Space target dynamic profile analysis method

Abstract

This invention relates to a method, electronic device, and storage medium for the dynamic profile analysis of space targets. The method includes the following steps: Step S1, Object group selection: Selecting two space targets as the analysis object group; Step S2, High-precision orbit calculation: Considering various perturbation factors, and combining error tolerance and dynamic step size control strategies, performing high-precision numerical integration calculations on the analysis object group; Step S3, Relative relationship calculation: Based on the high-precision orbit calculation results of Step S2, establishing a relative motion coordinate system and solving the relative motion differential equations; Step S4, Dynamic profile visualization: Integrating the high-precision orbit calculation results of Step S2 and the relative relationship calculation results of Step S3, synchronously displaying the comprehensive dynamic profile of the analysis object group in a three-dimensional scene. This invention can analyze the relative motion relationships of space targets and use a three-dimensional model to display the spacecraft motion profile, enabling a hierarchical and intuitive display of complex multi-target spatial relationships.

Description

Technical Field

[0001] This invention relates to the field of aerospace simulation and space situation analysis, and in particular to a dynamic profile analysis method for space targets that integrates high-precision orbit calculation, multi-target relative relationship analysis and three-dimensional visualization. Background Technology

[0002] In recent years, the global aerospace industry has entered a period of rapid development, with an explosive growth in the number of satellites in orbit, increasingly strained space orbital resources, and a continuous rise in demand for complex aerospace missions such as multi-satellite collaborative operations and space target monitoring. This trend places higher demands on aerospace mission planning, analysis, and visualization technologies, urgently requiring efficient and accurate simulation analysis tools to provide support. Aerospace simulation software, as a core tool for aerospace mission design, orbit analysis, and space target management, has become an indispensable technical support in the aerospace field.

[0003] Currently, mainstream analysis tools typically treat orbit prediction, relative motion analysis, and 3D visualization as separate functional modules, leading to the following technical problems:

[0004] Fragmented analysis process: Users need to manually configure data interfaces and formats between different software or modules, which is cumbersome and makes it difficult to achieve dynamic interaction and rapid iterative analysis.

[0005] It is difficult to balance computational accuracy and efficiency: The full numerical integration method using high-precision dynamic models has a large computational load, while commonly used analytical or linear simplified models ignore key perturbation forces, resulting in significant medium- and long-term forecast errors, which are difficult to meet the requirements of high reliability tasks.

[0006] Visualization and analysis are deeply disconnected: Existing 3D visualization tools mostly focus on showing the absolute motion trajectory of the target, lacking integrated analysis and linkage display of multi-dimensional dynamic parameters such as relative distance between targets, speed changes, and relative attitude of the orbital plane. This is not conducive to a deep understanding of complex scenarios such as spatial proximity, formation configuration, and maneuver strategies.

[0007] Therefore, there is an urgent need in this field for an automated analysis method that can deeply integrate high-precision computing, correlation analysis and multi-dimensional visualization to improve the completeness, efficiency and intuitiveness of analyzing complex spatial tasks. Summary of the Invention

[0008] This invention aims to overcome the shortcomings of the prior art and provide a method for cross-sectional analysis of dynamically changing spatial targets. The core concept of this invention lies in constructing an end-to-end automated analysis pipeline from data input to graphical output through collaborative innovation in three levels: process architecture, computational algorithm, and visualization mode. This systematically solves the problems of fragmented processes, imbalance between accuracy and efficiency, and lack of intuitive analysis.

[0009] To achieve the above-mentioned objectives, this invention provides a method for cross-sectional analysis of dynamically changing space targets, comprising the following steps:

[0010] Step S1, Object Group Selection: Select two spatial targets as the analysis object group;

[0011] Step S2, High-precision orbit calculation: Considering various perturbation factors, high-precision numerical integration calculation of the analysis object group is performed by combining error tolerance and step size dynamic control strategy; the error tolerance and step size dynamic control strategy combines the local error control mechanism of RKF78 integrator with the global confidence interval judgment standard based on Monte Carlo simulation in a closed loop, and dynamically controls the internal error tolerance and step size strategy of RKF78 in reverse.

[0012] Step S3, Relative Relationship Calculation: Based on the high-precision orbit calculation results in Step S2, establish a relative motion coordinate system with one spatial target in the high-precision orbit calculation as the reference body and the other spatial target as the target body, and solve the relative motion differential equation of the target body relative to the reference body;

[0013] Step S4, Dynamic Profile Visualization: Integrating the high-precision orbit calculation results from Step S2 with the relative relationship calculation results from Step S3, a comprehensive dynamic profile of the analysis object group is synchronously displayed in a three-dimensional scene.

[0014] According to one technical solution of the present invention, step S2 specifically includes:

[0015] Step S21: Obtain calculation data, including ephemeris data, atmospheric density data, and gravitational parameters of the sun and moon;

[0016] Step S22: Establish a precise orbit calculation model for the space target;

[0017] The precise orbit calculation model for the space target is a mechanical model that uses an inertial coordinate system, takes the two-body orbit calculation results as initial values, and superimposes perturbation accelerations. Factors affecting the perturbation accelerations include atmospheric disturbances, the Earth's non-spherical gravitational field, gravitational perturbations from the Sun and Moon, the effects of the solar wind and the Earth's magnetic field, and the satellite's own orbital perturbation forces. The orbit calculation results of the two-body model are obtained through the ephemeris data.

