43 results about "Lunar orbit" patented technology

The invention provides a lunar south pole multi-source fusion positioning method and system fusing a lunar orbit satellite, inertial navigation and wireless communication, and the method comprises the steps: building a multi-source fusion positioning measurement model through employing a lunar orbit satellite pseudo-range measurement value and a wireless communication time difference of arrival (TDOA) measurement value, and solving the position information through a least square method; the system error of the INS system is used as an error state, and the difference value between the INS system and the resolving position of the least square method is used as a measurement error state; and constructing a loose coupling model based on ESKF (Error State Kalman Filter) to calculate the error state and correct the calculation position of the INS, and updating the current INS internal information to obtain the accurate position of the lunar probe vehicle. According to the invention, the problems of limited coverage range and low positioning precision caused by insufficient lunar orbit satellites are effectively solved; the influence of INS error accumulation on the positioning accuracy is reduced, and finally the accurate position of the lunar probe vehicle is obtained; and the overall positioning precision and reliability of the system are obviously improved.

The invention provides a lunar conduction remote satellite formation system and a cooperative work method, the system is deployed in a lunar orbit, and through cooperative work among satellites in the system, the satellite formation system is matched with a ground Ka measurement and control station to jointly provide communication, navigation and remote sensing services for lunar users. The lunar conduction remote satellite formation system comprises a master satellite and slave satellites, wherein the master satellite comprises a satellite-ground high-speed measurement and control unit, an inter-satellite link unit, a communication load, a navigation load, an ultra-stable time frequency unit and a comprehensive electronic unit; the slave satellite comprises a measurement and control unit, an inter-satellite link unit, a communication load, a navigation load, a remote sensing load and a comprehensive electronic unit. Through the arrangement of the lunar orbit, the configuration of the communication navigation remote sensing function and the cooperation of satellites in the system, comprehensive communication and remote sensing services are provided for lunar users in the lunar space where signals of an earth navigation system cannot cover and signals of a ground measurement and control system are limited to cover.

The invention discloses a lunar principal element inversion method based on deep learning and self-supervision. The lunar principal element inversion method comprises the following steps: preprocessing a lunar orbiter remote sensing spectral image; constructing a first data set and a second data set based on the preprocessed remote sensing spectral image; performing one-dimensional convolution auto-encoder self-supervised learning through the first data set to obtain a feature extractor; performing feature extraction on the second data set by using a feature extractor, and performing regression training in combination with a support vector machine to obtain a lunar principal element inversion model; and inputting the full-moon hyperspectral data into the lunar major element inversion model, and inverting and drawing a lunar major element abundance distribution diagram. According to the method, deep learning and self-supervised learning are fused, features can be directly learned from the lunar orbit remote sensing spectral image, dependence on data labels and the difference between laboratory data and remote sensing data are effectively overcome, and the precision and efficiency of lunar surface principal element inversion are remarkably improved.

Systems and methods for predicting decrease in average sea surface temperature (SST) are disclosed. The method includes setting a value of a first Boolean flag when at least one of an eclipse season start date and an eclipse season end date in a particular calendar year occurs in a period of thirty days preceding a timestamp for a particular partition in a plurality of partitions segmenting the particular calendar year. The method includes inputting to a trained machine learning model the first Boolean flag. The trained machine learning model is trained using a plurality of time series data using lunar orbit characteristics. The method includes predicting, using the trained machine learning model, a probability of decrease in the average SST for a predefined time interval in the particular calendar year. The method includes invoking a weather forecasting system to generate weather forecast in dependence on the predicted probability.

