15 results about "Libration" patented technology

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 discloses a lunar balance dynamic parameter high-precision resolving method and system combining LRR and LLR technologies, and belongs to the field of deep space exploration and celestial body dynamics, and the method comprises the steps: constructing an LRR and LLR combined observation data set, and carrying out the preprocessing; constructing an LRR and LLR combined moon measurement model, and deducing a partial derivative of the model for a moon rotation parameter and a tidal LOVE number to obtain a combined partial derivative matrix; iterative solution is carried out through nonlinear least square to obtain dynamic parameters of the moon balance; and evaluating a resolving result and optimizing a resolving strategy, and optimizing a resolving scheme by dynamically adjusting the weight, improving a tide model and eliminating abnormal values. According to the method, the parameter calculation error is reduced, the technical advantages of two types of data are effectively integrated, the problem of insufficient observation continuity or precision of a single data source is solved, and reliable parameter support is provided for moon rotation and internal structure research.

The invention discloses a construction method and system for a low-thrust transfer orbit family between earth-moon translation point orbits, and belongs to the technical field of aerospace. The method comprises the following steps: firstly, solving a time optimal transfer orbit based on a continuous low-thrust spacecraft motion equation; then taking the orbit as a seed, performing forward and reverse phase continuation along a departure orbit and an arrival orbit by adopting a self-adaptive step length numerical continuation method, and generating a time optimal transfer orbit family covering different phases; secondly, mapping the time optimal orbit family into an energy optimal orbit family based on control law consistency, and expanding the transfer time range of the energy optimal orbit family through time continuation; and finally, gradually reducing homotopy parameters by using a homotopy method, detecting and processing thrust switching discontinuity in combination with a switching function, and converting the energy optimal orbit family into a fuel optimal transfer orbit family. According to the method, a continuous and complete orbit family is efficiently generated from a single solution, and systematicness and engineering applicability of orbit design are improved.

The invention discloses an earth-moon L1 point track crossing behavior prediction method and system based on machine learning. According to the method, firstly, an initial state set is generated based on an earth-moon plane circular restrictive three-body model, and an orbit sample data set is constructed through numerical integration; secondly, a multi-class behavior classification system is established, and sample data is automatically labeled; on the basis, an XGBoost algorithm is adopted for model training, a category self-adaptive threshold optimization strategy based on a verification set is introduced for the problem of data extreme imbalance, and the overall prediction precision and category fairness of the model are remarkably improved by independently optimizing an optimal decision threshold for each category and applying the optimal decision threshold to normalized probabilistic decision. According to the method, the technical blank of spacecraft behavior prediction in the earth-moon translation point area is filled, and an effective means is provided for space situation awareness.

A method for enhancing lunar regional navigation based on a libration point navigation constellation, belonging to the field of navigation technology. The navigation constellation in the present invention is composed of different numbers of libration point navigation satellites, and some or all of the orbits near the L1, L2, L4, and L5 libration points are selected as the orbits for the navigation constellation to operate. Different numbers of lunar surface base stations are added to this navigation constellation to jointly improve the observation geometry of the constellation with the libration point satellites, enhance the ranging signal, improve the autonomous orbit determination accuracy of the entire navigation constellation, and solve the problem of insufficient positioning accuracy of the lunar surface base station in the Earth-Moon rendezvous coordinate system. Its main purpose is to provide a high-precision navigation function for the lunar region. The present invention aims to be able to provide lunar surface base stations with high-precision positioning and achieve high-precision lunar regional navigation based on the lunar surface base stations, providing a new type of navigation service for functions such as resource exploitation of moving objects on the lunar surface and lunar surface patrol in the future.

The present invention discloses a near-lunar space wide-area surveillance system, comprising three observation satellites located in NRHO orbits at the Earth-Moon libration point L1, L2, and DRO orbits, respectively. The observation satellites carry a wide-field-of-view camera payload and a precision tracking camera payload. The wide-field-of-view camera payload is used for scanning, searching, and detecting wide-area airspace in near-lunar space, while the precision tracking camera payload is used for tracking and monitoring space targets. By leveraging the characteristics of the NRHO and DRO orbits, the system, using only three observation satellites, can achieve short-period, full-coverage detection of near-lunar space and tracking and monitoring of near-lunar space targets, while maintaining a low long-term presence cost.

