122 results about "Orbit/Orbital" patented technology

The invention relates to an earth-moon space low-thrust spacecraft fuel optimal collision avoidance trajectory implementation method, which comprises the following steps of: constructing a kinetic model based on calculation parameters of a spacecraft collision avoidance problem, and establishing a fuel optimal performance index of the spacecraft collision avoidance problem through a homotopy method; the optimal control problem is converted into a two-point boundary value problem, targeting solving is carried out, and the state and co-state parameters of the optimal orbit are obtained; and according to the obtained state and co-state parameters of the optimal orbit, solving to obtain collision avoidance optimal control and collision probability of the spacecraft. According to the method, the fuel optimal control problem is solved by adopting the indirect method, the convergence difficulty of the low-thrust fuel optimal problem is reduced through the homotopy method, and when the thrust is calculated, the calculation efficiency is high, the precision is high, and the method is suitable for the fuel optimal collision avoidance scene of the low-thrust spacecraft for a single target in the earth-moon space.

The invention relates to a method for designing a transfer orbit from a DRO orbit to a lunar low-orbit circular orbit. The method comprises the following steps: establishing a mapping data table and an orbital transfer speed increment three-dimensional data pair according to a preset DRO orbit size parameter and a DRO aircraft initial moment phase angle; dividing the orbital transfer speed increment three-dimensional data pair into two groups according to positive and negative conditions of an orbital transfer speed increment elevation angle, acquiring ascending node longitude and transfer flight time corresponding to the data pair of the speed increment in the minimum transfer process of each group to calculate orbital transfer moments, and selecting a minimum value in the two orbital transfer moments as a selected orbital transfer moment; correcting the phase at the selected orbital transfer moment, and calculating an orbital transfer speed increment projection according to the corrected phase; and adopting a Newton iteration method to take the projection of the selected orbital transfer moment and orbital transfer speed increment as an iteration initial value to carry out iteration solution, and obtaining the transfer orbital transfer moment and the transfer orbital transfer speed increment. By adopting the method, the transfer track design precision can be improved.

The invention discloses a satellite orbit space environment disturbance coping method and device and a computer program product, and the method comprises the steps: obtaining the current position data and current space environment data of a target orbit, the current space environment data comprises a first high-energy particle integral flux at a high-energy particle flux target position and a first upper atmosphere daily average atmospheric density; according to the current position data, querying a second high-energy particle integral flux and a second upper atmosphere daily average atmospheric density at the high-energy particle flux target position when the target orbit environment is quiet; calculating the relative disturbance quantity of the high-energy particle radiation environment of the target orbit and the maximum relative disturbance quantity of the atmospheric density of the upper layer; determining the current high-energy particle radiation environment disturbance degree and the current atmospheric environment disturbance degree in combination with the disturbance degree judgment range; and generating risk coping suggestions. According to the invention, the operation environment disturbance condition of the spacecraft can be accurately known, and the in-orbit operation safety of the spacecraft is improved.

The invention relates to an orbit identification method based on earth-moon collinear translation point orbit parameter characterization. The method comprises the following steps: describing a circular restrictive three-body problem as a chaotic Hamiltonian power system, converting a motion equation into a new coordinate system for series expansion, and taking a quadratic term of a Hamiltonian function in a polynomial form as a linear model of a translation point in the circular restrictive three-body problem; and carrying out complex transformation and regular transformation on the linearized model transformation to realize the decoupling of the hyperbolic unstable direction and the central direction of the central manifold, and constructing a mapping relation between CRTBP coordinates and characterization parameters to carry out parameter characterization. Selecting spacecraft coordinates from parameter characterization to calculate characterization parameters of a reference orbit and integral initial values of the coordinates, and then calculating mean square errors of a real orbit and the reference orbit; and by taking the mean square error as a target function and a preset constraint condition, constructing an orbit identification solving model and solving to obtain an orbit identification result. By adopting the method, the track identification accuracy can be improved.

