294 results about "Star sensor" patented technology

The invention discloses an agile satellite attitude fusion method and system considering multi-source reference difference, and belongs to the field of remote sensing satellite ground data processing, and the method comprises the steps: calculating a thermal environment factor based on satellite measurement data and satellite sensor installation parameters; extracting a space-time low-frequency error based on observation data of the star camera and the star sensor; performing normalization processing on the thermal environment factor, and constructing a relation model of the space-time low-frequency error and the thermal environment factor by adopting a multi-layer perceptron network to obtain compensated attitude data of the star sensor and the star camera; establishing a multi-source absolute attitude sensor data fusion model based on the compensated attitude data of the star sensor and the star camera; and carrying out absolute and relative attitude data fusion filtering processing on the processed gyroscope data and the fusion result of the star sensor and the star camera by adopting a bidirectional dynamic unscented Kalman filter to obtain a high-precision attitude data result. According to the method, the attitude determination precision of the agile satellite is improved, and the requirement of a high-precision remote sensing task is met.

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 the field of data processing, in particular to an on-orbit calibration method and device for a scale factor, and the method comprises the steps: obtaining attitude data of a spacecraft after maneuvering through a star sensor, converting the attitude data into an angular acceleration signal under a test mass coordinate system, constructing a Kalman filter in combination with an excitation voltage read by an inertial sensor, and carrying out the on-orbit calibration of the scale factor; on-orbit calibration of the angle scale factor is completed, and on-orbit calibration of the line scale factor is completed through the conversion relation between the angle scale factor and the line scale factor. By combining and considering related factors such as linear acceleration and angular acceleration, the calibration process can more comprehensively consider the motion state of the spacecraft, and more accurate calibration is provided. In the maneuvering process of calibrating the scale factor, other parameters such as centroid offset and rigidity of the inertial sensor of the spacecraft can be calibrated at the same time. Through the technical scheme of the invention, the additional mass increase can be reduced, the unnecessary control complexity is avoided, the reliability of the system is improved, and the cost is reduced.

The invention provides a satellite inertial installation error calibration method and system based on iterative filtering, and aims to solve the problem that the attitude measurement precision of a star sensor is influenced by an installation error between inertial navigation and the star sensor. The method comprises the following steps: constructing a state space model and a measurement model of iterative extended Kalman filtering through a star vector under a star sensor body coordinate system, an ephemeris vector under an inertial coordinate system and inertial information, and deducing and calculating to obtain a measurement Jacobian matrix; and substituting the measured Jacobian matrix into the measurement model of the iterative extended Kalman filter to iteratively estimate the satellite inertial installation error, namely the installation error of the inertial astronomical integrated navigation system, so that the function of calibrating the installation error of the inertial astronomical integrated navigation system is realized. The method eliminates the influence of the installation error between the inertial navigation system and the star sensor on the attitude measurement of the star sensor, is suitable for the application occasion of combined navigation, and is good in adaptability and high in precision.

The invention discloses a hotspot target tracking method and device based on load detection and star sensor resetting, and belongs to the field of spacecraft attitude control. The method comprises the following steps: estimating longitude, latitude and elevation information of a target according to a time scale of load target detection information and historical load visual axis ground pointing point information; according to the maximum attitude maneuverability of the system and a five-order polynomial guide model, the fastest moment of target recapture is determined by taking time as an iteration quantity, and the target attitude in the target recapture process is autonomously planned and closed-loop control is carried out; after the target is captured again, the control system performs closed-loop control according to load real-time detection information to ensure that the target is located in the center of a load view field, and the load performs continuous imaging; after tracking imaging is finished, a guiding process of returning to scanning imaging is autonomously planned; in the whole process, the current attitude is automatically reset according to a double-vector attitude determination result of the high-precision star sensor by taking the angular velocity as a constraint, so that the pointing precision is ensured.

The invention belongs to the technical field of space photoelectric detection, and provides a high-precision tracking and pointing system and method capable of autonomously formulating an observation task in orbit. The system comprises an information processing unit, a power supply management unit, a task management unit, a thermal control unit, a GNSS unit, a star sensor unit, a gyroscope unit, a tracking and pointing calculation unit and a servo turntable unit. The system can receive a platform instruction, integrate information of sensors such as a GNSS, a star sensor and a gyroscope, calculate the accurate position of a tracked celestial body, and adjust the direction of a servo turntable in real time so as to accurately track a target. The working method comprises the steps of system starting and initializing, observation task making, tracking and aiming mode preparing, aiming direction correcting, observation target tracking and aiming, exception handling and the like. The system can realize high-precision observation of a celestial body target without depending on attitude adjustment of a spacecraft platform, improves the observation efficiency and reliability, and is suitable for observation of celestial bodies such as pulsar on a space station platform.