The invention relates to a characterization parameter analysis method for an earth-moon collinear translation point orbit. 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; carrying out complex transformation and regular transformation on the linearized model transformation to realize decoupling of a hyperbolic unstable direction and a central direction of the central manifold, and constructing a mapping relation between CRTBP coordinates and characterization parameters to carry out parameter characterization; analyzing the two motion characteristics in the hyperbolic unstable direction and the central direction and the corresponding characterization parameters, determining the orbit of the target spacecraft through the phase difference of the angular variables in the characterization parameters, and obtaining the orbit transfer information of the target spacecraft through the change condition of the hyperbolic motion components in the characterization parameters. By adopting the method, characterization parameter analysis can be realized.

The application discloses a lunar orbit and libration parameter solving and forecasting method and system, and the method comprises the following steps: considering the interaction among the mass point gravity model based on the post-Newtonian framework, the moon figure, other planet mass points in the solar system and the influence of the earth figure, a high-precision lunar orbit dynamics model is established under the general relativity framework; considering the lunar moment of inertia and the lunar torque, a double-layer lunar libration attitude dynamics model is established; when the initial conditions of integration meet the preset accuracy, an adaptive step integrator and an adaptive step integrator with time delay effect are used to respectively forecast the orbit parameters and the libration parameters, and initial products of the orbit and the libration parameters with time sequence labels are obtained; when the initial conditions of integration do not meet the preset accuracy, an automatic differentiation-based nonlinear solver is used to process historical observation data, and the optimal solution of the to-be-estimated parameters in the initial conditions is obtained, and re-forecasting is performed.

The present invention discloses a method for designing an Earth-Moon space navigation constellation based on a special long-period orbit, comprising the following steps: step 1, designing Halo orbits of Earth-Moon L1 and L2 libration points in a dynamic model of a circular restricted three-body problem; step 2, designing a special long-period orbit using the invariant manifold of the Halo orbit, on which a spacecraft performs periodic motion; step 3, forming a navigation satellite constellation by deploying multiple navigation satellites on the special long-period orbit according to configuration parameters; step 4, setting a series of sample points in an ecliptic plane for simulating and analyzing the navigation performance of the constellation; at the sample points, first obtaining visible satellites according to a coverage judgment condition; then performing a star selection operation according to an optimal GDOP method, i.e., selecting a four-satellite combination with the smallest GDOP value for positioning; and calculating the minimum GDOP of each sample point at each time during the simulation process; and step 5, establishing a navigation constellation optimization model under constraint conditions, using the configuration parameters in step 3 as input, optimizing each configuration parameter of the constellation using an optimization algorithm to minimize the optimization index, and using the optimized parameters as the configuration parameters for the navigation constellation design. The present invention can solve the problem that existing navigation systems cannot meet the demand for high-precision navigation and positioning in the entire Earth-Moon space.

Embodiments of the present application disclose a transfer orbit design method from GTO to the Sun-Earth collinear libration point, a device and a medium; the method comprises: calculating and obtaining an invariant manifold connected with the Halo orbit around the libration point according to the spacecraft dynamics equation and the potential function in the Sun-Earth barycenter coordinate system; taking the geosynchronous transfer orbit as the earth parking orbit to construct the constraint variable, taking the Halo orbit as the target orbit to construct the control variable, and correcting the control variable by using the least square differential correction method; based on the set constraint condition and performance index, the optimal HOI point corresponding to the optimal performance transfer orbit is selected from the transfer orbit meeting the constraint condition; according to the optimal HOI point, the relationship between the control variable and the parking orbit constraint variable is analyzed, and the differential correction initial value expression for determining the control variable initial value of the multi-constraint orbit design is constructed; the transfer orbit meeting multiple constraints is obtained by using the multi-stage differential correction method according to the differential correction initial value expression.