The invention provides a low-energy transfer window rapid design method and system for earth-moon DRO, and relates to the technical field of aerospace. Based on a double-circle restrictive four-body model, an earth-moon transfer orbit database is constructed, the earth-moon transfer orbit database comprises a plurality of transfer orbits from a target near-earth orbit to a target long-distance retrograde orbit; determining a target track point from the track distribution diagram corresponding to each track leveraging category according to the track information track distribution diagram of each transfer track and the input near-earth track information and epoch time of the target spacecraft, and correcting the track information of the transfer track corresponding to the target track point based on an ephemeris model, and obtaining a target transfer orbit of the target spacecraft. According to the method provided by the invention, rapid and accurate planning of the earth-moon DRO low-energy transfer window can be realized, the processing time is reduced, and the planning efficiency is improved.

The invention provides a space target orbit maneuver identification method and system based on adaptive LSM, and the method comprises the steps: obtaining the orbit data of a space target, calculating the six orbit elements based on the orbit data, and coding the six orbit elements into a pulse sequence; and inputting the pulse sequence into an improved adaptive LSM for detection to obtain an orbit maneuver behavior identification result of the space target. According to the method provided by the invention, six orbital elements are calculated through the speed and the position of the orbital, then the orbital parameters are input into an improved LSM, and the improved LSM adopts an adaptive threshold neuron as a calculation unit, so that the threshold can be dynamically adjusted according to the current pulse response and attenuated along with time; meanwhile, an improved particle swarm optimization algorithm is combined to carry out optimization setting on an initial value of a neuron, so that the initial performance of the model is improved; and a learnable neuron connection weight based on an attention mechanism is introduced, so that the accuracy and reliability of orbit prediction are improved.

The invention discloses a real-time closed-loop pursuit game method and device for an elliptical orbit of a spacecraft, belongs to the crossing field of spacecraft orbit dynamics and intelligent game control, and particularly relates to the real-time closed-loop pursuit game method for the elliptical orbit of the spacecraft based on a convex optimization theory. The problems that a traditional orbit game control method is difficult in equilibrium solution, poor in strategy convergence and difficult in on-orbit application are solved. The method comprises the following steps: solving an equivalent discount unilateral minimum time transfer problem to obtain an open-loop optimal strategy pair; according to the optimal control strategy, the state of the spacecraft is subjected to recursion, the state of the escape spacecraft under unknown control is subjected to recursion, and a target point is updated. The real-time closed-loop pursuit game method and device for the elliptical orbit of the spacecraft are suitable for the real-time pursuit game of the elliptical orbit of the spacecraft.

The application provides an orbit transfer strategy optimization method and system for a large-ellipse frozen orbit satellite orbiting, comprising the following steps: acquiring a satellite-rocket separation orbit and a station parameter, calculating an elevation angle of the satellite and the station, and then determining a range of a true anomaly angle of pulse ignition meeting a measurement and control requirement; in combination with the range of the true anomaly angle of pulse ignition, generating a fuel most-probable pulse orbit change strategy of the satellite from an initial orbit to a target orbit, determining an orbit change speed increment, an ignition position and an ignition frequency; determining ignition-on and ignition-off conditions of each orbit change and a drift circle number between adjacent two orbit changes; according to the pulse orbit change strategy, respectively determining an initial value of an ignition time length, an ignition time and a thrust direction of each orbit change, and determining the final ignition time length, the ignition time and the thrust direction of each orbit change. The application adopts a non-apogee double-pulse orbit change strategy to design a fuel optimal orbit transfer strategy of a large-ellipse frozen orbit satellite orbiting, and solves the design method problem of the orbit transfer strategy.

