction when the collision probability exceeds the limit; and the propulsion system responds to the instruction to execute avoidance, verifies the effect by confirming the semi-major axis variation and executes track recovery by selecting an aircraft. According to the method, the ground-air collaborative screening and shift superposition enhancement technology is adopted, the problems that the low-precision ephemeris false alarm rate is high and the detection capability of a satellite-borne sensor on a dark and weak target is insufficient are solved, and the accuracy of evasion decision making and the execution reliability are improved.

The invention relates to a satellite laser autonomous calibration method based on multi-star sensor dynamic collaboration, multi-star sensors comprise a plurality of platform cabin star sensors and a plurality of laser self star sensors, and the platform cabin star sensors and the laser self star sensors are arranged on a satellite side plate. The satellite laser autonomous calibration method comprises the following steps: determining an installation matrix of a multi-star sensor and a satellite body system, and completing installation matrix parameter binding; by taking the platform cabin star sensor as an attitude calibration reference, correcting the mounting matrixes of the other platform cabin star sensors and the laser star sensor so as to complete in-orbit autonomous calibration; and calculating the residual error of the platform cabin star sensor and the laser star sensor, when the error quaternion residual error is lower than a set threshold value, ending the autonomous calibration process, and broadcasting the quaternion measured by the multiple star sensors. According to the installation layout of the satellite laser and the star sensor, the calibration process is simplified, a ground station does not need to participate, and the satellite can autonomously complete the calibration process in orbit.

The invention discloses a star sensor rapid tracking method based on dynamic window blind restoration, and belongs to the field of star sensor algorithms. The method comprises the following steps: estimating a coordinate estimation value of each star in an overlapped view field on a star map at the next moment through a historical star map, and extracting a local window image of a star point corresponding to each star in the overlapped view field; gaussian filtering, top-hat transformation and connected domain processing are carried out on the local window image, star point trailing information is extracted based on a Zernike moment sub-pixel edge extraction method, and a star point area is restored by using a Richardson-Lucy algorithm; then, star point centroids are extracted from the restored image, star points which are tracked successfully are screened out in combination with motion consistency, and initial attitude parameters are calculated through a QUEST method; and finally fusing the star point information newly entering the field of view to obtain attitude parameters of the star sensor under the reference coordinate system. According to the method, the image traversal range is greatly reduced, and the method has important engineering significance.

The invention relates to a satellite laser attitude data broadcasting method based on a multi-star sensor, the multi-star sensor comprises a platform cabin star sensor and a laser self-star sensor, and the platform cabin star sensor and the laser self-star sensor are arranged on a satellite side plate. The method comprises the following steps: judging the validity of a platform cabin star sensor and a laser star sensor, and determining the source of laser attitude data to be broadcasted; calculating a time difference; taking the attitude quaternion measured by a plurality of star sensors as a quaternion initial value, combining the time difference and the satellite angular velocity, and performing time delay correction by using a time delay integral correction algorithm to obtain an attitude quaternion after correcting the attitude deviation; and broadcasting the attitude quaternion after correcting the attitude deviation. According to the invention, the star sensor data of the laser is introduced and is combined with the data of other star sensors for comprehensive processing, so that the precision of attitude measurement is remarkably improved.

The invention discloses a star point motion parameter estimation method for a high-dynamic star sensor. The method specifically comprises the following steps: (1) acquiring a star map under a dynamic condition through a fixed star sensor; (2) carrying out filtering preprocessing on the fixed star sensor image obtained in the step (1); (3) carrying out threshold segmentation on the image processed in the step (2) and carrying out connected domain extraction to obtain rough star points; (4) selecting the brightest star point in the star points obtained in the step (3); (5) obtaining a sub-pixel star point track by using the brightest star point adjacent area obtained in the step (4) and adopting a Zernike moment sub-pixel edge extraction method; and (6) utilizing the sub-pixel star point trajectory obtained in the step (5), and adopting least square fitting to obtain star point motion parameters. The method has the beneficial effects that the precision of star point trailing length and angle estimation is improved, so that the extraction precision of star points in a dynamic environment is improved, and the precision of a fixed star sensor is improved.