The present invention provides a rapid design method and system for the transfer orbit of the Sun-Earth L5 libration point. Based on the heliocentric two-body orbit analytical model, the number of transfer loops, transfer time, and eccentricity of the transfer orbit are set, and the semi-major axis of the transfer orbit is calculated. According to the semi-major axis, eccentricity of the transfer orbit in step 1 and the orbital parameters of the Earth, the inclination of the transfer orbit is set, and the Earth zero-radius sphere of influence model is used to calculate the magnitude and direction of the hyperbolic excess velocity of the detector. It is set that the detector applies a braking impulse once when arriving at the L5 point, and the magnitude and direction of the impulse are calculated by using the transfer orbit parameters in steps 1-2. Taking the heliocentric two-body transfer orbit parameters calculated by the above steps as the initial value guess, the heliocentric multi-body orbit model is used to calculate the high-precision transfer orbit through iterative optimization. The method in the present invention realizes the trade-off among orbit accuracy, computational amount, and design efficiency, effectively reduces the number of iterations, reduces the computational amount, and ensures the robustness of numerical iteration.

The present invention discloses a navigation and positioning system for a user aircraft in the Earth-Moon space, which relates to the technical field of satellite navigation. The system includes navigation satellites and a user aircraft in the Earth-Moon space. Each navigation satellite is used to obtain the optimal estimates of the dynamic navigation ephemeris parameters and the correction amount of the inter-satellite measurement system deviation parameters according to the satellite attitude of the navigation satellite, the antenna phase center, the inter-satellite ranging measurements sent by the target navigation satellite, and the dynamic equation. The user aircraft in the Earth-Moon space obtains the orbit determination observation equation of the user aircraft according to the dynamic navigation ephemeris parameters, the optimal estimates of the correction amount of the inter-satellite measurement system deviation parameters, the dynamic equation of the navigation satellite, and the ranging information; and is used to obtain the state transition equation of the user aircraft according to the dynamic equation of the user aircraft in the Earth-Moon space. According to the state transition equation and the orbit determination observation equation of the user aircraft, the navigation and positioning of the user aircraft in the Earth-Moon space are realized. The present invention can improve the navigation and positioning accuracy and is more suitable for libration point navigation satellites.

The application discloses a bifurcation construction method, system, equipment, medium and program of a clover orbit family in a geolunar three-body system, and belongs to the technical field of aerospace. The method comprises the following steps: constructing a geolunar circular restricted three-body dynamics model according to a geolunar three-body system; obtaining a halo orbit family based on the geolunar circular restricted three-body dynamics model; positioning a halo orbit family bifurcation point, and constructing a transition orbit; obtaining a first clover orbit based on the transition orbit, and extending to obtain a clover orbit family. The application can obtain complex periodic orbit families derived near the libration points L1 and L2 in a restricted three-body system, and quickly constructs the periodic orbit family.

The application discloses a kind of coordinate determination methods of lunar observation remote sensing image, steps are as follows: the three-dimensional coordinates of each pixel point in lunar image are established;Determine the satellite view angle libration based on the lunar observation position of satellite;The azimuth between any two points on the moon is determined according to the geographical coordinates of the two points, the projection coordinates of two latitude and longitude coordinate points on the image plane and the azimuth thereof are determined;The angle relationship between the scanning direction when satellite sampling and the lunar fixed coordinate system is determined;The geographical coordinates of the observed lunar image under the lunar fixed coordinate system are determined.The method of the application can avoid the complex imaging model based on rigorous geometric relationship, and can solve the deficiency that the image matching method has strong dependence on image gray or texture information;Only a small amount of image processing and simple coordinate conversion can establish the accurate mapping relationship between image coordinate and moon surface coordinate.

This invention discloses an improved method for designing orbits near the Earth-Moon triangular translation point. Based on the quasi-periodicity of orbits near the Earth-Moon triangular translation point applicable to space missions, this method obtains terms with different periodic frequencies through spectral analysis and multi-target tracking. Then, based on the analytical solution of the orbit near the Earth-Moon triangular translation point under an ephemeris model, the corresponding coordinate and frequency coefficients are fitted. Using the semi-analytical solution of the orbit enables rapid orbit design, thus avoiding the performance overhead of numerical iteration, and allows for the design of orbits with specified amplitudes and phases.