The present invention provides a satellite constellation electric propulsion deorbiting control method and system, which relates to the technical field of spacecraft orbit design and control. The method includes: Step S1: According to the deorbiting mission requirements, select the characteristic point height of the deorbiting orbit, and calculate the semi-major axis and eccentricity of the target orbit; Step S2: Adjust the semi-major axis of the orbit, continuously fire along the tangential direction until the semi-major axis is reduced to the target value; Step S3: Adjust the eccentricity of the orbit, fire throughout the orbit, and fire along the direction perpendicular to the apsidal line until the eccentricity meets the target value; Step S4: Fire at the orbit characteristic point until the perigee height and apogee height meet the deorbiting requirements. The present invention can achieve rapid evaluation and optimization of the deorbiting orbit parameter set; by adopting the strategy of firing in different coordinate systems in stages, the coupling between the semi-major axis and the eccentricity is reduced, which is beneficial to simplifying the complexity of the actual mission deorbiting scheme.

The invention discloses a strict regression frozen orbit design method. The method comprises the following steps: calculating an initial estimated value of a target strict regression frozen orbit mean element number based on an analytic formula of a regression orbit and a frozen orbit under a third-order gravity field; under the high-order gravity field dynamic model, iteratively correcting the orbit semi-major axis and the orbit inclination angle based on the right ascension change rate of the ascending node and the deviation value of the longitude of the revisit ascending node; taking the geocentric vector difference of the revisit ascending node as an optimization target, taking the eccentricity rate as an optimization variable, and adopting an optimization algorithm to obtain a current optimal frozen eccentricity rate; iteratively correcting the semi-major axis of the orbit and the inclination angle of the orbit, and optimizing the eccentricity ratio to obtain the orbital number of the strictly frozen regression orbit.

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

The application discloses a kind of space target ultra-short arc initial orbit intelligent determination method, belong to aerospace technology field.The application is based on artificial intelligence technology to realize space target ultra-short arc initial orbit determination, there is no strict requirement to complex nonlinear problem, and can develop more complex perturbation condition and initial orbit determination under the operation law of multiple body, coverage is wider;The application generates multiple groups of target orbits randomly and calculates the observation and state quantity of each visible arc segment, constructs the sample data set with ultra-large capacity, which is helpful for neural network to learn the nonlinear relationship between observation and state quantity using rich samples;The application builds and trains long short-term memory network model, determines the optimal network model weight parameter by designing and adjusting hyperparameter, can obtain the fast prediction solution of unknown space target ultra-short arc initial orbit according to observation;The application can meet more space target ultra-short arc initial orbit determination tasks by considering different inputs and outputs.

The invention discloses a space-based angle-measurement-only target orbit determination method for compensating orbit position errors, and belongs to the technical field of satellite application measurement. The method comprises the following steps: establishing an error model of an empirical orbit dynamics model error introduction measurement error according to a real orbit dynamics model of a target; training the initial neural network according to a preset training sample to obtain an orbit position error estimation model; wherein the training sample is determined through historical orbit data and an empirical orbit dynamics model; and according to the orbit position error estimation model, compensating the error model, and processing the initial state of a to-be-measured target by using a deviation compensation least square method to obtain an optimal estimation value of the orbit of the to-be-measured target. The method can effectively improve the determination precision of the target track.

The utility model discloses an assembled astronavigation star orbit, and relates to the field of simulation models. The split mounting type astronavigation star orbit comprises a simulation star orbit unit, a driving unit is installed in the simulation star orbit unit, and the simulation star orbit unit is matched with the driving unit to be in transmission connection with a simulation celestial body unit; the simulation star orbit unit comprises a bottom plate and a simulation orbit, a plurality of clamping plates are clamped to the outer side of the bottom plate, and clamping grooves are formed in the clamping plates; the outer edge of the simulation track is attached to the inner walls of the multiple clamping grooves, and the simulation track is in transmission connection with the driving unit. According to the split mounting type astronavigation star orbit, in the using process, the simulation orbit can be mounted only by clamping the multiple clamping plates to the outer side of the bottom plate and enabling the outer edge of the simulation orbit to be attached to the inner walls of the multiple clamping grooves, at the moment, the driving unit and the simulation celestial body unit are correspondingly mounted, and then overall mounting of the astronavigation star orbit can be completed; installation is simple, and the device is suitable for users of different ages.