The embodiment of the invention provides a data processing method and device and electronic equipment, and relates to the technical field of satellites.The method comprises the steps that star sensor data obtained by collecting spacecraft attitude information through a star sensor and gyroscope data obtained by collecting spacecraft attitude information through a gyroscope are obtained and comprise data from the k-n moment to the k moment; carrying out fusion and linear interference filtering processing on the data of the star sensor and the gyroscope at each moment by using Kalman filtering to obtain preliminary prediction data of attitude information, and calculating a first residual value of the data of the star sensor; the star sensor data, the gyroscope data and the first residual value from the k-n moment to the k moment are input into the trained LSTM network, a second residual value of the preliminary prediction data at the k moment is predicted, and the second residual value represents an error of nonlinear interference; and compensating the second residual value to the preliminary prediction data to obtain attitude information of the spacecraft at the moment k. According to the embodiment of the invention, the error of spacecraft attitude information acquisition is reduced.

The application discloses a fault-tolerant implementation method of a star sensor storage system against single event upsets, and solves the short board of the existing star sensor storage system against single event upsets and the technical problem that the processor internal FLASH capacity is insufficient to implement triple modular redundancy and the like.The method comprises three parts of power-on loading fault-tolerant reading, external SRAM star table dynamic refreshing and internal FLASH data automatic repairing.The ECC error processing is accurate and efficient, and is suitable for the hardware characteristics.The hardware parameters of 32 bytes corresponding to 10 bits of ECC and 32 byte error positioning granularity are adapted, single-bit error is only recorded and not switched to backup, the data accuracy is guaranteed by relying on hardware error correction, double-bit error is temporarily switched to backup reading, the error correction efficiency and data reading fluency are considered, and the real-time operation requirement of the star sensor is met.The application greatly improves the reliability of the star sensor in long-term operation in orbit, and effectively reduces the failure rate in orbit and operation and maintenance cost.

The star sensor acquires a star point image through a detector, a data processor scans the obtained star point image using a 3*3 template; judges whether the gray value of the pixel point in the center of the template is greater than an image threshold T1 and whether the gray values of the 8 pixels around the pixel point are all less than the image threshold T1, and obtains a star point image after removing isolated points; the data processor divides the star point image after removing isolated points into blocks, divides the size of a sub-block into 32*32 pixels, and then calculates the image mean value in each sub-block; judges the mean value of each sub-block, and filters out the background and noise with high gray value; performs a pixel traversal operation on the star point image after removing isolated noise using the determined image threshold T1, and sends the pixels not less than the threshold T1 to a target extraction module. The star point image processing method under a complex background can judge and filter the star point image under a complex background, and can stably output the attitude information of the star sensor.

The invention discloses a method and a system for determining an optimal attitude of a high-dynamic star sensor, and belongs to the technical field of spaceflight measurement and control. Aiming at anisotropic distribution of centroid positioning errors caused by'trailing 'of a star point image under high dynamic condition, the method comprises the following steps: acquiring a trailing star point image and extracting an observation star vector; calculating a prior attitude matrix; analyzing the anisotropy error and trajectory orientation of the center of mass in the motion trajectory direction and the vertical direction; calculating an error component in an image plane coordinate axis direction; and constructing a weighted least square problem based on the error component, and updating a prior attitude to obtain an optimal attitude matrix. By accurately matching the weight of anisotropy error distribution, noise in a large error direction is effectively suppressed, and information in a high-confidence direction is fully utilized, so that the limitation of a traditional method is overcome, attitude estimation under a high-dynamic condition is ensured to meet the minimum variance unbiased estimation requirement, and the accuracy of attitude estimation is improved. And the precision and reliability of attitude determination are obviously improved.

The application provides a multi-pose measuring device relative installation error calibration and compensation method and system, and relates to the technical field of spaceflight. The application comprises the following steps: S1, measuring a star sensor pose and converting coordinates; S2, calibrating high-precision inertial navigation and star sensor relative installation errors; S3, calibrating high-precision inertial navigation and low-precision inertial navigation self-alignment relative installation errors; and S4, compensating for the relative installation errors. The application calibrates and compensates for the relative installation errors among the high-precision inertial measurement combination, the low-precision inertial measurement combination and the star sensor by using the inertial navigation static self-alignment and the star sensor static pose measurement in the static state, so as to realize the consistency of the space pose reference in the dynamic multi-source combined navigation information fusion.

