15 results about "Double star" patented technology

An inter-satellite pointing target attitude high-precision real-time online prediction method, first, the ground uses the orbit determination program corrected by on-orbit data observation to generate a double-satellite orbit position and velocity value sequence in a predetermined time period, and samples according to the control period, then for each sampling time, the double-satellite orbit position and velocity is used to establish the local orbit coordinate system, and the component of the target star position vector in the local orbit coordinate system is calculated, the pitch and yaw target attitude angle values are calculated according to the target attitude angle definition, the pitch target angle sequence is linearly averaged and estimated, the average orbit period value and the long-term change rate parameter are calculated, and the modified sequence after deducting the long-term change rate is given, for the pitch target angle correction sequence and the yaw target angle sequence, data of a specified average orbit period number is taken from a specified initial time, the fitting coefficient is calculated, finally the generated fitting coefficient is uploaded to the satellite, and the satellite uses the high-precision fitting determination formula to calculate the inter-satellite pointing target attitude angle online every control period.

The invention discloses a navigation mode transition method, an electronic device and a medium, in a working process of a first navigation mode, geometric angles among a mother satellite, a service spacecraft and space debris are monitored, and when the geometric angles meet a mode switching requirement, the position of a target orbit under a J2000 coordinate system is obtained through double-satellite geometric estimation; and taking the position as an initial position calculated in the second navigation mode, and switching to the second navigation mode. Before the navigation mode is switched, the target orbit position is determined through double-satellite geometric estimation, the single-satellite navigation positioning precision can be rapidly improved, a good initial value is provided for double-satellite angle measurement filtering, and the convergence process is accelerated.

The invention discloses an astronomical experimental device for simulating the photometric change of double food stars, and the device comprises a fixing unit which comprises a supporting part located at the lower part, an angle display part disposed at the top end of the supporting part, and a length display part transversely disposed on the angle display part. The limiting component is arranged on the angle display component, and the angle display component and the length display component are connected in series through the limiting component; the limiting unit comprises a fixing part located on the length display part and an adjusting part arranged below the fixing part. The invention provides a device capable of meeting more use requirements, improving the use efficiency of the device, facilitating the smooth proceeding of experimental work, and providing a device capable of recording the luminosity change and simulating the luminosity change conditions of different stages of a food double-star experiment by controlling the movement and the shielding degree of a star, so that the experimental efficiency of the food double-star experiment is improved. And thus, a scientific research scheme in the aspect of eating double stars in astronomy can be better assisted for working personnel.

This invention discloses a near-space target localization method based on dual-satellite imaging, comprising: simultaneously acquiring original observation images of the same target using two low-Earth orbit (LEO) satellites; acquiring the centroid coordinates of the target region in the two original observation images respectively, and performing coarse localization based on the centroid coordinates to obtain the coarse localization point coordinates of the near-space target; performing forward path correction to generate a simulated image of the target observed by the two LEO satellites under the current localization; performing residual correction on the estimated spatial position of the target to obtain the residual-corrected spatial position of the target; and repeating the above steps based on the residual-corrected spatial position in the current iteration until the iteration conditions are met to obtain the precise localization of the target. By constructing a closed-loop technical scheme of "dual-satellite synchronous imaging - coarse localization - non-uniform atmospheric and imaging system effect coupling correction - image centroid comparison - iterative optimization", the accurate inversion of the three-dimensional position of near-space targets can be achieved.

The application provides a method and device for optimal resource allocation decision of star group cooperative observation of space targets. The method comprises: obtaining input parameters, state parameters, constraint parameters and optimization variable parameters; based on the input parameters, the state parameters and the constraint parameters, calculating a benefit coefficient of observation of each target by each satellite, and updating each benefit coefficient based on the optimization variable parameters; calculating a cost coefficient matrix of double-star cooperative observation of targets according to the updated benefit coefficients, and constructing an optimal resource allocation decision model based on the cost coefficient matrix, so as to allocate satellite resources by using the optimal resource allocation decision model. According to the scheme, real-time and rapid optimal solving on the star can be facilitated, so that optimal resource allocation is realized.