[0018] Step S23: Solve the orbital motion equations using an RKF78 integrator. Combine Monte Carlo simulation analysis with dynamic reverse adjustment of the RKF78 integrator's internal error tolerance and step size strategy within the global confidence interval to optimize computational efficiency and perform accurate orbit calculations for the space target.

[0019] According to one technical solution of the present invention, step S23 specifically includes:

[0020] Step S231: Construct the motion state vector of the spatial target and its differential equation;

[0021] Step S232: Set the initialization conditions for the calculation, including the initial time point, initial solution, and error tolerance;

[0022] Step S233: Calculate the approximate solutions with 7th-order and 8th-order accuracy within each time step;

[0023] Step S234: Estimate the local truncation error based on the difference between the two approximate solutions calculated in step S233;

[0024] Step S235: Adjust the time step based on the local truncation error;

[0025] Step S236: Advance all time steps sequentially according to the adaptive step size to complete the spatial target orbit calculation at the target time;

[0026] Step S237, Repeated Iteration: Analyze the contribution of perturbation error through Monte Carlo simulation, use global accuracy as the criterion, dynamically adjust the error tolerance of the integrator in reverse, and execute steps S232 to S236 until the result meets the preset accuracy requirements of the confidence interval.

[0027] According to one technical solution of the present invention, step S234 specifically includes:

[0028] Calculate the difference between two approximate solutions This is used to estimate the local truncation error at the current step size. , is represented as:

[0029]

[0030] in, and They are respectively At each step, approximate solutions with 7th and 8th order precision are obtained through calculation.

[0031] Step S235 specifically includes:

[0032] Define scalar error Given an error tolerance τ, calculate the new step size using the following formula. :

[0033]

[0034] in, For safety reasons, and This is the step size variation limiting factor.

[0035] like ≤τ, accept and with a new step size Proceed to the next time step to continue integration; otherwise, reject the result of the current time step and proceed with the new step size. Recalculate the results at the current time step.

[0036] According to a technical solution of the present invention, in step S237, the error tolerance of the reverse dynamic adjustment integrator specifically includes:

[0037] The initial trajectory error and key model parameter errors are modeled as random vectors. Draw N independent samples from a random vector. ;

[0038] For each independent sample drawn This allows for the correction of initial conditions and model parameters, using the current error tolerance. The RKF78 integrator is used for orbit prediction to obtain the position sample at the target time. ;

[0039] Calculate the mean of N location samples. The sample covariance matrix S is expressed as:

[0040]

[0041] Using a 1σ confidence level, the confidence interval for the sphere radius is approximately calculated using the quantiles of the sample error modulus. , is represented as:

[0042]

[0043] according to If the accuracy is less than the preset confidence interval, the calculation result is considered to meet the preset confidence interval accuracy requirement, and efficiency is improved by increasing the error tolerance; if If the accuracy exceeds the preset confidence level, the calculation result is considered not to meet the preset confidence interval accuracy requirements, and the integration requirements should be tightened to reduce the error tolerance.

[0044] According to one technical solution of the present invention, step S3 specifically includes:

[0045] Step S31: Select the RTN coordinate system, describe the linear approximate relationship of the relative motion of the two space targets through the CW equation, and construct the differential equation of relative motion;

[0046] Step S32: Solve the relative motion differential equation to obtain the position and velocity changes of the target body relative to the reference body;

[0047] Step S33: Introduce an atmospheric density model to correct the calculation results based on step S32, thereby optimizing the calculation results of the relative relationship of spatial targets.

[0048] According to one technical solution of the present invention, step S33 specifically includes:

[0049] The difference in tangential drag acceleration between the target body and the reference body is added as a non-secondary term to the tangential equation of the relative motion differential equation, and the relative motion differential equation is solved again to obtain the corrected solution, thus completing the optimization of the calculation results of the relative relationship of the space target.

[0050] The difference in tangential drag acceleration between the target body and the reference body is expressed as:

[0051]

[0052] In the formula, ρ is the atmospheric density. and Let represent the ballistic coefficients of the target body and the reference body, respectively. The ballistic coefficient of spacecraft i is expressed as: , Indicates the drag coefficient. This represents the cross-sectional area of ​​a spacecraft facing the wind. The spacecraft mass is represented by v; v is an approximate value for the orbital velocity.

[0053] The tangential equation of the relative motion differential equation with the addition of non-secondary terms is expressed as:

[0054]

[0055]

[0056] In the formula, Indicates tangential relative acceleration. It is the orbital angular velocity of the reference body. It is the gravitational constant of the central celestial body. It is the semi-major axis of the reference body orbit. It represents the difference in acceleration of external forces in the tangential direction.

[0057] According to one technical solution of the present invention, step S4 specifically includes:

[0058] Step S41: Initialize the 3D profile simulator and add spatial targets;

[0059] Step S42: Obtain the positions of the two space targets based on the accurate orbit calculation results of the space targets, and calculate the Euclidean distance between the two space targets;

[0060] Step S43: Obtain the relative position and relative velocity of the two space targets based on the calculation results of the relative relationship between the two space targets, and calculate the orbital plane relationship between the two space targets;

[0061] Step S44: Display the spacecraft motion profile on the three-dimensional model based on the calculation results of Step S42 and Step S43. The spacecraft motion profile includes the absolute motion trajectory, relative distance, relative velocity, and relative relationship of the orbital plane of the space target.