The application relates to a design method of a moon global positioning constellation based on a lunar space DRO reference orbit, wherein the initial orbit state and the orbit period of any DRO reference orbit in a DRO reference orbit to be deployed are determined based on the initial phase of the DRO reference orbit; the initial orbit state of any satellite to be deployed on the initial phase of the DRO reference orbit is determined based on the orbit state corresponding to the target phase; the orbit period of the any satellite is determined based on the orbit period of the any DRO reference orbit; the target orbit state corresponding to the any satellite is obtained based on the orbit period of the any satellite and the initial orbit state of the any satellite; and the moon global positioning constellation under a four-body model is constructed based on the target orbit state corresponding to each satellite. The moon global positioning constellation constructed by the application has higher stability and higher moon surface coverage capacity compared with the existing near-moon constellation.

The present invention discloses a method for designing a companion satellite orbit for space station aimed at early warning of space debris threats, and belongs to the technical field of aerospace. The present invention provides a method for designing a companion satellite orbit for space station aimed at early warning of space debris threats, which writes the CW equation into a parameterized form by combining the parametric equation of the space ellipse, and then uses the geometric meaning of the parameters in combination with the azimuth and distance constraints to give the orbit design result. The design method proposed in the present invention makes full use of the geometric characteristics and mathematical geometry methods of the CW equation, and has the advantages of being intuitive, having obvious physical meaning, and being easy to calculate; at the same time, the present invention supplements the design method of the flyby orbit for early warning of space debris impact in the existing field, and provides ideas, references and technical support for subsequent references in engineering fields.

This invention discloses a method and apparatus for planning a timed, fixed-point landing and timed takeoff for lunar sampling missions. The method includes: determining the orbital root number and orbital insertion time of the orbit in the first control phase of the lunar sampling mission, as well as the nominal parameters for timed, fixed-point landing and timed takeoff. Based on the nominal parameters, orbital root number, and orbital insertion time, the first control phase is jointly planned with subsequent control phases to obtain a first deviation, and the first deviation is iterated based on a first variable. When the first deviation is less than a convergence threshold, the target variable for the first control phase is determined, allowing the satellite to achieve timed, fixed-point landing and timed takeoff based on the target variables of each control phase. This achieves dynamic joint planning with timed, fixed-point landing and timed takeoff as the final goal for each control phase, effectively improving the mission's control reconfiguration capability, avoiding additional control operations, and reducing resource consumption.

This invention relates to the field of collision warning technology, specifically to a spacecraft collision warning method for a large number of space targets. The method includes: selecting a multi-data parallel screening warning approach; screening secondary target orbits within the TLE data set of the complete collision target orbits using set pre-screening thresholds for date expiration, apogee / perigee pre-screening safety margins, orbital intersection pre-screening thresholds, and time pre-screening thresholds; performing SGP4 orbit prediction on the primary target orbit and the initially screened secondary target orbits; obtaining the minimum spatial distance and collision probability between spacecraft within the primary and secondary target orbits; summarizing the primary and secondary target orbit numbering information, as well as the minimum spatial distance, collision probability, and the time corresponding to the minimum spatial distance, into a collision warning result; and summarizing all collision warning results to output a collision warning set information. This invention can filter out targets that are unlikely to intersect with the spacecraft to the greatest extent possible, reducing the computational load of subsequent orbit prediction.

