194 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 provides a solar synchronous orbit quick response SAR satellite load platform integrated configuration, which comprises a satellite body, an SAR antenna, a solar array, a phased array data transmission antenna, a star sensor and a body-mounted solar cell, the satellite body is in a quadrangular prism shape, the cross section of the satellite body is an isosceles trapezoid, the satellite body comprises a group of opposite parallel surfaces and a group of opposite inclined surfaces, the phased array data transmission antenna and the star sensor are respectively arranged on two parallel surfaces, and the SAR antenna and the solar array are respectively arranged on two inclined surfaces; the body-mounted solar cell piece and the solar array are installed on the same inclined plane and located on the inner side of the solar array, and the solar array comprises two solar array plates and has an unfolded state and a folded state. According to the method, load platform integration is taken as a design thought, so that the AIT process of the SAR satellite and the time required by AIT are greatly shortened; meanwhile, the attitude requirements of energy and load work are met, and the reliability of the initial injection stage is improved through sufficient energy.

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 provides an integrated navigation method combining starlight refraction positioning with inertial navigation in a near space, and belongs to the technical field of starlight refraction navigation positioning. The method comprises the following steps: firstly, determining a state model of starlight refraction positioning in a near space by taking an error equation of an inertial navigation system of a near space aircraft as a state equation; designing a near space starlight refraction positioning algorithm, identifying a refraction star by using a star sensor on the aircraft, acquiring a refraction angle as a measurement, and establishing a measurement model of starlight refraction positioning in a near space; and obtaining the optimal estimation value of the system state variable by using an extended Kalman filter (EKF) method. The method has the advantages of being suitable for near space, high in reliability, high in autonomy and the like, and can be finally used for pure astronomical positioning and astronomical / inertial integrated navigation.

The invention discloses a star point coordinate extraction and star sensor attitude optimization method based on deep learning. The whole process of pseudo star map generation, star point coordinate extraction and attitude calculation is optimized. Multi-factor interference such as spacecraft jitter and optical distortion is simulated through dynamic star point modeling, and a high-fidelity pseudo star map is generated; a noise adaptive detection network is designed to fuse denoising and key point detection, and star point positioning precision in a low signal-to-noise ratio environment is improved; and optimizing attitude solution, constructing a loss function based on star point coordinates, and performing back propagation to optimize attitude parameters to realize an algorithm closed loop. According to the method, real data can be expanded to realize effective training of the model, and the positioning precision is improved; the star point coordinate positioning precision is improved to a sub-pixel level (RMSElt; 0.3 pixel); an end-to-end optimization framework reduces error accumulation in a traditional assembly line, and the robustness of the system is improved by 40%.

The utility model provides a star-sensitive load support assembly and a satellite, and relates to the technical field of satellites. The utility model provides a star-sensitive load bracket assembly and a satellite. The star-sensitive load bracket assembly comprises a load, the carbon fiber bracket is arranged on the surface of the load; and the star-sensitive element is arranged on the carbon fiber bracket, and the star-sensitive element, the load and the carbon fiber bracket are fixedly mounted into a whole. The thermal expansion coefficient of the carbon fiber in the carbon fiber bracket is low, the thermal deformation of the material is small, and the change of the relative position caused by the thermal deformation between the load and the star sensor is reduced. The star-sensitive element is directly connected with the load into a whole through the carbon fiber bracket, so that a transmission path between the star-sensitive element and the load is shortened, random angle change caused by thermal deformation on the transmission path is further reduced, and the reorbit shooting precision and the integrity of the bracket assembly are improved.

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.

A laboratory assessment method for the stellar vector observation accuracy of a star sensor includes: establishing a star sensor calibration and accuracy assessment system, jointly modeling and calibrating the star sensor's intrinsic and extrinsic parameters to obtain calibration results; statically or dynamically rotating a turntable to collect data at certain angles or time intervals; converting the imaged star point coordinates in the collected data into star sensor observation vectors based on the calibrated intrinsic and extrinsic star sensor parameters and the corresponding lens distortion model and pinhole imaging model; and uniformly linearly transforming the star sensor observation vectors to the turntable's initial position coordinate system to obtain the observation vectors; estimating the concentration parameter of the obtained observation vectors; and converting the concentration parameter into the standard deviation parameter of a normal distribution to obtain the star sensor's stellar vector observation accuracy. This method can more accurately estimate the stellar vector observation accuracy of a star sensor, requires no changes to the existing system, and simplifies calculations.