[0062] According to one aspect of the present invention, an electronic device is provided, comprising: one or more processors, one or more memories, and one or more computer programs; wherein the processor is connected to the memory, and the one or more computer programs are stored in the memory; when the electronic device is running, the processor executes the one or more computer programs stored in the memory to cause the electronic device to perform a profile analysis method for dynamic changes of a space target as described in any of the above technical solutions.

[0063] According to one aspect of the present invention, a computer-readable storage medium is provided for storing computer instructions, which, when executed by a processor, implement a profile analysis method for dynamic changes of a space target as described in any of the above technical solutions.

[0064] Compared with the prior art, the present invention has the following advantages:

[0065] (1) An integrated analysis process from computation to visualization was constructed: The high-precision orbit calculation, which includes error propagation analysis, enhanced relative motion solution, and multi-dimensional information fusion visualization were innovatively integrated into a coherent automated process. Users can complete the entire process from data input, precise analysis to intuitive display without switching between multiple platforms, which greatly improves the efficiency and depth of analysis for complex space missions.

[0066] (2) Balancing high accuracy and high efficiency in orbit prediction: By combining adaptive step size control of the RKF78 integrator with global confidence interval evaluation based on Monte Carlo simulation, a dynamic closed-loop control mechanism is formed. This mechanism can intelligently optimize the allocation of computational resources while ensuring the long-term accuracy of orbit prediction, effectively overcoming the technical contradictions of high computational load of high-precision models and significant errors of simplified models in traditional methods. In addition, in relative motion analysis, an atmospheric density model is introduced specifically for drag perturbation compensation for low-Earth orbit targets, correcting the radial drift error caused by neglecting atmospheric drag in traditional linear models, and significantly improving the reliability of relative position and velocity prediction between low-Earth orbit space targets.

[0067] (3) Provides a multi-dimensional and interconnected in-depth analysis perspective: In the three-dimensional visualization scene, not only is the absolute motion trajectory of the space target displayed synchronously, but also multi-dimensional dynamic relationship parameters such as relative distance, relative speed, and relative attitude of the orbital plane are integrated. Real-time linkage between the visualization interface and the analysis data is realized, which greatly enhances the intuitive understanding and analysis depth of complex scenarios such as space approach events, formation configuration evolution, and orbital maneuver strategies. Attached Figure Description

[0068] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly described below. Obviously, the drawings described below are merely some embodiments of the present invention, and those skilled in the art can obtain other drawings based on these drawings without creative effort.

[0069] Figure 1 A flowchart illustrating a method for cross-sectional analysis of dynamic changes in a space target according to an embodiment of the present invention;

[0070] Figure 2 This is a schematic diagram illustrating a three-dimensional profile of a cross-sectional analysis method for dynamically changing spatial targets provided in an embodiment of the present invention. Detailed Implementation

[0071] The description of the embodiments in this specification should be taken in conjunction with the accompanying drawings, which should form part of the complete specification. In the drawings, the shape or thickness of the embodiments may be exaggerated and may be indicated in a simplified or convenient manner. Furthermore, parts of the various structures in the drawings will be described separately; it is worth noting that elements not shown in the figures or not described in words are in a form known to those skilled in the art.

[0072] The descriptions of the embodiments herein, including any references to directions and orientations, are for ease of description only and should not be construed as limiting the scope of the invention. The following description of preferred embodiments involves combinations of features, which may exist independently or in combination; the invention is not particularly limited to the preferred embodiments. The scope of the invention is defined by the claims.

[0073] like Figure 1 As shown, the present invention provides a method for cross-sectional analysis of dynamic changes in space targets, comprising the following steps:

[0074] Step S1, Object Group Selection: Select two spatial targets as the analysis object group;

[0075] Step S2, High-precision orbit calculation: Considering various perturbation factors, and combining error tolerance and step size dynamic control strategy, high-precision numerical integration calculation is performed on the analysis object group;

[0076] When performing high-precision orbit calculations, various perturbation factors are considered. The local error control mechanism of the RKF78 integrator is combined with the global confidence interval judgment standard based on Monte Carlo simulation in a closed loop. The internal error tolerance and step size strategy of RKF78 are dynamically adjusted in reverse to optimize the calculation efficiency while ensuring the calculation accuracy.

[0077] Step S2 specifically includes:

[0078] Step S21: Obtain calculation data, including ephemeris data, atmospheric density data, and gravitational parameters of the sun and moon;

[0079] This invention is designed for the precise orbit calculation of space targets using three-dimensional profile analysis. It first acquires daily updated precise ephemeris data to ensure the accuracy of the ephemeris data and thus the accuracy of the basic data.

[0080] Step S22: Establish a precise orbit calculation model for the space target; the precise orbit calculation model for the space target is a mechanical model that uses an inertial coordinate system, takes the orbit calculation results of the two-body model as the initial value, and superimposes perturbation acceleration; the factors affecting the perturbation acceleration include atmospheric disturbance, Earth's non-spherical gravitational field, gravitational perturbation of the Sun and Moon, the effects of solar wind and Earth's magnetic field, and the orbital perturbation force of the satellite itself; the orbit calculation results of the two-body model are obtained by ephemeris data.