The invention provides a multi-star sensor attitude fusion calculation method and system, and the method comprises the steps: S1, calculating the projection of an optical axis of a star sensor j in a spacecraft body system according to an installation quaternion from a measurement coordinate system of the star sensor j to the spacecraft body system; s2, according to the quaternion from the inertial system of the star sensor j to the spacecraft body system, calculating the projection of the optical axis of the star sensor j under the inertial system; and S3, according to the vector sum, in combination with a Moore-Penrose generalized inverse theory of the matrix, calculating to obtain a spacecraft multi-star sensor fusion attitude conversion matrix. The invention aims to fully utilize the attitude information of all the star sensors carried by the spacecraft to carry out multi-star sensor attitude fusion calculation when the quantity of the star sensors carried by the spacecraft is more than 2, so as to realize high-precision spacecraft attitude measurement.

This invention relates to star sensors, specifically to lenses for shortwave imaging detectors in all-weather star sensors and their design methods. It addresses the problem that shortwave imaging detectors are not yet widely used in all-weather star sensors, resulting in a lack of suitable optical lenses. The invention comprises a first lens, a second lens, a third lens, a fourth lens, a fifth lens, a sixth lens group, and a seventh lens arranged sequentially along the optical path. The sixth lens group is a cemented lens, including a first cemented lens and a second cemented lens arranged sequentially along the optical path. This six-lens-plus-one-cemented-lens configuration results in low lens distortion, uniform dispersion under a telecentric optical path, which is beneficial for centroid extraction and improves the lens's response in the near-infrared band, enabling all-weather detection.

The application relates to a kind of moving base star sensor and high-precision time synchronization method of inertial navigation, and is mainly related to the technical field of integrated navigation, and the main steps include: real-time acquisition of the motion parameter data output by IMU;Using second-order extrapolation algorithm, based on IMU historical data, the equivalent IMU data corresponding to the star sensor data update time is calculated, the output frequency difference is adapted;Through asynchronous Kalman filtering algorithm, combined with IMU data and star sensor measured data, time update and measurement update are completed, and dynamic error is inhibited;Using high-precision crystal oscillator to optimize the clock source of inertial navigation system, the inherent error of clock output is reduced;Through multi-link cooperation, the time delay error of star sensor is compensated;Using FPGA interrupt latch, time alignment at the hardware level of star and inertial is realized.The method can effectively solve the time asynchronous problem of star sensor and inertial navigation data in the online calibration stage during flight, and guarantee the navigation accuracy and stability of integrated navigation system.

The invention discloses an aerodynamic radiation transfer calculation method based on equivalent Lambert body surface element tracing, and the method comprises the steps: simplifying a conventional radiation transfer equation into a form based on the Kirchhoff's law based on the optically thin and non-scattering characteristics of a hypersonic velocity external flow field; the brightness spectrum of the equivalent Lambert area light source is obtained by integrating the emission coefficient along the observation sight line; then, enabling the transmission process of aerodynamic radiation of an aircraft external flow field to the star sensor to be equivalent to a Lambert surface radiation source located at the top end of a light shield, discretizing a light source surface and a detector image surface into infinitesimal elements, and performing tracing calculation on irradiance contributions generated by all the light source surface infinitesimal elements to a single detector image element; finally, the irradiance distribution of the pneumatic radiation on the star sensor detector is efficiently and accurately obtained; according to the method, the transmission process of aircraft flow field pneumatic radiation to the star sensor is reasonably simplified, the calculation efficiency of pneumatic radiation transmission is remarkably improved, and the method is suitable for engineering prediction and evaluation.

The invention relates to a star sensor, in particular to a lens for a short-wave imaging detector in an all-day star sensor and a design method of the lens, and solves the problem that the short-wave imaging detector is not popularized on the all-day star sensor at present, so that almost no applicable optical lens exists. The optical system comprises a first lens, a second lens, a third lens, a fourth lens, a fifth lens, a sixth lens group and a seventh lens which are sequentially arranged along an optical path, the sixth lens group is a glued lens and comprises a first glued lens and a second glued lens which are sequentially arranged along the optical path, the mode that the six lenses and one glued lens group are adopted, the lens distortion is small, and the imaging quality is high. The defocused spots are uniform in a quasi telecentric light path, so that the mass center extraction of the light spots is facilitated, the response of the lens in a near-infrared band is improved, and all-day detection can be carried out.