This invention relates to a clustering method for observation targets based on continuous observation constraints of two satellites, comprising the following steps: at each moment, determining whether the observation target is physically potentially visible relative to the satellite using off-axis angle constraints and detection distance constraints; when the observation target is physically potentially visible relative to the satellite, using direction vector similarity analysis to extract the spatial nearest neighbor targets of the observation target, constructing a set of satellite line-of-sight candidate directions pointing towards the geometric center of the observation target group; selecting the optimal satellite line-of-sight candidate direction that covers the most observation targets from the set of satellite line-of-sight candidate directions; and setting the sliding window length to seconds, selecting the optimal set of observation targets at each moment within the sliding window, and taking the intersection of all optimal sets of observation targets to select the intersection target group that is always located in the optimal field of view corresponding to the optimal satellite line-of-sight candidate direction within the sliding window as the continuous observation target set.

The application discloses a space-based space target trajectory segmentation fitting method based on infrared radiation characteristics, and comprises the following steps: (1) acquiring initial positioning information of a space target in a ground-fixed coordinate system through double-star joint observation; (2) monitoring the infrared radiation intensity of the space target in real time through an infrared sensor, identifying the classification points of the space target in different flight stages according to the infrared radiation intensity and signal-to-noise ratio change, and completing target trajectory segmentation; (3) for different flight stages, the least square method is introduced through an adaptive factor, the trajectory optimization fitting of the space target in the space-based scene is completed, high-precision positioning and tracking are realized, and the motion characteristics of the target are acquired. By using the application, the overfitting or underfitting problems in the traditional method can be reduced, so that the precision and reliability of space-based space target detection and tracking are improved, and the application is suitable for a complex space target observation environment.

The application discloses a kind of navigation star library construction methods of approximate uniform distribution under small field of view condition, comprising: (1) using VMT algorithm, set threshold to SAO star table is segmented, only keep following navigation star, store in star library StarLib;(2) in star library StarLib, eliminate the navigation star of brightness and position not fixed, star distance value is too small, that is, eliminate variable star and binary star, obtain basic star library (StarLib1);(3) using approximate uniform distribution of spherical spiral reference point, according to the angle distance between reference point and star, star, the combination characteristics of three star density to rough navigation star, obtain preliminary navigation star library (StarLib2) and alternative star library (StarLib3);(4) on the basis of StarLib2, using KernelSVM-SMO from nonlinear distribution StarLib3 star library, select navigation star, calculate the average value of angle distance between selected star and existing navigation star in current field of view, judge whether less than set threshold, to effectively avoid redundant star, obtain final navigation star library (StarLib4).

The invention discloses a satellite-borne laser terminal rotary table pointing zero offset calibration method using double star sensors. Relates to the technical field of spacecraft attitude measurement and pointing control crossing, in particular to a satellite-borne laser terminal rotary table pointing zero offset calibration method using double star sensors. According to the method, a complete attitude transfer chain model is established, inertial attitude data output by a platform star sensor and inertial attitude data output by a terminal star sensor are fused, so that the measurement relative attitude of the terminal star sensor relative to a satellite platform is constructed, and the measurement attitude and a pre-established theoretical attitude chain model are subjected to high-precision alignment; a nonlinear estimation model about the azimuth coding angle zero offset and the pitching coding angle zero offset of the rotary table is constructed, the azimuth and pitching zero offset values of the rotary table can be autonomously estimated in an on-orbit mode with high precision by solving the problem, and finally on-orbit autonomous calibration of the pointing zero offset of the rotary table is achieved. The whole process does not need ground intervention, and the autonomy and long-term stability of satellite pointing control are effectively improved.

The invention relates to the technical field of spacecraft attitude determination, and particularly provides a double-star sensor attitude determination method and an attitude determination precision verification method and system.The method comprises the steps that absolute time and trigger time are generated based on a time synchronization signal; acquiring current output data of the two star sensors; comparing the timestamps with the triggering moments, and processing quaternions corresponding to the triggering moments according to different comparison results to obtain time-synchronized target quaternions; calibrating through a theodolite to obtain a first mounting matrix and a second mounting matrix of the first star sensor and the second star sensor to the same reference; and fusing the time-synchronized target quaternion based on the first installation matrix and the second installation matrix to obtain the attitude quaternion after fusion of the double star sensors. According to the double-star-sensor attitude determination method and the attitude determination precision verification method and system, time alignment of cross-equipment data is realized, and independent, repeatable and high-confidence evaluation of measurement precision of the star sensors and fusion algorithms thereof is realized at the same time.

