71 results about "Orbit+Face" patented technology

The invention discloses a fuel-saving control on-satellite autonomous planning method for rapid accompanying flying, fly-around flying and sweeping flying of an elliptical orbit coplanar target, and relates to a fuel-saving control on-satellite autonomous solving method for rapid approaching and high-precision accompanying flying / fly-around flying / sweeping flying from a long-distance position on the same orbit surface. Comprising the following steps: providing a three-time planning and five-time maneuvering long-distance approaching accompanying flight / fly-around / grazing flight configuration forming method; the efficient iteration method is suitable for solving the initial value of a minimum fuel control analytical solution and the final value of an iterative precise solution of a high-precision orbit extrapolation model through a satellite autonomous solution two-body model. The method is suitable for a near-earth space random elliptical orbit satellite to quickly approach a long-distance target on an approximate same orbit to carry out accompanying flight / fly-around / grazing flight, and the control precision can reach a ten-meter level; and the calculation amount is small, and the method is suitable for on-satellite autonomous task planning and solving.

The invention discloses a satellite constellation autonomous deployment method and system, belongs to the technical field of aerospace orbit control, and solves the problems of long deployment period, high personnel investment, slow collision avoidance response, increased orbit safety risk and the like caused by a traditional satellite deployment method. The method comprises the following steps: constructing a constellation network according to initial orbit parameters of a satellite and orbit injection time and orbit parameters of different stages; outputting an optimal deployment strategy of each stage of the constellation through a genetic algorithm; the right ascension of the ascending node and the phase are adjusted to a target position together while the height is deployed in place through a coupling equation, and ignition constraints are constructed; generating an orbital transfer plan according to the thrust and a satellite initial orbit needing to be deployed; an orbital transfer instruction is automatically generated according to the orbital transfer plan; according to on-orbit telemetering, the orbital transfer progress condition is judged, and strategy closed-loop adjustment is carried out; cooperative deployment of multi-orbit-plane satellites is achieved, and fuel and time cost are saved. The method is particularly suitable for multi-orbit-plane collaborative deployment and autonomous management of large-scale satellite constellations.

The invention relates to the technical field of constellation satellite maintenance regulation and control, and discloses an intelligent regulation and control system and method for low-orbit Walker constellation configuration maintenance, and the method comprises the steps: obtaining the change trend of an in-plane phase angle and an inter-plane distance of an orbit of a satellite in real time through a high-precision orbit prediction model; inputting the in-plane phase angle and the inter-plane spacing into a pre-trained decision model, judging whether the parameters exceed a preset configuration maintenance threshold, and positioning an unstable satellite; if the parameters exceed the limit, the model outputs the optimal orbit maneuvering parameters of the satellite based on dual-target constraints of minimum fuel consumption and shortest service interruption; a phase maintenance demand and an on-satellite observation task are fused, an optimal control scheme is generated, and constellation configuration is maintained in combination with task planning and implementation; continuously monitoring the satellite constellation configuration state, and maintaining the stable operation of the constellation. According to the invention, automatic maintenance of Walker constellation configuration is realized, and the control precision and efficiency are improved.

A method for node sensitivity analysis affecting the degradation of the communication performance of a giant constellation. The present invention relates to the field of spacecraft system simulation. It solves the problem that in the prior art, there is a lack of a method for measuring the sensitivity of the constellation to changes in the number of satellites, which brings changes in performance such as communication delay. The method includes: reading the parameter values under different two-line element data segments according to the format of the two-line element data record and saving them to the SGP4 orbit model structure variable; dividing different sub-constellations within the giant constellation according to the orbital inclination or the semi-major axis of the orbit; setting that the satellites within each orbital plane form a routing domain, setting the gateway on / off condition, setting that the satellite nodes can communicate within the domain, and restricting the connectivity of the nodes in different domains; updating the connectivity between the nodes, outputting the on / off situation and delay between the nodes; and performing routing and outputting the routing result to complete the node sensitivity analysis. It is also applicable to the field of measuring the communication delay of giant constellations.