A method for measuring and calibrating installation errors of a star sensor and an inertial navigation system relates to the technical field of integrated navigation, and specifically comprises the following steps: establishing a star sensor body coordinate system, an inertial navigation system body coordinate system and a space reference coordinate system; the method comprises the following steps: constructing an attitude transfer model containing an installation error angle based on a coordinate conversion relation of a star sensor and an inertial navigation system in space; an observation equation is constructed on the basis of an attitude matrix output by a star sensor and navigation parameters output by an inertial navigation system, and online identification and calibration of installation errors are achieved in combination with a linear Kalman filtering algorithm. The method does not need to depend on external reference equipment, accurate solving of the installation error can be completed through measurement data of the star sensor and the inertial navigation system carried by a carrier, the method is suitable for an inertia / starlight integrated navigation system of an aircraft, and the three-axis installation error generated when the star sensor and the inertial navigation system are fixedly connected and installed can be accurately identified and calibrated.

The application discloses a method for improving star detection capability by fusing energy and polarization information, and belongs to the technical field of star sensor detection. In order to solve the problem of insufficient star detection capability of a near-earth space all-time star sensor in a strong atmospheric background light interference environment, the star sensor integrated with a micro-polarization plate array is used to obtain four gray star images through one-time exposure of the sky; a polarization degree calculation value is calculated to obtain a polarization degree star image and determine a sky background polarization degree; a polarization degree segmentation threshold is set; the polarization degree star image is subjected to threshold segmentation to obtain a polarization degree star image segmentation result; a weight coefficient is determined according to the ratio of the radiant energy of the star and the sky background; the sky background polarization degree is compared, and the gray star image or the accumulated star image with the highest signal-to-noise ratio is subjected to weighted processing based on different conditions to obtain a final high signal-to-noise ratio star image after processing. The star detection capability is effectively improved.

The application discloses a satellite solar cell array rotation angle estimation method, which comprises the following steps: step 1, establishing an earth inertial coordinate system, a satellite body coordinate system and a solar cell array coordinate system, and defining a satellite solar cell array rotation angle; step 2, establishing a state equation of an augmented state composed of a sun vector, the solar cell array rotation angle and a gyro constant drift noise; step 3, calculating an augmented state quantity estimation value by using an extended Kalman filter; and step 4, calculating a satellite solar cell array rotation angle estimation value. The method fuses gyro angular velocity and sun sensor vector data, constructs a nonlinear dynamic model containing a sun vector, a rotation angle and a gyro bias, and designs an EKF algorithm to estimate the solar cell array rotation angle in real time, thereby solving the strong dependence of a traditional scheme on a star sensor and a mechanical sensor, and significantly improving the autonomous recovery capability and energy guarantee reliability of a satellite in an extreme fault scene.

The application relates to a method for positioning the position of a star point in a star map, relates to the technical field of computer image processing, and solves the problems of low efficiency of the existing star map processing method and low output frequency of a star sensor, etc. The method uses a histogram mode function of a detector, firstly collects a star map, obtains a row of data containing the mean values of all columns of the histogram and carries out searching on the data, obtains the column corresponding to the peak value data, then configures the detector into a normal output mode, collects a star map again, carries out accumulation and processing on the data of all rows, obtains a column of data containing the accumulation sum of all rows of data and carries out searching on the data, obtains the row corresponding to the peak value data, and finally, according to all column and row data, the data are taken as the center points of star point coordinates, the star point pixels in the stored star map are searched in sequence, and relevant information such as the star point centroid is obtained. The method does not need to search the whole star map, greatly improves the star point searching efficiency, and shortens the star point extraction time.

The application discloses a star-inertial integrated outfield dynamic calibration method based on a refracting surface constraint mechanism, and relates to the technical field of calibration methods, and specifically discloses the following technical scheme: a relationship model of an observed star position and a zenith direction is established based on a refracting surface constraint mechanism and observation imaging of a star, a first zenith direction measurement model under a star sensor system corresponding to an observation time and an observation position is established based on the relationship model; a second zenith direction measurement model under an inertial navigation system of the zenith direction of a current observation position is established through a filter combination model based on real-time measurement values of angular velocity and angular acceleration of the inertial navigation system; a zenith direction integrated conversion matrix model between the inertial navigation system and the star sensor system is established by simultaneously solving the first zenith direction measurement model under the star sensor coordinate system and the second zenith direction measurement model under the inertial navigation system; and a star-inertial integrated outfield optimization calibration model is established by changing the longitude and latitude of the observation position and the observation direction for multiple times, so that star-inertial integrated outfield calibration is carried out.