The invention discloses a thermal control method and device for improving the performance of a lunar orbit laser corner reflector, and belongs to the technical field of space flight and aviation. The method comprises the following steps: in a non-test stage, analyzing illuminated and thermal environment data of the corner reflector in a lunar orbit, and designing heat dissipation and heat insulation measures, so that the corner reflector is in a low-temperature storage stage; in the test preparation stage, heating belts are started for preheating, temperature measurement data of all areas of the corner reflector are collected, and the duty ratio of all the heating belts is independently adjusted in combination with a novel PID temperature control algorithm and an external heat flow prediction model; in the test process, the heating belt is started, all areas of the corner reflector are monitored and adjusted in real time, and when temperature deviation occurs, the duty ratio of the heating belt is adjusted through a novel PID temperature control algorithm and an external heat flow prediction model; and after the test is completed, the corner reflector enters the low-temperature storage stage again. According to the invention, high precision and high reliability of the laser corner reflector on the lunar orbit can be ensured, and the success probability of laser ranging on the lunar orbit spacecraft corner reflector by the ground is improved.

This invention is embodied in a system and device for transporting assets (small satellites, commodities, or other payloads) in space. The preferred system comprises a scaffold of relay stations strategically orbiting one or more celestial bodies. The orbits are complimentary to each other so that the relay stations can transfer mass from one station to another. In this way, the system can support different logistical networks connecting the space between orbits such as Low Earth orbit (LEO), geostationary orbit (GEO), and Low Lunar orbit (LLO), and even between Earth and Mars.

The present invention discloses a calculation method for the requirements of satellite determination and orbit determination accuracy for inter-satellite signal acquisition in a lunar orbit formation, belonging to the technical field of deep space exploration. The method includes measuring the antenna pattern, transmitted signal power, and receiving sensitivity of the main satellite, measuring the antenna pattern, receiving sensitivity, and transmitted EIRP value of the slave satellite, measuring the installation error and antenna mechanism deformation error of the high-gain directional antenna of the main satellite, measuring the attitude control accuracy of the main satellite, and calculating the position error requirement of the slave satellite relative to the main satellite according to the total error requirement of the main satellite antenna pointing, the installation error of the antenna, the antenna mechanism deformation error, and the beam pointing error caused by the attitude control accuracy of the main satellite; determining the requirements for the determination and orbit determination accuracy of the main satellite and the slave satellite based on the position error. The present invention can put forward the requirements for determination and orbit determination to the ground TT&C system in advance, or select an appropriate inter-satellite distance according to the determination and orbit determination accuracy that can be provided by the ground, improving the success probability of establishing an inter-satellite link.

The application discloses a heat control method and device for improving performance of a lunar orbit laser corner reflector, and belongs to the technical field of aerospace. The method comprises the following steps: in a non-test stage, analyzing the illumination and thermal environment data of the corner reflector in a lunar orbit, and designing heat dissipation and heat insulation measures so that the corner reflector is in a low-temperature storage stage; in a preparation test stage, starting a heating belt for preheating, collecting temperature data of each region of the corner reflector, combining a novel PID temperature control algorithm and an external heat flow prediction model, and independently adjusting the duty cycle of each heating belt; in a test process, starting the heating belt, and performing real-time monitoring and adjustment on each region of the corner reflector; when the temperature deviates, the duty cycle of the heating belt is adjusted through the novel PID temperature control algorithm and the external heat flow prediction model; and after the test is completed, the corner reflector reenters the low-temperature storage stage. The application can ensure high precision and high reliability of the laser corner reflector in the lunar orbit, and improves the success probability of laser ranging of the corner reflector on a lunar orbit spacecraft on the ground.

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 relates to the technical field of lunar orbit image analysis, in particular to an annular mountain detection method based on orbit camera images, a training data set is constructed based on the lunar orbit camera images, and the training data set comprises an annular mountain edge thermodynamic diagram label, an offset direction label and a radius label based on multi-interval coding; training a deep learning network by using the data set, and setting a network output layer into three branches which are respectively used for predicting an edge thermodynamic diagram, an offset direction and a radius multi-interval; inputting a to-be-detected image into the network, obtaining a prediction result, screening an effective edge pixel set based on the thermodynamic diagram, and performing direction and radius decoding on the effective edge pixel set to generate an annular mountain center candidate point; the mean shift clustering algorithm is adopted to perform instance separation on the candidate points to obtain a ring-shaped mountain detection result, and edge prediction and multi-interval coding are combined to improve the detection rate and positioning precision of weak-edge, small-scale and adjacent ring-shaped mountains.