[0081] The core of accurate orbit calculation lies in "a suitable mechanical model + an efficient numerical integration algorithm". The mechanical model of this invention uses an inertial coordinate system (J2000 geocentric inertial frame), with the solution to the two-body problem as the initial value, and superimposed perturbation accelerations. Since satellite orbit measurements are affected by various factors, including but not limited to atmospheric disturbances, the Earth's non-spherical gravitational field, gravitational perturbations from the Sun and Moon, the effects of solar wind and the Earth's magnetic field, and the satellite's own orbital perturbations, these factors can all lead to errors in orbital parameters. Therefore, error analysis and control are crucial steps in improving calculation accuracy. This invention comprehensively considers the accuracy of profile analysis and processing efficiency, introducing a high-order Earth gravitational field model (such as EGM2008); an atmospheric drag model NRLMSISE-00, which is applicable to the Earth's upper atmosphere (50-2000 km) and considers the influence of solar activity and geomagnetic activity on density; and a solar / lunar gravitational model: directly using the instantaneous positions of the Sun and Moon (high-precision solar and lunar ephemeris) to calculate their gravitational acceleration on the target.

[0082] Step S23: Solve the orbital motion equations using the RKF78 integrator, and optimize computational efficiency by combining the error tolerance of global confidence analysis based on Monte Carlo simulation and the dynamic reverse adjustment strategy of time step to perform accurate orbit calculation for the space target.

[0083] Step S23 specifically includes:

[0084] Step S231: Construct the spatial target motion state vector and its differential equation, expressed as:

[0085]

[0086] in, It is the position vector of the spacecraft in the geocentric inertial frame. It is a velocity vector, and its differential equation is:

[0087]

[0088] In the formula, Represents the total perturbation acceleration. , This represents the acceleration due to two-body perturbation. This represents the acceleration due to the non-spherical gravitational perturbation of the Earth. This represents atmospheric drag acceleration. and These represent the gravitational perturbation acceleration of the Moon's third body and the gravitational perturbation acceleration of the Sun's third body, respectively.

[0089] Step S232: Set the initialization conditions for accurate orbit calculation. The initialization conditions include parameters such as initial time point, initial solution and error tolerance. The initial solution includes the spacecraft's initial position vector and velocity vector, which are obtained by converting ephemeris data according to orbital elements.

[0090] Step S233: Calculate approximate solutions with 7th and 8th order accuracy at each time step.

[0091] When calculating approximate solutions with 7th and 8th order accuracy, according to Moment State calculate Moment State ,in, h is the time step of the calculation;

[0092] In progress Moment State During the calculation, firstly, according to Moment State Calculate the 13 intermediate slopes :

[0093] =

[0094] =

[0095] =

[0096] …

[0097] =

[0098] in, and It is a fixed coefficient, obtained by looking up a table.

[0099] Then, using the 13 intermediate slopes obtained from the calculation We construct approximate solutions with 7th and 8th order accuracy, respectively, as follows:

[0100]

[0101]

[0102] in, and These represent the weighting coefficients for the 7th-order precision solution and the 8th-order precision solution, respectively.

[0103] Step S234: Estimate the local truncation error based on the difference between the two approximate solutions calculated in step S233, expressed as:

[0104]

[0105] Step S235: Adjust the time step based on the local truncation error:

[0106] Specifically, scalar error is defined. Given an error tolerance τ, calculate the new step size using the following formula. :

[0107]

[0108] in, For safety factors (usually taken as 0.8~0.9), and This is the step size variation limiting factor.

[0109] like ≤τ, accept and with a new step size Proceed to the next time step to continue integration; otherwise, reject the result of the current time step and proceed with the new step size. Recalculate the results at the current time step.

[0110] Step S236: Advance all time steps at once according to the adaptive step size to complete the spatial target orbit calculation at the target time;

[0111] Step S237, Repeated Iteration: Analyze the contribution of perturbation error through Monte Carlo simulation, use global accuracy as the criterion, dynamically adjust the error tolerance of the integrator in reverse, and execute steps S232 to S236 until the calculation result meets the preset accuracy requirements of the confidence interval.

[0112] The error tolerance of the integrator is adjusted dynamically in reverse, specifically including:

[0113] The initial trajectory error and key model parameter errors are modeled as random vectors. (e.g., random vectors satisfy a multivariate normal distribution) N independent samples are drawn from a random vector. The initial orbital error is the error in the spacecraft's initial position and velocity obtained through ephemeris calculations; the critical model parameter error is the calculation error of various perturbation terms.

[0114] For each independent sample drawn This allows for the correction of initial conditions and model parameters, using the current error tolerance. The RKF78 integrator is used for orbit prediction to obtain the position sample at the target time. ;

[0115] Calculate the mean of N location samples. The sample covariance matrix S is expressed as:

[0116]

[0117] Using a 1σ confidence level (e.g., confidence level α = 0.683 corresponds to 1σ), the confidence interval for the sphere radius is approximately calculated using the quantiles of the sample error modulus. , is represented as:

[0118]

[0119] like If the accuracy is less than the preset confidence interval, the calculation result is considered to meet the preset confidence interval accuracy requirement. To improve efficiency, the integration requirement can be relaxed, and the error tolerance can be increased. If the result exceeds the preset accuracy, it is considered that the calculation result does not meet the preset accuracy requirement of the confidence interval, and the integration requirement should be tightened to reduce the error tolerance.

[0120] Step S3, Relative Relationship Calculation: Based on the high-precision orbit calculation results in Step S2, establish a relative motion coordinate system with one spatial target in the high-precision orbit calculation as the reference body and the other spatial target as the target body, and solve the relative motion differential equation of the target body relative to the reference body;

[0121] The core of calculating the relative relationships of space targets is to establish a relative motion coordinate system between the "reference body and the target body," focusing on the influence of the central celestial body's gravity, perturbation force, and initial relative state. By solving the differential equations of relative motion, the changes in the target body's position and velocity relative to the reference body are obtained.