1. The name of the design product: firework base (double star). 2. The use of the design product: for firework base. 3. The design points of the design product: in shape. 4. The picture or photo that best indicates the design points: perspective view 1.

The invention discloses a double-star positioning and error analysis inhibition method for a small-view-field star sensor, which comprises the following steps of: firstly, reading measurement data of the small-view-field star sensor, selecting two navigation stars, calculating star vectors of the navigation stars, aligning data of the two navigation stars at different moments to the same moment, resolving a double-star positioning result, converting the double-star positioning result into spherical coordinates, and carrying out error analysis on the spherical coordinates. The unique position of the carrier is determined according to data at the previous moment, and finally effective positioning results are screened through criteria. According to the method, aiming at the characteristics of the small-view-field star sensor, the invalid positioning result caused by error amplification is removed by processing observation navigation star data and fully considering an error amplification mechanism of double-star positioning, so that the confidence coefficient of the double-star positioning result of the small-view-field star sensor is improved.

The application relates to a kind of double-star electromagnetic formation dynamics modeling and control method of port Hamilton system, comprising: defining two slave spacecraft center-of-mass position coordinates in ECI coordinate system and LVLH coordinate system;According to the center-of-mass position coordinates of slave spacecraft, the kinetic energy and potential energy of slave spacecraft in ECI coordinate system are obtained;According to the center-of-mass position coordinates of slave spacecraft and the coordinate transformation matrix, the Lagrangian function L in LVLH coordinate system is obtained;The Euler-Lagrange equation is established, according to Legendre transformation, the generalized momentum coordinates p j are selected, the Hamilton function H is obtained, and the port Hamilton dynamics model is established;An electromagnetic formation controller is designed according to the trajectory tracking passive control method based on interconnection-damping distribution;The electromagnetic satellite formation trajectory tracking magnetic dipole strength control law is obtained according to the electromagnetic formation controller.The application retains the nonlinear term of the system, embodies the inherent dissipation characteristics of the system, can avoid the problem of reducing the accuracy of the system caused by linearization, and ensures the passivity of the system.

The invention discloses a method and a system for identifying cyanobacterial bloom risk based on double-satellite cooperation. The method comprises the following steps: S1, constructing an A-satellite lightweight chromaticity boundary triggering model and a B-satellite refined spectrum risk classification model based on historical data; s2, preprocessing and identifying a target water body area by using an A satellite receiving image, and converting the multispectral reflectivity value of each water body pixel into a CIE chromaticity space to obtain a chromaticity coordinate; judging whether the chromaticity coordinates exceed a preset boundary or not, and if the chromaticity coordinates meet a triggering condition, scheduling a satellite B through an inter-satellite communication link; s3, the satellite B responds to the scheduling instruction, analyzes a task triggering report, obtains an image extraction pixel multi-dimensional feature system, calculates risk scores by using a spectrum risk classification model, grades the risk scores, and outputs an identification result; according to the method, the A star uses the chromaticity boundary triggering model to screen and dispatch the B star, and the B star uses the spectrum risk classification model to perform refined evaluation, so that double-star cooperative identification of the cyanobacterial bloom risk is realized.

The application discloses a kind of on-orbit calibration methods of follow sight measurement error based on GNSS, relative position relationship between two stars is calculated by using double star GNSS differential data, and this is used as the reference value of follow sight measurement calibration, the measurement error of spaceborne follow sight radar is obtained by calculation, and the corresponding relationship between the measurement error of spaceborne follow sight radar and the position of the sun is obtained by further fitting, for ground number correction compensation.The application reduces the influence of follow sight radar measurement system error, improves the precision of follow sight radar on-orbit measurement, and can meet the high-precision relative measurement requirements of close-range double-star companion flying task.