The invention discloses a star sensor fixed star detection capability analysis method based on a polarization degree image, and belongs to the technical field of star sensor detection. In order to solve the problems of few detection means and low utilization rate of sky light radiation information during star observation of a near-earth space all-time star sensor, the method comprises the following steps: constructing a star sensor integrated with a focal plane micro-polarizer array, and acquiring four gray star maps through short exposure by using the star sensor integrated with the focal plane micro-polarizer array; a star sky Stokes vector diagram is calculated; calculating star sky polarization degree distribution to obtain a polarization degree star map; constructing a fixed star polarization signal-to-noise ratio calculation formula in the polarization degree star map by utilizing a physical process of converting radiation energy into polarization degree; the calculation formula calculates the polarization signal-to-noise ratio of the fixed star in the field of view; and comparing the polarization signal-to-noise ratio of the fixed star with a signal-to-noise ratio threshold value detected by the near-earth space all-time star sensor, and judging the detectable capability of the fixed star in the polarization degree star map. According to the method, the robustness of the near-earth space all-day star-sensitive technology to a complex environment is improved.

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 provides a laser terminal calibration method and system based on inter-satellite orbit mutual transmission, and the method comprises the steps: carrying out the static calibration of a laser terminal through employing a star sensor or ground assistance before the establishment of an inter-satellite communication link, and enabling the pointing error of an optical axis to be smaller than 0.1 degree; after a communication link is established, orbit parameters of two satellites are exchanged in real time, a relative orbit normal vector is calculated through coordinate transformation, and a fusion reference is dynamically generated in combination with a static calibration result. According to the method, the real-time track mutual transmission data is used as dynamic correction input, the initial static calibration result is not only relied on any more, and fusion benchmark optimization of calibration and track mutual transmission is achieved. Compared with a traditional static calibration method depending on a ground station or a single-star inertial reference, the method breaks through the precision bottleneck of static calibration through the fusion of the orbit dynamic data and the static calibration result, effectively eliminates the orbit dynamic error, improves the laser calibration precision to a micro-radian level, and improves the calibration precision. And the establishment efficiency and reliability of the inter-satellite laser link are greatly improved.

The invention relates to the technical field of fusion positioning, in particular to a multi-sensor fusion positioning method, which comprises the following steps of: acquiring original positioning data from a plurality of sensors in real time, and preprocessing the positioning data; designing an independent local state estimation filter for each sensor, and outputting a local state estimation value; inputting each local state estimation value into a federated filter for multi-source information fusion, and outputting an optimal positioning result after fusion; and generating a fused positioning track according to an optimal positioning result output by federated filtering. According to the invention, an independent local state estimation filter is designed to carry out preprocessing and local state estimation on data of multi-source sensors such as an inertial measurement sensor, a GNSS receiver and a star sensor, a federated filtering architecture is adopted to carry out covariance weighted fusion, and the measurement precision and dynamic characteristics of each sensor are fully considered.

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 invention discloses a rapid star map lossless compression algorithm, and aims to solve the problems of large memory space and low compression efficiency of the existing lossless compression algorithm. According to the algorithm, the efficient compression of star map data is realized through the combination of a context modeling module, a predictive coding module and a run length coding module. The method specifically comprises the steps that a flat area and a non-flat area are divided according to local gradient values, run length coding is adopted in the flat area, and predictive coding is adopted in the non-flat area; the predictive coding is realized through gradient quantization merging, predictive error calculation and Rice coding; and run length coding is realized by counting the length of continuous same pixel points and performing table look-up coding. The algorithm optimizes the number of context types, reduces the storage demand, and improves the compression efficiency by dynamically correcting the predicted value. The method is suitable for application scenes with high requirements on data integrity, such as star sensors.