[0122] This invention uses the RTN coordinate system (Radial-Transverse-Normal) for relative trajectory calculation, and describes the relative motion of two spacecraft using the Clohessy-Wiltshire (CW) equations as a linear approximation. Furthermore, considering the significant impact of atmospheric drag on low-Earth orbit (<500km) space targets, which can cause radial drift in the relative orbit, this invention incorporates an atmospheric density model (such as NRLMSISE-00) for correction.

[0123] Step S3 specifically includes:

[0124] Step S31: Select the RTN coordinate system, describe the linear approximation relationship of the relative motion of the two space targets using the CW equations, and construct the differential equations of relative motion; the differential equations of relative motion are expressed in the following form:

[0125] Radial (x direction):

[0126] Tangential direction (y direction): ¨

[0127] Normal (z-direction): ¨

[0128] in, , , 、¨ , 、¨ and These represent radial relative acceleration, tangential relative velocity, radial relative position, tangential relative acceleration, radial relative velocity, normal relative acceleration, and normal relative position, respectively. It is the orbital angular velocity of the reference body. It is the gravitational constant of the central celestial body. It is the semi-major axis of the reference body orbit.

[0129] Step S32: Solve the differential equations of relative motion to obtain the position and velocity changes of the target body relative to the reference body;

[0130] Define relative state vector Its analytical solution can be written as:

[0131]

[0132] in, It is the initial relative state of the target body with respect to the reference body. Here is the CW state transition matrix:

[0133]

[0134] This matrix allows direct determination of the initial state of the target body relative to the reference body. Calculate the relative position and velocity of the target body with respect to the reference body at any time t.

[0135] Step S33: Introduce an atmospheric density model to correct the calculation results based on step S32, thereby optimizing the calculation results of the relative relationship of spatial targets.

[0136] In step S33, the difference in tangential drag acceleration between the target body and the reference body is added as a non-secondary term to the tangential equation of the relative motion differential equation, and the relative motion differential equation is solved again to obtain a corrected solution, thus completing the optimization of the calculation results of the spatial target relative relationship.

[0137] The drag acceleration of spacecraft i is Where i=R represents the reference body and i=T represents the target body; ballistic coefficients , Indicates the drag coefficient. This represents the cross-sectional area of ​​a spacecraft facing the wind. Let represent the spacecraft mass; ρ be the atmospheric density (calculated using models such as NRLMSISE-00); and v be an approximate value for the orbital velocity. The tangential drag acceleration difference is approximated as a constant and expressed as:

[0138]

[0139] The difference in tangential resistance acceleration As a non-homogeneous term added to the tangential equation, it can be expressed as:

[0140]

[0141] in, .

[0142] The corrected solution is obtained by resolving the relative motion differential equations, and the final optimized relative state is:

[0143]

[0144] in, It is the correction term for the analytical solution of the CW equation caused by atmospheric drag perturbation, and its core component is This describes the tangential position drift caused by the drag difference. From the second drift term This correction significantly improves the accuracy of long-term relative trajectory prediction for low-Earth orbit targets.

[0145] Step S4, Dynamic Profile Visualization: Integrating the high-precision orbit calculation results from Step S2 with the relative relationship calculation results from Step S3, a comprehensive dynamic profile of the analysis object group is synchronously displayed in a three-dimensional scene.

[0146] In this invention, the high-precision trajectory calculation results of step S2 and the relative relationship calculation results of step S3 are integrated to synchronously display the absolute motion trajectory, relative distance, relative speed and orbital plane relationship of the space target in a three-dimensional scene, forming a comprehensive dynamic profile.

[0147] The 3D profile display enables multi-dimensional comprehensive analysis and presentation of dynamically changing spatial targets. This invention introduces Tianditu image map, place name annotation service, and elevation model to construct the analysis environment. It is implemented based on the open-source Cesium component and supported by a high-performance discrete graphics card.

[0148] Step S4 specifically includes:

[0149] Step S41: Initialize the 3D profile simulator and add spatial targets;

[0150] Initializing the 3D profile simulator includes creating the simulation environment, setting the coordinate system, setting the start time, setting the camera position, and creating distance display labels.

[0151] Adding space targets includes: creating spacecraft objects, loading 3D models, setting initial orbital parameters, and storing them in the satellite list. The initial orbital parameters include the semi-major axis, eccentricity, inclination, right ascension of the ascending node, argument of perigee, and true anomaly.

[0152] Step S42: Obtain the positions of the two space targets based on the accurate orbit calculation results of the space targets, and calculate the Euclidean distance between the two space targets;

[0153] Based on the precise orbit calculation results of the space target in step S1, obtain the position vectors of the two spacecraft, and calculate the Euclidean distance between the two spacecraft according to the following formula:

[0154] Euclidean distance =

[0155] in, This indicates the expression for finding the square root.

[0156] Step S43: Obtain the relative position and relative velocity of the two space targets based on the calculation results of the relative relationship between the two space targets, and calculate the orbital plane relationship between the two space targets;

[0157] Specifically, it includes the following steps:

[0158] Step S431: Based on the calculation results of step S3, obtain the spacecraft's position vector and velocity vector;

[0159] Step S432: Calculate the orbital inclination and right ascension of the ascending node of the two spacecraft based on their position and velocity vectors;

[0160] Step S44: Display the spacecraft motion profile on the three-dimensional model based on the calculation results of Step S42 and Step S43.