This invention provides a novel integrated attitude sensor for satellites, relating to the fields of onboard attitude control and electronics. It includes: a star sensor body; a solar sensor body mounted on the star sensor body; a magnetometer body within the star sensor body; and a gyroscope body within the star sensor body. By integrating traditional discrete structures into a single device, the size and weight are reduced, and costs are significantly lowered. Magnetometer and gyroscope data are fused in real-time within an FPGA. This device features rapid adaptability, modular design supporting plug-and-play functionality, and integrated design improving EMC performance. It solves the problems of existing traditional satellite attitude control systems, where various attitude sensors are independently configured and equipped with dedicated processors, resulting in high hardware redundancy, large overall size and weight, high power consumption and cost, hindering further improvements in attitude measurement accuracy and dynamic performance, and leading to high overall design complexity.

The invention discloses a short-wave infrared near-earth space all-time polarization star sensor star map simulation method, and relates to the technical field of star sensors. The problems that polarization information is not considered in current near-earth space all-day star map simulation, and a star map processing algorithm lacks a data set are solved. The method comprises the following steps: determining the orientations of the sun and the star according to observation time, observation position and observation direction; light field modeling of a target and a background is realized according to the geographic position of the observation point; determining an exposure time simulation value according to the polarization degree of the detected sky area; the radiation energy information is converted into the photoelectron number through photoelectric conversion; superposing detector noise and image quality degradation influence to obtain an original star map; and resolving a polarization degree star map according to the energy information received by the polarization star sensor. The method has a good application prospect in the field of all-day astronomical autonomous navigation of a near-earth space platform, and can provide key data support for optimization design and navigation algorithm test of the polarization star sensor.

The invention belongs to the technical field of photoelectric navigation, and discloses an all-optical navigation method and device of an integrated polarization sun fixed star inertial sensor. The device comprises a polarized light sensor, a sun sensor, a star sensor, an inertial measurement unit, an information comprehensive processing module, a clock module and a power supply module, and all-time and high-precision navigation positioning is realized through cooperative work of multiple sensors. The method comprises the following steps: selecting an optimal navigation observation source according to radiation differences of fixed stars, the sun and atmospheric polarized light in different time periods, and completing integrated navigation positioning with an inertial navigation system. The system overcomes the limitation of a single navigation mode, remarkably improves the all-time navigation capability of the system, has the advantages of small size, light weight, low power consumption, high precision and the like, and is suitable for navigation application scenes sensitive to space and cost, such as small and medium-sized unmanned aerial vehicle platforms.

The present application relates to the technical field of spacecraft attitude control, and particularly relates to a load attitude determination method and device based on an included angle monitoring star sensor, wherein the method comprises: obtaining angle conversion relationships between a reference mirror coordinate system and a star sensor measurement coordinate system, between a central prism coordinate system and the reference mirror coordinate system, and between the central prism coordinate system and a star sensor monitoring coordinate system in a ground phase; obtaining angle conversion relationships between the central prism coordinate system and the star sensor monitoring coordinate system, and between the star sensor measurement coordinate system and a J2000 coordinate system in an on-orbit phase; and combining a corresponding relationship between the star sensor measurement coordinate system and the star sensor monitoring coordinate system to determine an angle conversion relationship between the central prism coordinate system and the J2000 coordinate system in the on-orbit phase. The present application can obtain high-precision load pointing.

The invention discloses a hybrid star sensor daytime star measuring device, which relates to the technical field of celestial navigation, and comprises a star measuring assembly, a two-dimensional fast reflector, a control circuit board and a shell, all the parts are integrally installed on the shell, the satellite measuring assembly is arranged in the shell and comprises a telephoto lens, a multi-band-pass satellite observation window and a satellite measuring camera, reflected light rays are converged on the satellite measuring camera arranged behind the telephoto lens through the telephoto lens, the multi-band-pass satellite observation window is arranged between the telephoto lens and the satellite measuring camera and divided into a plurality of areas, and the satellite measuring camera is arranged in the shell. Only light of different specific wavebands can pass through each region, and the light is moved to the optimal region of the multi-band-pass star observation window by combining small-angle adjustment of the reflector, so that rapid selection of the wavebands is realized; the two-dimensional fast reflecting mirror and the satellite measuring assembly are coaxially arranged, the projection position is rotated and offset through two-axis control, and the small view field is expanded into the large view field by fast rotation and image splicing. The device solves the problem of background saturation of satellite measurement in the atmosphere in the daytime, and improves the signal-to-noise ratio and the adaptability.