The invention discloses a fixed star detection capability improving method fusing energy and polarization information, and belongs to the technical field of star sensor detection. In order to solve the problem that a near-earth space all-day star sensor is insufficient in fixed star detection capability in a strong atmospheric background stray light interference environment, the method comprises the following steps: exposing the sky once by using a star sensor integrated with a focal plane micro-polarizer array to obtain four gray star maps; calculating a polarization degree calculation value to obtain a polarization degree star map, and determining the sky background polarization degree; setting a polarization degree segmentation threshold value; performing threshold segmentation on the polarization degree star map to obtain a polarization degree star map segmentation result; determining a weight coefficient according to the radiation energy ratio of the fixed star to the sky background; and comparing the sky background polarization degrees, and performing weighting processing on the gray star map or the accumulated star map with the highest signal-to-noise ratio based on different conditions to obtain a final processed high-signal-to-noise-ratio star map. According to the invention, the fixed star detection capability is effectively improved.

This invention discloses an on-orbit evaluation method for the single-frame attitude accuracy of a star sensor, comprising: S1, calculating the single-frame attitude covariance matrix of the star sensor by combining the synthesized value of the variance of noise from all star vector measurements; S2, evaluating the three-axis attitude error of the star sensor based on the attitude covariance matrix; S3, calculating the single-frame attitude quality factor of the star sensor based on the three-axis attitude error value; and S4, evaluating the attitude error of the star sensor based on the quality factor. This invention achieves on-orbit real-time calculation and evaluation of the single-frame attitude accuracy of a star sensor by calculating the covariance matrix of the star sensor and constructing a "quality factor".

The utility model discloses a portable star sensor sun evasion performance testing device. The portable star sensor sun evasion performance testing device comprises a supporting mechanism, an angle adjusting mechanism and a light source module, the supporting mechanism is a total station tripod, and the angle adjusting mechanism is mounted on a top platform of the total station tripod; the driving end of the angle adjusting mechanism is connected with the light source module; and the light source module is connected with the portable power supply. The device has the advantages that the device comprises the light source module, the angle adjusting mechanism and the supporting mechanism, the light source module provides a light source, the angle adjusting mechanism adjusts the angle of the light source module, the supporting mechanism adopts a total station tripod and is convenient to carry, and the whole device can be rapidly arranged and put into use in the field; and carrying out sun avoidance performance test on the star sensor.

The invention provides an integrated calibration method, device and equipment for photoelectric non-uniformity of an optical star measurement sensor, and relates to the technical field of spacecraft celestial navigation. The method comprises the following steps: constructing lattice structure light to form array image spots on a star sensor, resampling to establish an actual PSF model, and realizing optical system aberration calibration of a corresponding area; sequentially displacing the array image spots, and calibrating the electrical response non-uniformity of each pixel by using the mapping relation between imaging data and PSF; and performing precise positioning on the image spot array after pixel response correction to obtain an image spot high-precision positioning result, and calibrating optical distortion of each pixel according to the high-precision positioning result. Therefore, the problems of photoelectric non-uniform coupling and limited calibration precision in calibration of the star measurement sensor are effectively solved, and high-precision and integrated calibration of optical aberration, distortion and electrical response of the star measurement sensor is realized.

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 present invention provides an on-orbit identification method for gyro nonlinearity and pointing system errors, comprising: calculating the platform angular velocity based on star sensor dynamic data; establishing an equation relationship between the gyro system error and the platform angular velocity; constructing observation equations for the nonlinear error and installation error angle of each gyro meter head; using the least squares method to perform on-orbit identification of the gyro meter head system error; and calculating the installation pointing angle of the gyro meter head in the body coordinate system. On-orbit identification of gyro system errors can effectively reduce the attitude angular velocity calculation error and gyro integrated attitude angle calculation error introduced by gyro nonlinearity and installation error in high-dynamic satellites. The present invention can simultaneously identify the nonlinear error and installation angle pointing error of the gyro meter head. The gyro meter head system error identification dimension is high, and the gyro system error estimation accuracy is high. This method can effectively improve the attitude angular velocity measurement accuracy and spatial pointing measurement accuracy of remote sensing satellites under high dynamic conditions.

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.