[0161] The pseudocode for displaying a 3D profile is as follows:

[0162] Initialize the simulator:

[0163] Create a simulation environment;

[0164] Set the J2000 coordinate system;

[0165] Set the start time;

[0166] Set the camera position;

[0167] Create distance display labels;

[0168] Add spacecraft (name, model, orbital parameters):

[0169] Create a spacecraft object;

[0170] Load 3D model;

[0171] Set the initial orbit (semi-major axis, eccentricity, inclination, right ascension of the ascending node, argument of perihelion, true anomaly).

[0172] Stored in the satellite list;

[0173] Calculate the Euclidean distance (spacecraft 1, spacecraft 2): Obtain the positions of the two spacecraft;

[0174] Euclidean distance = ;

[0175] Calculate orbital plane relationship (spacecraft 1, spacecraft 2): Obtain the current position of the spacecraft;

[0176] FUNCTION

[0177] / / Input: Spacecraft position vector r and velocity vector v

[0178] / / Output: Orbital plane parameters (orbital inclination i, right ascension of ascending node) )

[0179] Calculate_orbit_plane(r,v): / / Calculates the orbital angular momentum vector

[0180] h=cross_product(r,v); / / Calculate the track inclination angle

[0181] i=arccos(hz / magnitude(h)); / / Calculate the right ascension of the ascending node

[0182] N=cross_product((0,0,1),h);

[0183] if Ny>=0: arccos(Nx / magnitude(N));

[0184] else: arccos(Nx / magnitude(N)); / / Returns the orbital plane parameters of the two spacecraft respectively.

[0185] RETURN(i, );

[0186] In this invention, the dynamic profile analysis method for space targets solves the problem of three-dimensional profile analysis of the motion state of multiple space targets during the execution of important missions in a space environment through three steps: accurate trajectory calculation, relative relationship calculation, and three-dimensional profile display. The accurate trajectory calculation of space targets comprehensively considers the computational accuracy and efficiency in engineering scenarios. Using the solution to the two-body problem as the initial value, perturbation accelerations output from the Earth's non-spherical gravitational field model EGM2008, the atmospheric drag model NRLMSISE-00, and the solar / lunar gravitational model can be superimposed. The RKF78 integrator is used to solve the orbital motion equations. A closed-loop combination is established between the local error control mechanism of the RKF78 integrator and the global confidence interval judgment criterion based on Monte Carlo simulation. This dynamically and inversely adjusts the internal error tolerance and step size strategy of the RKF78, dynamically optimizing computational efficiency while ensuring computational accuracy. This method is suitable for orbital calculations requiring both high accuracy and high efficiency in space target profile analysis scenarios. The relative relationship calculation uses the RTN coordinate system and describes the linear approximation of the relative motion of the two spacecraft using the CW equation. Error elimination methods are proposed to address major sources of error in the calculation process, such as atmospheric drag. The 3D profile display incorporates Tianditu imagery, place name annotation services, and an elevation model to construct the analysis environment. Based on the open-source Cesium component, it analyzes the relative motion relationships of spatial targets, such as relative distance, relative velocity, and orbital plane relationships between spacecraft. A Web-based 3D model is used to display the spacecraft motion profile, providing a hierarchical and intuitive display of complex multi-target spatial relationships.

[0187] According to one aspect of the present invention, an electronic device is provided, comprising: one or more processors, one or more memories, and one or more computer programs; wherein the processor is connected to the memory, and the one or more computer programs are stored in the memory; when the electronic device is running, the processor executes the one or more computer programs stored in the memory to cause the electronic device to perform a profile analysis method for dynamic changes of a space target as described in any of the above technical solutions.

[0188] The processor can be a central processing unit (CPU), but it can also be other general-purpose processors, digital signal processors (DSPs), application-specific integrated circuits (ASICs), field-programmable gate arrays (FPGAs), or other programmable logic devices, discrete gate or transistor logic devices, discrete hardware components, etc. A general-purpose processor can be a microprocessor or any conventional processor.

[0189] According to one aspect of the present invention, a computer-readable storage medium is provided for storing computer instructions, which, when executed by a processor, implement a profile analysis method for dynamically changing space targets as described in any of the above technical solutions.

[0190] Computer-readable storage media can include any medium capable of storing or transmitting information. Examples of computer-readable storage media include electronic circuits, semiconductor memory devices, ROM, flash memory, erasable ROM (EROM), floppy disks, CD-ROMs, optical disks, hard disks, fiber optic media, radio frequency (RF) links, and so on. Code segments can be downloaded via computer networks such as the Internet and intranets.