The invention discloses a satellite deployment method, system and device and a storage medium, and relates to the technical field of satellite deployment. The method comprises the following steps: for a backup satellite adjusted on the same orbital plane, obtaining a relative phase angle between the backup satellite and a failure satellite after the backup satellite and the failure satellite run after the transfer time based on the transfer time required by the backup satellite to climb to a target orbit from a drift orbit, and obtaining the drift time based on the relative phase angle; for the backup satellite adjusted across the orbital plane, the running time of the backup satellite is obtained, the right ascension of the relative ascending node between the backup satellite and the failed satellite after the backup satellite and the failed satellite run after the running time is obtained based on the running time, and the adjusting time is obtained based on the right ascension of the relative ascending node; and setting the ignition time of the backup satellite based on the drift time and the adjustment time. The method has the beneficial effects of low energy consumption and short consumed time.

The invention discloses a satellite network laser inter-satellite link establishment method and system, electronic equipment and a storage medium, and mainly solves the problem of low availability of a satellite network established in the prior art due to sun-sun influence. According to the scheme, the method comprises the following steps: inputting constellation configuration parameters of a satellite network, and calculating satellites which are uniformly distributed on each orbit surface; according to the input constellation configuration parameters, establishing a same-orbit inter-satellite laser link between all satellites in the constellation and adjacent satellites in the same orbit plane where the satellites are located; according to the constellation configuration parameters, calculating boundary phase factors used for judging different-orbit link connection modes; and according to the phase factor of the constellation and the boundary phase factor, establishing a different-orbit laser inter-satellite link between satellites on all adjacent orbits, and completing the construction of the whole-network inter-satellite laser link. According to the method, the optimal link connection configuration can be screened out in the satellite network construction stage, the availability of the satellite network is improved, and the method can be used for inter-satellite link sun-sun avoidance.

The application aims to provide a remote sensing satellite descending node local time control method based on orbit height adjustment, which has simple control and high control precision. The method comprises the following steps: S1, acquiring orbit elements of a remote sensing satellite, wherein the orbit elements are instantaneous elements or two-line elements; S2, calculating a current descending node local time of the satellite based on the orbit elements; S3, calculating a drift deviation of the current descending node local time from a nominal descending node local time △T ; S4, calculating a current orbit plane rotation rate based on a J2 perturbation parameter of the earth; S5, calculating a difference between the current orbit plane rotation rate and an earth revolution rate; S6, designing an expected orbit plane rotation rate based on the difference calculated in the step S5; S7, calculating an expected orbit height according to the expected orbit plane rotation rate; and S8, adjusting the orbit height of the satellite to the expected orbit height. The application is applied to the technical field of remote sensing satellite measurement and control.

The application relates to a rotation error correction method of a distributed autonomous navigation constellation, and comprises the following steps: S1, each satellite calculates the orbit plane orientation error of the satellite in orbit and distributes the result to all visible satellites in the constellation in real time; S2, each satellite fits a local constellation rotation quantity according to the orbit plane orientation error of the satellite and the collected orbit plane orientation error; S3, each satellite determines a virtual satellite orbit observation quantity according to the local constellation rotation quantity of the satellite; S4, each satellite constructs a conditional constraint equation of virtual observation according to the virtual satellite orbit observation quantity and the second type of non-singularity orbit root number; and S5, each satellite solves the satellite orbit through filtering according to the conditional constraint equation of virtual observation, so that the overall rotation effect of the constellation is inhibited and corrected. The application can inhibit the rotation effect of the constellation, improve the distributed autonomous orbit determination precision of the satellite, does not increase the inter-satellite data interaction communication pressure, and can stably maintain the constellation configuration for a long time.