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 invention provides a nonlinear degree evaluation method suitable for a Lyapunov orbit near an earth-moon space stable translation point, and belongs to the field of celestial body measurement and celestial body mechanics. Comprising the following steps: generating a plurality of sample points near an initial value of a specified reference orbit according to Gaussian distribution of an initial error; calculating a propagation result of each sample under a floating-point number framework at any time length, gradually increasing an order k, integrating a reference orbit under a DA framework to obtain a k-order Taylor expansion of an integral end moment state quantity relative to an initial moment state quantity, calculating a k-order approximation value of a sample point by using the Taylor expansion, further calculating a local non-linear index eta, and researching the change of the local non-linear index eta; and when eta is smaller than the acceptable tolerance, determining the non-linear degree of the simulation data as the corresponding k at the moment, and ending the search order calculation. The high-order uncertainty propagation method based on the high-order uncertainty propagation method has the advantages that the high-order uncertainty propagation method ensures that the precision is close to the uncertainty expression methods such as Monte Carlo, and meanwhile, the efficiency is greatly improved.

The application discloses a transfer orbit design method and system of a lunar orbit large ellipse frozen orbit based on eccentricity vector targeting. The transfer orbit lifting type is determined in advance in a transfer stage, the transfer orbit is selected as a lifting orbit, launching and arriving are lifted, the lunar orbit is ensured as a large ellipse frozen orbit with a far point above the south hemisphere of the moon, the eccentricity vector targeting is used to generate a transfer orbit meeting task constraints, the near point amplitude angle is adjusted near the far point after the moon is captured, the lunar orbit large ellipse frozen orbit is efficiently generated, and finally, a pulse is applied at the near point according to task requirements to transfer to a large ellipse frozen orbit with a required period.

The invention discloses a rotating tether system maneuvering method for load transmission between different-plane orbits, and the method systematically designs a load throwing mechanism by constructing a double-body rotating tether system architecture: firstly, planning a load release mode based on the characteristics of a double-body rotating tether system; deducing speed increment and position increment parameters required for realizing the release mode; by calculating the total mass and length of the tether and combining the relative position vector and the velocity vector between the end body satellites required by load throwing, the key parameters of the throwing rotation surface are accurately calculated; and finally, constructing a rotating surface maneuvering control strategy of a full task period based on the parameters. Compared with a traditional coplanar orbit transfer scheme, the non-coplanar orbit transfer task of the effective load from the initial inclined low earth orbit to the target equator geostationary orbit is achieved in a breakthrough mode. The method has obvious technical reference value and engineering application prospect in orbit inclination angle adjustment maneuvering tasks in a satellite launching stage.

The invention discloses a low-orbit satellite everyday transit circle calculation method, which comprises the following steps of: firstly, determining the descending node longitude of an orbit where each point on a target point visual space circle is located, uniformly distributing the points on the target point visual space circle and enabling each point to correspond to two orbits, and then calculating the transit circle number of a low-orbit satellite on the basis of the descending node longitude of each point, determining a descending node longitude range which is continuously visible to the target point, and finally determining the average orbit number of the satellite which continuously passes through the visible space circle of the target point every day based on the descending node longitude range. According to the calculation method, the transit circle number is determined based on the geometric coverage characteristics of the satellite and the ground target point, and the transit characteristics of the orbit under different orbit heights and inclination angles can be rapidly obtained.

The present invention provides a semi-analytical multi-pulse elliptical orbit out-of-plane rendezvous optimization method. The method first uses a four-pulse rendezvous velocity increment rapid estimation method to obtain estimated four-pulse initial parameters. A precise dynamics model is then used to predict the difference between the orbit after the four-pulse maneuver and the target orbit to be rendezvoused. An analytical expression is then given to calculate a correction algorithm for the four-pulse size and timing. By iteratively repeating the correction and prediction steps, the optimal four-pulse vector and application timing are ultimately obtained, resulting in a zero difference between the orbit after the pulse maneuver and the target orbit to be rendezvoused. The method can be applied to low-orbit space debris removal and on-orbit servicing mission orbit optimization design, resolving the issues of multiple local optimal solutions and low computational efficiency during long-term out-of-plane rendezvous optimization of perturbed orbits.