[0191] This invention relates to a method, electronic device, and storage medium for the dynamic profile analysis of space targets. The method includes the following steps: Step S1, Object group selection: Select two space targets as the analysis object group; Step S2, High-precision orbit calculation: Considering various perturbation factors, and combining error tolerance and step size dynamic control strategies, perform high-precision orbit numerical integration calculations on the analysis object group; The error tolerance and step size dynamic control strategy combines the local error control mechanism of the RKF78 integrator with the global confidence interval judgment standard based on Monte Carlo simulation in a closed loop, dynamically and inversely controlling the internal error tolerance and step size strategy of the RKF78; Step S3, Relative relationship calculation: Based on the high-precision orbit calculation results of Step S2, establish a relative motion coordinate system with one space target in the high-precision orbit calculation as the reference body and the other space target as the target body, and solve the relative motion differential equations of the target body relative to the reference body; Step S4, Dynamic profile visualization: Integrate the high-precision orbit calculation results of Step S2 and the relative relationship calculation results of Step S3, and synchronously display the comprehensive dynamic profile of the analysis object group in a three-dimensional scene. This invention solves the problem of three-dimensional profile analysis of the motion state of multiple targets in space during the execution of important missions in a space environment through three steps: accurate trajectory calculation of space targets, relative relationship calculation, and three-dimensional profile display. It fills the gap in the comprehensive analysis capabilities of existing technologies, provides more comprehensive and efficient technical support for complex aerospace missions, and has become a key problem that urgently needs to be solved in the field of aerospace simulation. It also lays an important practical foundation and application value for the proposal of this invention.

[0192] Furthermore, it should be noted that the present invention can be provided as a method, apparatus, or computer program product. Therefore, embodiments of the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Moreover, embodiments of the present invention can take the form of a computer program product implemented on one or more computer-usable storage media containing computer-usable program code.

[0193] Embodiments of the present invention are described with reference to flowchart illustrations and / or block diagrams of methods, terminal devices (systems), and computer program products according to embodiments of the invention. It will be understood that each block of the flowchart illustrations and / or block diagrams, and combinations of blocks in the flowchart illustrations and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, embedded processor, or other programmable data processing terminal device to produce a machine, such that the instructions, which execute via the processor of the computer or other programmable data processing terminal device, generate instructions for implementing the flowchart illustrations. Figure 1 One or more processes and / or boxes Figure 1 A device that provides the functions specified in one or more boxes.

[0194] These computer program instructions may also be stored in a computer-readable storage medium that can direct a computer or other programmable data processing terminal device to operate in a particular manner, such that the instructions stored in the computer-readable storage medium produce an article of manufacture including instruction means, which are implemented in a process Figure 1 One or more processes and / or boxes Figure 1 The functions specified in one or more boxes. These computer program instructions may also be loaded onto a computer or other programmable data processing terminal equipment to cause a series of operational steps to be performed on the computer or other programmable terminal equipment to produce a computer-implemented process, thereby providing instructions that execute on the computer or other programmable terminal equipment for implementing the process. Figure 1 One or more processes and / or boxes Figure 1 The steps of the function specified in one or more boxes.

[0195] It should also be noted that, in this document, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or terminal device that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or terminal device. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or terminal device that includes said element.

[0196] Finally, it should be noted that the above description represents a preferred embodiment of the present invention. It should be pointed out that although preferred embodiments have been described, those skilled in the art, once they understand the basic inventive concept of the present invention, can make various improvements and modifications without departing from the principles described herein. These improvements and modifications should also be considered within the scope of protection of the present invention. Therefore, the appended claims are intended to be interpreted as including both the preferred embodiments and all changes and modifications falling within the scope of the embodiments of the present invention.

Claims

1. A method for cross-sectional analysis of dynamic changes in space targets, characterized in that, Includes the following steps: Step S1, Object Group Selection: Select two spatial targets as the analysis object group; Step S2, High-precision orbit calculation: Considering various perturbation factors, high-precision numerical integration calculation of the analysis object group is performed by combining error tolerance and step size dynamic control strategy; the error tolerance and step size dynamic control strategy combines the local error control mechanism of RKF78 integrator with the global confidence interval judgment standard based on Monte Carlo simulation in a closed loop, and dynamically controls the internal error tolerance and step size strategy of RKF78 in reverse. Step S3, Relative Relationship Calculation: Based on the high-precision orbit calculation results in Step S2, establish a relative motion coordinate system with one spatial target in the high-precision orbit calculation as the reference body and the other spatial target as the target body, and solve the relative motion differential equation of the target body relative to the reference body; Step S4, Dynamic Profile Visualization: Integrating the high-precision orbit calculation results from Step S2 with the relative relationship calculation results from Step S3, a comprehensive dynamic profile of the analysis object group is synchronously displayed in a three-dimensional scene.

2. The method for cross-sectional analysis of dynamic changes of space targets according to claim 1, characterized in that, Step S2 specifically includes: Step S21: Obtain calculation data, including ephemeris data, atmospheric density data, and gravitational parameters of the sun and moon; Step S22: Establish a precise orbit calculation model for the space target; The precise orbit calculation model for the space target is a mechanical model that uses an inertial coordinate system, takes the orbital calculation results of the two-body model as initial values, and superimposes perturbation acceleration. Factors affecting the perturbation acceleration include atmospheric disturbances, the Earth's non-spherical gravitational field, gravitational perturbations from the Sun and Moon, the effects of the solar wind and the Earth's magnetic field, and the satellite's own orbital perturbation force. The orbital calculation results of the two-body model are obtained through the ephemeris data. Step S23: Solve the orbital motion equations using an RKF78 integrator. Combine Monte Carlo simulation analysis with dynamic reverse adjustment of the RKF78 integrator's internal error tolerance and step size strategy within the global confidence interval to optimize computational efficiency and perform accurate orbit calculations for the space target.