The application discloses an initialization method of a conic surface under a projection linearization track surface, and belongs to the technical field of calculation, reckoning or counting. The method first obtains a preliminary estimation of a target track through three groups of observation data; determines a conic surface containing an initial track surface according to initial track information and geometric knowledge; estimates the missing height information in combination with the conic surface, and obtains more position information of the target in combination with the existing angle information; projects the position information of the target into a two-dimensional section, re-estimates the track parameters by using a least square algorithm, and more accurate track parameters can be obtained; then the iteration process is repeated by using the newly obtained track information until a certain initial orbit determination accuracy requirement is met. The algorithm firstly fits the height information of the target by using a conic surface coordinate system, and realizes linear fitting of the position vector by using a track surface projection method, so that the calculation amount is small and the accuracy is high.

The application discloses a heterogeneous satellite constellation configuration correction method based on stable earth coverage performance. The method comprises the following steps: constructing a first query table and a second query table; for each orbit plane, using a working orbit parameter optimization method, combining the first query table and the second query table, analyzing and calculating dynamic visibility conditions and a visibility list, and then correcting the working semi-major axis and the latitude argument AoL of each satellite on the orbit plane; for each orbit plane, according to the working semi-major axis correction value of the satellite on the orbit plane, combining the first query table, calculating the right ascension of the ascending node RAAN correction value of the satellite; and the working semi-major axis correction value, the latitude argument AoL correction value and the right ascension of the ascending node RAAN correction value of each satellite on all orbit planes constitute the configuration parameters of the heterogeneous satellite constellation. The application restrains the drift of the earth coverage index of the heterogeneous satellite constellation, reduces the constellation control frequency, and avoids the additional increase of the number of satellites due to the drift of the coverage index and the multiple artificial iterative design.

The invention provides a low-orbit constellation deployment method and device based on an evolution strategy, and belongs to the technical field of universe navigation. The method comprises the following steps: designing the configuration of a target low-orbit constellation, screening from the target low-orbit constellation to obtain a minimum sub-constellation, and determining an evolution strategy of deploying an orbit surface in the minimum sub-constellation first and then deploying the rest orbit surfaces in the target low-orbit constellation, thereby ensuring that the target low-orbit constellation can be deployed in the deployment process of the target low-orbit constellation. The network capacity of the target low-orbit constellation is gradually increased, so that the target low-orbit constellation is matched with the user demand, the deployment cost is reduced, the overall service capability in the constellation deployment period is improved, and the service efficiency of the target low-orbit constellation is higher.

The invention discloses a fuel-saving control on-satellite autonomous planning method for rapidly transferring an elliptic orbit different-plane target to the same orbit, and particularly relates to a fuel-saving control on-satellite autonomous solving method for rapidly transferring the elliptic orbit different-plane target to the same orbit from a different orbit plane. Comprising the following steps: judging orbital plane normal correction or orbital shape correction according to orbital elements of an initial satellite and a target; solving a fuel-saving orbital plane normal correction control strategy according to orbital plane normal orbital element difference of the satellite and the target; solving a fuel-saving orbit shape correction control strategy according to the in-orbit element difference of the orbit planes of the satellite and the target; and a high-precision orbit extrapolation model is adopted to carry out iterative correction on the fuel-saving control strategy. The method is suitable for orbital transfer of any elliptical orbit target by any near-earth space elliptical orbit satellite, and the control precision can reach ten meters; and the calculation amount is small, and the method is suitable for on-satellite autonomous task planning and solving.

The invention discloses a two-way rocket launching window searching method, system and device and a medium, and the method comprises the steps: rapidly calculating two nearest launching windows of an ascending orbit and a descending orbit of a rocket according to the attributes of a central body, target orbit plane parameters, the geographic coordinates of launching points and the flight duration of the rocket; according to the method, the problem is converted into numerical calculation based on space geometry, the crossing moment is directly solved by using the geometrical relationship between the normal vector of the orbital plane and the inertial position vector of the launching point location and adopting efficient numerical methods such as dichotomy and the like, and complex dynamic iteration is avoided. According to the method, the calculated amount is greatly reduced, rapid searching of the launching window is achieved, the method is suitable for an interstellar detection task scene needing rapid response or multiple times of calculation, and the real-time performance of task planning is remarkably improved.