The invention provides an optical navigation method and device for an earth-moon spacecraft, electronic equipment and a storage medium. The method comprises the steps that an optical image of the moon is shot through image acquisition equipment, the optical image is analyzed according to imaging parameters of the image acquisition equipment, the angle diameter of the moon is obtained, the target distance of the earth-moon spacecraft relative to the moon is calculated according to the angle diameter of the moon, and the optical image comprises the surface of the moon; measuring starlight angular distance information between the earth-moon spacecraft and the moon by adopting a star sensor, wherein the starlight angular distance information is used for representing the angular distance between the navigation star and the centroid of the moon detected by the star sensor; and calculating the current orbit position and the current attitude of the earth-moon spacecraft by adopting a navigation solution algorithm according to the target distance and the starlight angle distance information.

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 the technical field of avionics and flight control, in particular to a multi-source fusion avionics control system and method, a multi-source acquisition and compression module synchronously acquires inertial measurement, star sensor and radar data under a unified time reference, and generates observation data through random projection compression; the fusion twinning module obtains a fusion state and a hypergraph based on compressed observation data, generates pneumatic, electromagnetic and structural twinning environments in real time on a shared three-dimensional grid through physical constraint diffusion, and outputs a registration result by adopting entropy regularization optimal transmission; the decision execution module maps discrete control variables quantified by a flight control computer into an Isinus model, obtains control sequence seeds through quantum annealing, generates a decision instruction through pulse sinus strategy network optimization, issues the decision instruction to a control execution unit to generate control execution data, and updates a hypergraph in combination with execution feedback for next cycle use. According to the method, the attitude estimation precision and the control robustness under the conditions of high maneuverability and strong interference 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 relates to the field of aerospace thermal control, in particular to a star sensor light shield based on foam carbon and phase change materials, which comprises a foam carbon main body, and two phase change materials with different melting points are filled in pores of the foam carbon main body; the graphene film is co-cured on the outer surface of the foam carbon main body and completely covers the outer surface of the foam carbon main body; the heating sheets are adhered to the outer surface of the graphene film in the circumferential direction of the foam carbon main body to form different heating areas; the thermistors are adhered to the positions, corresponding to the heating areas, of the outer surface of the graphene film; the outer surface of the graphene film is covered with the multi-layer heat insulation assembly; and the temperature controller is used for adjusting the output power of the heating sheet by adopting a self-adaptive temperature control algorithm according to the temperature data acquired by the thermistor and the state of the phase change material, and controlling the temperature of the star sensor light shield within a preset temperature range. The direct solar radiation heat flow on the light shield can be effectively inhibited and dredged, and the high thermal stability control of the light shield is realized.

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 relates to a flight calibration method and system for structural parameters of a star sensor and an inertial unit, and the method comprises the following steps: setting parameters and environmental parameters of the star sensor and the inertial unit, and setting an initial state of a structural parameter deviation estimator; after the inertial unit and the star sensor output data, performing time propagation on estimation values and estimation covariance of constant zero offset, zero offset drift and structural parameter deviation; calculating attitude increments of the inertial unit and the star sensor by using the angular velocity and star observation information, and constructing a measurement equation of structural parameter deviation; solving a corresponding regularization least square problem by using a regularization robust filtering framework, and completing estimation value updating and estimation covariance updating of constant value zero offset, zero offset drift and structure parameter deviation; and after the stable estimated value of the structure parameter deviation is obtained, the structure parameters are compensated. According to the method, high-precision structural parameter calibration of the star sensor and the inertial unit can be realized, and complicated astronomical-ground coordinate conversion can be carried out without the aid of time service positioning information.

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 near-earth space all-day star sensing system based on a sub-wavelength metal grating and a working method thereof, and belongs to the technical field of star sensors. In order to increase the number of fixed stars which can be detected by the near-earth space all-time star-sensitive system, the near-earth space all-time star-sensitive system comprises an optical system, an InGaAs detector, a spectrum optical filter, a micro-polaroid array, a memory, a processor, a control unit and a time unit, the InGaAs detector is respectively connected with the micro-polaroid array, the memory and the control unit; the micro-polaroid array is connected with the optical system, and the optical system is connected with the spectrum optical filter; the memory is respectively connected with the processor and the time unit; and the processor is connected with the control unit. According to the invention, the polarization filtering work of the sky background is realized, and the number of observable fixed stars is effectively increased on the basis of ensuring the miniaturization of a near-earth space all-day star-sensitive system.