3. The method for cross-sectional analysis of dynamic changes of space targets according to claim 2, characterized in that, Step S23 specifically includes: Step S231: Construct the motion state vector of the spatial target and its differential equation; Step S232: Set the initialization conditions for accurate orbit calculation. The initialization conditions include the initial time point, the initial solution, and the error tolerance. The initial solution is the orbit calculation result of the two-body model. Step S233: Calculate the approximate solutions with 7th-order and 8th-order accuracy within each time step; Step S234: Estimate the local truncation error based on the difference between the two approximate solutions calculated in step S233; Step S235: Adjust the time step based on the local truncation error; Step S236: Advance all time steps sequentially according to the adaptive step size to complete the spatial target orbit calculation at the target time; Step S237, Repeated Iteration: Analyze the contribution of perturbation error through Monte Carlo simulation, use global accuracy as the criterion, dynamically adjust the error tolerance of the integrator in reverse, and execute steps S232 to S236 until the calculation result meets the preset information interval accuracy requirements.

4. The method for cross-sectional analysis of dynamic changes of space targets according to claim 3, characterized in that, Step S234 specifically includes: Calculate the difference between two approximate solutions This is used to estimate the local truncation error at the current step size. , is represented as: in, and They are respectively At each step, approximate solutions with 7th and 8th order precision are obtained through calculation. Step S235 specifically includes: Define scalar error Given an error tolerance τ, calculate the new step size using the following formula. : in, For safety reasons, and This is the step size variation limiting factor; like ≤τ, accept and with a new step size Proceed to the next time step to continue integration; otherwise, reject the result of the current time step and proceed with the new step size. Recalculate the results at the current time step.

5. The method for cross-sectional analysis of dynamic changes of space targets according to claim 4, characterized in that, In step S237, the error tolerance of the reverse dynamic adjustment integrator specifically includes: The initial trajectory error and key model parameter errors are modeled as random vectors. Draw N independent samples from a random vector. ; For each independent sample drawn This allows for the correction of initial conditions and model parameters, using the current error tolerance. The RKF78 integrator is used for orbit prediction to obtain the position sample at the target time. ; Calculate the mean of N location samples. The sample covariance matrix S is expressed as: Using a 1σ confidence level, the confidence interval for the sphere radius is approximately calculated using the quantiles of the sample error modulus. , is represented as: according to If the accuracy is less than the preset confidence interval, the calculation result is considered to meet the preset confidence interval accuracy requirement, and efficiency is improved by increasing the error tolerance; if If the accuracy exceeds the preset confidence level, the calculation result is considered not to meet the preset confidence interval accuracy requirements, and the integration requirements should be tightened to reduce the error tolerance.

6. The method for cross-sectional analysis of dynamic changes of space targets according to claim 1, characterized in that, Step S3 specifically includes: Step S31: Select the RTN coordinate system, describe the linear approximate relationship of the relative motion of the two space targets through the CW equation, and construct the differential equation of relative motion; Step S32: Solve the relative motion differential equation to obtain the position and velocity changes of the target body relative to the reference body; Step S33: Introduce an atmospheric density model to correct the calculation results based on step S32, thereby optimizing the calculation results of the relative relationship of spatial targets.

7. The method for cross-sectional analysis of dynamic changes of space targets according to claim 6, characterized in that, Step S33 specifically includes: The difference in tangential drag acceleration between the target body and the reference body is added as a non-secondary term to the tangential equation of the relative motion differential equation, and the relative motion differential equation is solved again to obtain the corrected solution, thus completing the optimization of the calculation results of the relative relationship of the space target. The difference in tangential drag acceleration between the target body and the reference body is expressed as: In the formula, ρ is the atmospheric density. and Let represent the ballistic coefficients of the target body and the reference body, respectively. The ballistic coefficient of spacecraft i is expressed as: , Indicates the drag coefficient. This represents the cross-sectional area of ​​a spacecraft facing the wind. The spacecraft mass is represented by v; v is an approximate value for the orbital velocity. The tangential equation of the relative motion differential equation with the addition of non-secondary terms is expressed as: In the formula, Indicates tangential relative acceleration. It is the orbital angular velocity of the reference body. It is the gravitational constant of the central celestial body. It is the semi-major axis of the reference body orbit. It represents the difference in acceleration of external forces in the tangential direction.

8. The method for cross-sectional analysis of dynamic changes of space targets according to claim 6, characterized in that, Step S4 specifically includes: Step S41: Initialize the 3D profile simulator and add spatial targets; Step S42: Obtain the positions of the two space targets based on the accurate orbit calculation results of the space targets, and calculate the Euclidean distance between the two space targets; Step S43: Obtain the relative position and relative velocity of the two space targets based on the calculation results of the relative relationship between the two space targets, and calculate the orbital plane relationship between the two space targets; Step S44: Display the spacecraft motion profile on the three-dimensional model based on the calculation results of Step S42 and Step S43. The spacecraft motion profile includes the absolute motion trajectory, relative distance, relative velocity, and relative relationship of the orbital plane of the space target.

9. An electronic device, characterized in that, include: One or more processors, one or more memories, and one or more computer programs; wherein the processor is connected to the memory, and the one or more computer programs are stored in the memory, and when the electronic device is running, the processor executes the one or more computer programs stored in the memory to cause the electronic device to perform the profile analysis method for dynamic changes of space targets as described in any one of claims 1 to 8.

10. A computer-readable storage medium, characterized in that, Used to store computer instructions, which, when executed by a processor, implement a profile analysis method for dynamic changes of space targets as described in any one of claims 1 to 8.

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