The application discloses an antenna switching method for inter-satellite networking. When the included angle between the direction vector of any antenna in a closed state and the direction vector of the opposite satellite is less than or equal to a preset value within a continuous specified time length, the antenna is opened, and the currently opened antenna is closed.

The invention provides a navigation constellation autonomous orbit determination method based on space-based observation, which comprises the following steps: establishing inter-satellite routing, and permanently establishing a link based on phases between satellites in the same orbit plane; a laser inter-satellite link observation value is obtained, and a vector between every two satellites in the same orbital plane is obtained based on laser inter-satellite link distance measurement and angle measurement observation, that is, the laser inter-satellite link observation value represents vectors from the satellite i to the satellite j; coordinate system conversion: converting the laser inter-satellite link observation value from a satellite body coordinate system to an inertial system; establishing an observation equation; establishing a motion equation; and generating a normal equation, combining the observation equation and the motion equation to generate the normal equation, and performing inversion to obtain an orbit determination result. According to the method, the feasibility of realizing autonomous orbit determination of the navigation constellation through laser inter-satellite link and star-sensitive observation is explained. The direction precision of the laser inter-satellite link is equivalent to or even higher than that of a fixed star sensor, and a high-precision and long-term stable orientation reference can be provided for a normal equation, so that the autonomous orbit determination precision and reliability are improved.

The invention provides a satellite switching method, a satellite switching device and electronic equipment, which can be applied to the technical field of satellite communication. Based on the method, when the user equipment needs to be switched from a source satellite to a target satellite, a first angle difference between the user equipment and the source satellite based on a first orbital plane and a second angle difference between the user equipment and the target satellite based on a second orbital plane can be firstly obtained; wherein the first orbital plane is an orbital plane where a source satellite is located, and the second orbital plane is an orbital plane where a target satellite is located; determining target switching time according to the first angle difference and the second angle difference; correspondingly, the user equipment switches from the source satellite to the target satellite based on the target switching time. By calculating and utilizing the angle difference based on the orbital plane, the related data processing process can be effectively simplified, and the corresponding switching time can be quickly and accurately determined, so that the satellite switching of the user equipment can be accurately and timely realized.

The present invention discloses a method for analyzing the optimal pursuit start azimuth based on the condition of pulsed maneuver in orbital pursuit and evasion, including: setting the pursuit and evasion to be carried out on the same orbital plane, establishing an orbital coordinate system with the initial position of the evading spacecraft as the origin based on the orbital altitude and the initial relative distance between the evading spacecraft and the pursuing spacecraft, and determining the initial states of the evading spacecraft and the pursuing spacecraft; based on the initial states of the evading spacecraft and the pursuing spacecraft, using a pulsed pursuit and evasion game strategy solution method based on game tree search to obtain the relative distance under the optimal strategy; the pursuing spacecraft starts to pursue from different start azimuths respectively, compares the relative distances at all azimuths under the optimal strategy, and obtains the optimal pursuit start azimuth under this initial condition. The present invention takes into account the limited action set of the pursuit and evasion spacecraft and the influence of the natural evolution of the spacecraft, obtains the optimal pursuit start azimuth under the same orbital plane, and is helpful for the subsequent research on the selection of the initial position in orbital pursuit and evasion.