The invention discloses an FPGA (Field Programmable Gate Array)-based star sensor high-frame-frequency attitude measurement method, which belongs to the technical field of star sensor attitude measurement, and comprises the following steps: a star point data processing module completes star point data processing and FPGA implementation: longitude and latitude data normalization of star points and pixel coordinate data normalization of the star points on a star sensor are realized on the FPGA; the attitude matrix calculation FPGA is realized: a star point data processing module is sequentially connected with three matrix multiplication modules and a matrix inversion module on the FPGA to form an attitude matrix solving module; and according to a generalized least square method, the three matrix multiplication modules and the matrix inversion module are combined to calculate an attitude matrix. According to the method, the attitude calculation time is shortened, and the attitude update rate of the star sensor is improved.

The present application relates to a kind of attitude prediction formula star tracking method, it is related to star sensor technical field.The method includes the following steps: obtaining historical angular velocity data;Predict the next frame angular velocity;Predict the next frame attitude quaternion;Project the sky area star to image plane;Window matching is carried out between projected star and the next frame observation star;Calculate the next frame accurate attitude quaternion;The target of window matching success is used for attitude solution, obtain the next frame accurate attitude;Repeat the above steps, realize autonomous star tracking.The attitude prediction formula star tracking method of the present application, by polynomial fitting historical angular velocity data, can quickly and accurately predict the next frame angular velocity, significantly improve real-time and accuracy;Can well adapt to large angular velocity maneuver, enhance the tracking ability of star sensor in complex dynamic environment;Reduce the dependence on external sensor, improve the independence and anti-interference ability of system, while simplifying the calculation process, improve overall computing efficiency.

The application relates to a method for establishing a high-stability inertial reference of a satellite star sensor, belonging to the technical field of spacecraft control, which comprises the following steps: integrating the star sensor and a load accelerometer; using a separated two-stage light shield to suppress the stray light of the star sensor in the cabin; using a sectional heat control coating outside the two-stage light shield, and performing anti-stray-light treatment on the mounting surface of the light shield; calculating the inertial attitude of the load accelerometer by using the star sensor output, and establishing an inertial attitude model of the load accelerometer; calibrating the installation matrix of the star sensor based on the load accelerometer, and establishing a unified inertial measurement reference between the star sensor and the load accelerometer; and establishing a unified inertial measurement reference between multiple star sensors.

The application discloses a method for improving the attitude precision of a single-line array satellite, and comprises the following steps: calculating the geometric positioning error of a control area according to a geometric positioning model; solving the low-frequency error model parameters through an image geometric positioning model; obtaining the variation law of the low-frequency error with time and latitude according to the low-frequency error model parameters, and calculating the low-frequency error value of the attitude at any time; and correcting the attitude value of the single-line array photographing time according to the obtained low-frequency error value, eliminating the low-frequency error caused by the star sensor, and further improving the geometric precision of the single-line array satellite.

The invention relates to a laser communication terminal calibration method and system for a large satellite constellation, and relates to the technical field of laser communication, and the method comprises the steps: configuring laser terminals in the constellation, configuring a first laser terminal with a fixed star sensor for a very few satellites as a laser master satellite, and configuring a second laser terminal for other satellites as a laser slave satellite; starting a relay calibration process by taking at least one self-calibrated laser master satellite as a starting point: calibrating between the laser master satellite and an uncalibrated laser slave satellite, then taking the calibrated laser slave satellite as a new calibration source, and continuing to calibrate other laser slave satellites; the process is repeated until all satellites are calibrated; in each calibration, pointing calibration is firstly executed, and then receiving and transmitting coaxial calibration is executed. According to the invention, a global unified reference is established through a very few star-sensitive terminals and relay calibration, the consistency of the pointing precision of the whole constellation is ensured, the constellation-level high-precision calibration is realized, and the constellation cost and weight are greatly reduced.