Systems and methods for predicting increase in average sea surface temperature (SST) are disclosed. The method includes setting a value of a first Boolean flag when at least one of an eclipse season start date and an eclipse season end date in a particular calendar year occurs in a period of thirty days preceding a timestamp for a particular partition in a plurality of partitions segmenting the particular calendar year. The method includes inputting to a trained machine learning model the first Boolean flag. The trained machine learning model is trained using a plurality of time series data using lunar orbit characteristics. The method includes predicting, using the trained machine learning model, a probability of increase in the average SST for a predefined time interval in the particular calendar year. The method includes invoking a weather forecasting system to generate weather forecast in dependence on the predicted probability.

The invention provides a steady-state measuring device and method for measuring radiation heat flux of an extraterrestrial space, and the device comprises a mounting bottom plate which is made of a thermal insulation material; the first heat flow meter is assembled in the first area of the mounting bottom plate; comprising a first measuring assembly and a second measuring assembly, the second measuring assembly is stacked on the surface of the first measuring assembly, and the diameter of the second measuring assembly is smaller than that of the first measuring assembly; the surface layer of the first measuring assembly is provided with a first compensating plate. The surface layer of the second measuring assembly is provided with a first temperature-sensing wafer. The second heat flow meter is assembled in the second area of the mounting bottom plate; comprising a third measuring assembly and a fourth measuring assembly, the fourth measuring assembly is stacked on the surface of the third measuring assembly, and the diameter of the fourth measuring assembly is smaller than that of the third measuring assembly; the surface layer of the third measuring assembly is provided with a second compensating plate. The surface layer of the fourth measuring assembly is provided with a second temperature-sensitive wafer. The lunar orbit solar radiation, the lunar reflection and the lunar infrared radiation heat flow can be measured at the same time.

The invention discloses a method and system for dynamically reconstructing a launch window for a lunar orbit, and relates to the technical field of lunar exploration launch orbit design. The method comprises the following steps: receiving the initial transfer orbit elements of the lunar orbit, and calculating the orbit elements at the moment of satellite-rocket separation according to the initial transfer orbit elements of the lunar orbit; inputting the orbit elements at the moment of satellite-rocket separation into an accurate dynamics equation, iteratively obtaining the accurate orbit elements of the first lunar orbit at the take-off moment, changing the take-off moment, and iteratively obtaining the accurate orbit elements of multiple lunar orbits; performing polynomial fitting on the bound elements of the accurate orbit elements of the multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and binding the elements of a launch vehicle as the polynomial coefficients corresponding to each orbit, so as to realize dynamic reconstruction of the lunar orbit according to the bound elements of the launch vehicle.

The present invention discloses a dual-pulse co-tangent optimal orbit change method for a large elliptical orbit around the moon. First, based on the known initial state and target orbit state of the lunar orbit, a dual-pulse co-tangent orbit change control strategy based on a two-body nonlinear model is calculated. Through two pulse controls along the tangent direction, the average semi-major axis, average eccentricity and average latitude argument of the large elliptical orbit are adjusted simultaneously. The velocity increment minimum problem is converted into a one-dimensional nonlinear function minimum problem to solve the optimal orbit change strategy, thereby achieving precise transfer between any two elliptical orbits. Then, using the two-body orbit model design as the initial value, the two-body model is corrected by numerically integrating the average orbit elements under the influence of perturbations, thereby obtaining a real orbit change control strategy for the average orbit parameters under the influence of perturbations. This solves the problem that the short-period term introduced by directly using the instantaneous orbit element description turns the problem into a time-varying problem that is difficult to solve, thereby improving the engineering practical value of the algorithm.