The application provides a large-scale low-orbit satellite distributed authentication and key management method, which comprises the following steps: a same-orbit-plane satellite group leader generates a batch message authentication code, and sends a first authentication request to a ground control center; the ground control center calculates an expected batch message authentication code based on the first authentication request, compares the batch message authentication code, and derives a session key after the comparison is consistent, makes the satellite locally derive the same session key through a broadcast response, and takes the candidate authentication key and the candidate anonymous identity derived in the current round of authentication as the input of the next round of satellite-ground authentication; the same-orbit-plane satellite derives an inter-satellite shared bootstrapping authentication key by using an orbit public physical parameter, a second authentication request containing an anonymous identity and a location proof is generated by a satellite based on the inter-satellite shared bootstrapping authentication key, and a session seed is calculated by a receiving satellite based on the second authentication request, and the session seed is taken as an authentication key for the next inter-satellite authentication. The large-scale low-orbit satellite distributed authentication and key management is realized.

The application discloses a kind of unloading method and device of asymmetric configuration satellite angular momentum, the method includes: by solar sensor, the solar vector under satellite body system is acquired, and the solar vector is converted into the solar vector under orbit system;The included angle β of the solar vector under the orbit system and orbit plane is calculated, and the rotation angle of solar wing rotation axis and the preset angle of biasing axis are set according to the included angle β of the solar vector under the orbit system and orbit plane;The angular momentum of flywheel system is calculated and is converted into the angular momentum under satellite body system;According to the modulus of satellite to be unloaded angular momentum and preset unloading threshold value are compared, according to the comparison result, it is judged whether magnetic moment device unloading needs to be started, and according to the comparison result, the control instruction of magnetic moment device is output.The application effectively reduces angular momentum accumulation and space moment accumulation;With the advantages of saving fuel, low power consumption requirement and high attitude stability, it provides a theoretical basis for angular momentum unloading of weak magnetic field planet explorer.

An embodiment of the present application discloses an ephemeris prediction method and device, which relate to the field of satellite navigation technology and are used to reduce the amount of ephemeris parameter broadcast data. The method may include: a server obtains EOP data and historical ephemeris data for a preset time period; performs satellite orbit prediction and satellite clock error prediction based on the EOP data and historical ephemeris data to obtain a predicted orbit and predicted clock error for a future preset time period; after fitting the predicted orbit into orbital parameters, encodes the orbital parameters of satellites belonging to the same orbital plane to obtain an orbital parameter code for each orbital plane; after fitting the predicted clock error into clock error parameters, encodes the clock error parameters to obtain a clock error parameter code; and sends the ephemeris parameter code to a terminal device, where the ephemeris parameter code includes a clock error parameter code and an orbital parameter code for each orbital plane.

The application discloses a method for calibrating a roll angle of an emergency anti-solar time of a satellite on the earth, which is beneficial to reestablishing a three-axis attitude of the satellite on the earth. Firstly, an angle between a projection of a sun vector on an orbit plane and an orbit system-X axis is calculated; then, whether a current orbit position meets a requirement is determined according to the angle between the projection of the sun vector on the orbit plane and the orbit system-X axis, and whether an absolute attitude reference is available at present is determined, if the current orbit position meets the requirement and the absolute attitude reference is available at present, the roll angle is calibrated at a current control beat; otherwise, the roll angle is not calibrated at the current control beat; if the roll angle is calibrated at the current control beat, the roll angle is calculated based on the absolute attitude reference; if the roll angle is not calibrated at the current control beat, the roll angle is calculated by using gyro inertia angular velocity integration. The application can provide high-precision and reliable attitude information for reestablishing the attitude of the satellite on the earth.