The present invention discloses a method and system for optimizing the orbital parameters of a double star, including: Step 1) According to the six orbital elements of the primary star, calculate the time interval to obtain the <time, position, velocity> information matrix of the primary star and the theoretical solution set of the <time, position, velocity> information matrix of the secondary star; Step 2) Set the constraint conditions, establish an optimization model for the orbit of the secondary star, and according to the set solution granularity, use the dynamic mutation genetic algorithm. By setting the mutation operation probability that changes dynamically with the iteration step size, optimize the six orbital elements of the secondary star to obtain the actual solution of the <time, position, velocity> information matrix of the secondary star, and use the number of occurrences of the actual solution in the theoretical solution set as the fitness function; Step 3) Calculate the fitness function to obtain the continuous duration of maintaining communication between the two stars of the double star, and judge whether the continuous duration remains unchanged. If it is judged to be yes, go to Step 4), otherwise, go to Step 2); Step 4) Output the positions of the primary star and the secondary star that meet the conditions.

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 application discloses a satellite orbit space environment disturbance coping method, equipment and computer program product. The method comprises the following steps: acquiring current position data and current space environment data of a target orbit, wherein the current space environment data comprises a first high-energy particle integrated flux at a target position of a high-energy particle flux and a first high-altitude atmospheric daily average atmospheric density; querying a second high-energy particle integrated flux at the target position of the high-energy particle flux and a second high-altitude atmospheric daily average atmospheric density when the target orbit environment is calm according to the current position data; calculating a high-energy particle radiation environment relative disturbance amount and a high-altitude atmospheric density maximum relative disturbance amount of the target orbit; determining a current high-energy particle radiation environment disturbance degree and a current atmospheric environment disturbance degree in combination with a disturbance degree discrimination range; and generating a risk coping suggestion. The application can accurately understand the disturbance situation of a spacecraft operating environment and improve the safety of the spacecraft in-orbit operation.

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 application discloses an online trajectory planning method for the final stage of a carrier rocket aiming at an elliptical target orbit and belongs to the field of carrier rocket trajectory and guidance, which comprises the following steps: step one, constructing the orbit entry terminal constraint and planning performance index aiming at the elliptical target orbit; step two, establishing the system dynamics constraint of the orbit entry of the final stage of the carrier rocket; step three, convex approximation is conducted on the planning orbit entry terminal constraint and performance index function; step four, state affine approximation and pseudo-spectrum discretization are conducted on the system dynamics constraint of step two; and step five, the optimal trajectory of the final stage entering the elliptical orbit is generated online through sequential convex iteration. The method of the application is suitable for the joint reconstruction of the final stage trajectory and the adaptive guidance of the carrier rocket launching the elliptical target orbit mission, can realize the online generation of the optimal trajectory with high precision and fast convergence, and is favorable for improving the reliability of the carrier rocket and enhancing the launch service fulfillment capability.

The present invention provides a geomagnetic detection method, device, launch vehicle, electronic device, and storage medium, relating to the field of aerospace technology. The method includes: when a remaining orbit final stage passes over multiple target detection areas on a current target orbit, controlling a geomagnetic detection payload carried by the remaining orbit final stage to obtain geomagnetic measurement data corresponding to the multiple target detection areas in the current target orbit; determining a first orbit change control parameter for controlling the remaining orbit final stage to change its orbit from the current target orbit to the next target orbit based on orbital parameters of the current target orbit and orbital parameters of the next target orbit; causing an onboard computer to control the remaining orbit final stage to change its orbit to the next target orbit based on the first orbit change control parameter; and constructing an Earth magnetic field model based on the geomagnetic measurement data corresponding to each target detection area in the multiple target orbits. The method and device provided by the present invention improve the equipment reuse rate of the launch vehicle's final stage and reduce the cost of geomagnetic detection.