The invention discloses a satellite attitude control method and system when gyro data cannot be used and a satellite, and relates to the technical field of satellite attitude control, the method comprises the following steps: when it is determined that the gyro data of the satellite cannot be used, judging whether star sensing data collected by a star sensor configured by the satellite in a current control period is valid, if yes, executing the step 2, and if not, executing the step 3; determining the attitude data of the satellite at the current moment according to the star sensor data collected by the star sensor in the current control period; and when a preset condition is met, performing attitude control on the satellite according to the attitude data of the satellite at the current moment. In the invention, when the gyro data of the satellite cannot be used, the attitude data of the satellite at the current moment can be determined by using the star sensor data in the current control period acquired by the star sensor configured on the satellite, so that the attitude control can be effectively and timely performed on the satellite, the task implementation can be guaranteed to the greatest extent, and the satellite attitude control accuracy is improved. The engineering implementation operability is high.

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 provides a geometric recalibration method and device based on a star group time sequence image and electronic equipment, and relates to the technical field of satellite data processing.The geometric recalibration method based on the star group time sequence image comprises the steps that the long-period included angle variable quantity between any two star sensors is counted according to time sequence observation vectors of the star sensors, and the time sequence observation vectors of the star sensors are calculated according to the long-period included angle variable quantity; decomposing the variable quantity of the included angle into a three-axis attitude change angle; performing aberration compensation according to the original data of the star sensor, and determining a three-axis attitude correction value; determining a three-axis optimized attitude angle based on the three-axis attitude change angle and the three-axis attitude correction value of the long-period change of the included angle; and a re-calibration model is determined according to the three-axis optimization attitude angle, and the re-calibration model is solved to obtain an optimal calibration model for image calibration, so that the technical problem of space reference change of the satellite group long-time sequence image in the prior art is solved.

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 provides a two-dimensional expansion type SAR satellite in the field of SAR satellites. The two-dimensional expansion type SAR satellite comprises a satellite body, an SAR antenna, a solar array, a phased array data transmission antenna, a star sensor, an antenna measurement and control antenna, a ground wire measurement and control antenna, an electric thruster and other key components. A unique two-dimensional unfolding mechanism is adopted in the satellite design, so that the satellite can be loaded with a larger SAR (Synthetic Aperture Radar) antenna; the solar array is folded and pressed on the side face of the satellite body, the phased array data transmission antenna and the ground wire measurement and control antenna are installed on the side face (to the ground) of the satellite body, and the star sensor is installed on the top of the satellite body through a carbon fiber support. Space envelope resources in the fairing are fully utilized, the satellite is small in size in the folded state, the satellite structure is stable and reliable, and the fairing can adapt to various observation tasks.

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.

A star sensor installation device comprises an installation support, a heat dissipation plate and a temperature control system, the installation support comprises at least one installation face, the installation face is provided with a star sensor installation interface, a star sensor is installed on the installation support through the star sensor installation interface, the heat dissipation plate is flexibly connected to the installation face, and the temperature control system is connected to the heat dissipation plate. The temperature control system is arranged on the mounting bracket and is used for carrying out temperature control on the mounting bracket, the star sensor comprises a conduction part, and the conduction part is in contact with the heat dissipation plate, so that heat energy generated in the working process of the star sensor is diffused to the outside through the heat dissipation plate. A satellite is also provided.

The invention relates to the technical field of satellite orbit determination, in particular to a low-orbit giant constellation autonomous orbit determination method based on a star sensor and inter-satellite angle measurement. The method comprises the following steps: step 1, determining a constellation configuration by using inter-satellite quantity measurement: expressing each satellite state quantity in a constellation by using six elements, constructing a partial derivative matrix of inter-satellite quantity measurement relative to state information, constructing a linear equation set according to the state quantity, the quantity measurement and the partial derivative matrix, and determining a constellation configuration; and solving the satellite state quantity describing the constellation configuration through an improved least square method. Step 2, correcting constellation drift by using a star sensor and a camera quantity measurement and combining the constellation configuration in the step 1: acquiring a true value of an included angle between an adjacent satellite direction vector and a certain fixed star direction vector by using the star sensor and the camera, then calculating an estimated value of the included angle between the adjacent satellite direction vector and the certain fixed star direction vector, and correcting the constellation drift; and the difference between the real value and the estimated value of the included angle is minimized to realize constellation correction. By adopting the scheme, orbit determination can be realized without relying on a ground observation station, and data is easy to manage.