The invention discloses an automatic division method and system for a satellite constellation orbital plane, and relates to the technical field of spaceflight measurement and control. The method comprises the following steps: firstly, calculating an orbital plane unit normal vector based on a satellite orbit inclination angle and an ascending node right ascension, and performing first-level spatial pointing clustering by adopting a density clustering algorithm to form an orbital plane group; secondly, in the same orbital plane group, by taking a semi-major axis and an eccentricity rate as characteristics, automatically determining an optimal shell layer number by utilizing a Gaussian mixture model and a Bayesian information criterion, and realizing second-stage orbital energy shell layer division; and finally, in the same shell layer, sorting and interval statistical analysis are performed on the right ascension of the ascending node, recursive subdivision is performed through dynamic threshold segmentation in combination with a range-based dual-condition verification mechanism, and fine division of a third-level orbital plane is completed. The method realizes full-process automation from orbit data to division results, has the advantages of high division precision, strong adaptability, good engineering practicability and the like, and is suitable for operation and maintenance management of large-scale heterogeneous satellite constellations.

The application discloses a method for quickly changing orbit of an elliptical orbit out-of-plane target to the same orbit with the least fuel control on a satellite, and a method for independently solving the least fuel control on a satellite for quickly changing orbit from an out-of-plane orbit to the same orbit, which comprises the following steps: judging whether to perform normal orbit plane correction or orbit shape correction according to orbit elements of an initial satellite and a target; solving a least fuel normal orbit plane correction control strategy according to a normal orbit plane element difference between the satellite and the target; solving a least fuel orbit shape correction control strategy according to an in-orbit plane element difference between the satellite and the target; and iteratively correcting the least fuel control strategy by using a high-precision orbit extrapolation model. The method is suitable for orbit changing of any elliptical orbit satellite in near-earth space to any elliptical orbit target, and the control precision can reach ten meters; and the calculation amount is small, and the method is suitable for solving a task planning on a satellite.

This invention discloses a method for initial orbit determination of a non-cooperative target under space-based angle-only extremely short arc observation, belonging to the technical field of initial orbit determination. This invention addresses the failure of existing initial orbit determination methods under extremely short arc conditions. The method includes: using the unit normal vector of the target orbital plane as the variable to be optimized; obtaining the relationship between the slant distance and the unit normal vector based on the geometric relationship between the target position vector, the slant distance between the space-based platform and the target position, the target azimuth unit vector, and the space-based platform position vector; transforming the solution of the unit normal vector into the solution of the X-axis and Y-axis components of the unit normal vector based on the normalization characteristics of the unit normal vector, and determining the constraints of the solution; and employing a heuristic algorithm to solve for the X-axis and Y-axis components of the unit normal vector, thereby determining the unit normal vector and achieving initial orbit determination. This invention improves the accuracy and robustness of initial orbit determination.

The invention provides a sun-synchronous orbit inclination angle adjusting method and device, and the method comprises the steps: determining a position relation between an initial orbit and a target orbit, enabling two intersection points of the initial orbit and the target orbit to be located above an equator, and taking one of the intersection points as a satellite orbit transfer point; determining an orbital transfer angle required for adjusting the satellite from the initial orbital plane to the target orbital plane; according to the system state of the satellite and the task target, the single ignition duration of the satellite and the ignition moment and shutdown moment of the satellite are comprehensively determined; determining a satellite ignition direction; the speed increment of satellite monopulse ignition is analyzed and calculated, and the total impulse required for changing the inclination angle of the satellite orbit is determined according to the speed increment; the total satellite ignition orbit number is analyzed and calculated according to the total impulse and the actual impulse generated by each time of ignition; and based on the satellite orbital transfer point, the satellite ignition direction, the satellite ignition time and the shutdown time, the satellite is controlled to carry out secondary ignition orbital transfer, and then inclination angle adjustment is completed.

The invention discloses a three-pulse near-moon braking fixed-time and fixed-point landing aiming method. The method comprises the steps that input parameters are obtained; determining a power descending point moment initial value, a second-time near-moon braking moment, a four-device separation point moment initial value, a third-time near-moon braking moment and a lunar surface take-off moment; a finite thrust model and an accurate forecasting model are adopted for orbit control strategy solving, and orbit parameters at the four-device separation point moment are obtained through calculation; inner and outer ring strategy solving is carried out according to four-device separation point moment orbit parameters; and outputting an orbit control strategy calculation result. According to the method, the characteristic of perturbation rate difference of lunar elliptical orbit planes of different periods is utilized, a 4h lunar parking elliptical orbit is designed, the parking orbit flight time corresponding to the orbit plane adjustment amount is determined by establishing the drift rate difference between the 4h lunar parking elliptical orbit and a 2h target lunar orbit plane, the flight time of the nearly one month from capturing to descending is fully utilized, and the flight speed of the target lunar orbit plane is determined. On the premise of not additionally consuming a propellant, high-precision aiming of a landing point is realized.

The invention provides a spacecraft cluster density deployment method and device based on decoupling control, and belongs to the technical field of universe navigation. According to the method, the state of each spacecraft in the spacecraft cluster can be directly and accurately adjusted from the inside and outside of the orbital plane through the density control speed field of the spacecraft cluster, and the density control effect of the spacecraft cluster is improved while the calculation complexity of density control of the spacecraft cluster is reduced.

This invention provides a method and apparatus for adjusting the inclination of a sun-synchronous orbit. The method includes: determining the positional relationship between an initial orbit and a target orbit, such that the two intersection points of the initial orbit and the target orbit are located above the equator, and using one of the intersection points as the satellite's orbit change point; determining the orbit change angle required for the satellite to adjust from the initial orbital plane to the target orbital plane; comprehensively determining the satellite's single ignition duration, ignition time, and shutdown time based on the satellite's system status and mission objectives; determining the satellite's ignition direction; analytically calculating the velocity increment of the satellite's single pulse ignition, and determining the total impulse required to change the satellite's orbit inclination based on the velocity increment; analytically calculating the total number of satellite ignitions based on the total impulse and the actual impulse generated by each ignition; and controlling the satellite to perform a single ignition orbit change to complete the inclination adjustment based on the satellite's orbit change point, ignition direction, ignition time, and shutdown time.

The embodiment of the application discloses a two-degree-of-freedom discontinuous rotation control method and system for a solar wing, the discontinuous rotation control method comprises the following steps: obtaining the angle between a sun vector and the normal of a solar wing along the direction of orbit operation and the angle between the sun vector and the normal of the solar wing in the direction perpendicular to the orbit plane, so as to obtain the angle between the sun vector and the normal of the solar wing in a first coordinate projection plane and a second coordinate projection plane, and further obtain the driving target instruction of a first control shaft and a second control shaft; determining the position of the satellite, if the satellite is in a sunlit area, driving the first control shaft to make the sun vector coincide with the normal of the solar wing in the first coordinate projection plane, then driving the second control shaft to adjust the position angle of the solar wing relative to the orbit plane, so that the normal of the solar wing coincides with the sun vector in the second coordinate projection plane; if the satellite is in a shadow area, driving the first control shaft to execute a rotation instruction.

The invention discloses a high-precision space target orbit forecasting method based on numerical fitting, and relates to the technical field of space target orbit forecasting, frequency domain decomposition is combined with a harmonic compensation model, and seasonal system drift caused by sun-moon gravity is thoroughly eliminated; high-frequency components peel off attitude noise through light pressure parameterization and smooth difference, and perturbation coupling error transmission is avoided; a harmonic order adaptive mechanism captures a main disturbance term with the minimum calculation amount; carrying out least square batch estimation to realize coefficient primary solution; the template library pre-loading technology effectively compresses the time consumed for new target initialization; in addition, a geomagnetic modulation factor is introduced to dynamically suppress compensation distortion during a magnetic storm period, a sliding window mechanism guarantees real-time updating of a coefficient, and stable forecasting is still maintained under an extreme event; attitude measurement is replaced by luminosity curve inversion, so that non-cooperative target forecasting becomes possible; tangential component directional compensation ensures the stability of an orbital plane, and reduces the fuel consumption of satellite position protection.