The application discloses a method for processing and analyzing whole-process errors of a lunar orbit transfer of a large-ellipse frozen orbit, and belongs to the technical field of deep space exploration orbits. The method comprises the following steps: setting whole-process error terms; adding an orbit entry error to an orbit parameter at an orbit entry point, separately calculating measured orbit errors of each stage to obtain orbit control velocity increments of the corresponding stages under the measured orbit errors; analyzing the influence of the whole-process errors on total velocity increments and judging whether a probe can enter a mission orbit; analyzing the influence of the whole-process errors on measurement and control conditions and judging whether the measurement and control conditions are met; analyzing the influence of the whole-process errors on orbit safety and autonomous orbit departure at the end of a mission and judging whether the probe will be controlled to depart from the orbit, if yes, the whole-process error analysis is ended, and if no, a reserved velocity increment is used for orbit correction. Compared with single analysis of errors of each stage, the application is more systematic and closer to a real situation, and the analysis result can provide a reference for overall and each subsystem design.

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 and device for determining the orbit of a lunar-orbiting satellite based on radio and laser ranging, which relates to the technical field of lunar-orbiting satellite orbit calculation, and includes: acquiring the observation data of radio tracking stations and laser stations; establishing a high-precision dynamic model of the lunar-orbiting satellite; establishing a mathematical model for radio and laser ranging of the lunar-orbiting satellite, processing the observation data according to the observation equation, and obtaining the solution values of the dynamic parameters of the satellite at the initial orbit determination moment; substituting the solution values of the dynamic parameters into the satellite perturbed motion equation to calculate and obtain the orbit determination result of the lunar-orbiting satellite. The present invention uses combined orbit determination by radio ranging and laser ranging to achieve precise orbit determination of the lunar-orbiting satellite. Compared with the traditional orbit determination method, the accuracy and reliability are greatly improved, and the requirements for the number and geometric distribution of ground stations are reduced, thus reducing the measurement and control pressure on existing ground stations.

The invention discloses a full-time reliable lunar-based communication topology architecture, and belongs to the technical field of deep space exploration. The framework comprises a lunar surface base station and a lunar orbit base station, and the lunar surface base station is used for establishing an uninterrupted earth-moon backbone communication link with a ground measurement and control station; the lunar orbit base station is used for establishing a relay link with the ground measurement and control station when the lunar orbit base station is visible to the ground; a lunar orbit-lunar surface link is established between the lunar orbit base station and the lunar surface base station, a lunar orbit inter-satellite link is established between every two visible lunar orbit base stations, and earth-moon communication service and lunar communication service are jointly provided for lunar users. The method does not need a ground measurement and control station for forwarding, is suitable for tasks needing quick response, can dynamically adjust communication resources and paths according to different communication requirements and priorities, and improves the efficiency of earth-moon communication and moon-based communication.

This invention discloses a method for determining the lunar orbit expansion rate using a lunar optical clock, comprising: acquiring the relative frequency deviation between a ground optical clock and a lunar optical clock obtained through continuous measurement via a two-way Earth-Moon time-frequency transfer link; the two-way Earth-Moon time-frequency transfer link is established between a lunar surface station and a ground station for realizing optical clock time-frequency signal comparison and laser ranging, wherein both the lunar surface station and the ground station are equipped with a frequency deviation better than 10. ‑20 This invention utilizes an optical clock of magnitude [missing information]; based on the gravitational redshift theory of general relativity, a physical relationship model is established between the rate of change of relative frequency deviation and the lunar orbit expansion rate; based on this physical relationship model, the lunar orbit expansion rate is calculated from the measured rate of change of relative frequency deviation. This invention achieves direct, continuous, and high-precision measurement of lunar orbit expansion, overcoming the limitations of traditional lunar laser ranging technology, and providing a new generation of measurement methods for the study of Earth-Moon system dynamics and fundamental physics verification.

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

This invention provides a rapid prediction method for lunar orbit rendezvous guidance, achieving high-precision and rapid prediction to meet mission requirements. The prediction method of this invention applies the gravitational field acceleration calculation method used for GPS autonomous orbit determination of Earth orbit satellites to the field of lunar orbit rendezvous. Addressing the problems that the algorithm cannot be directly applied to the design of lunar orbit rendezvous guidance strategies, improvements and upgrades are made one by one, resulting in a high-precision and rapid algorithm suitable for lunar orbit rendezvous flight control missions to meet mission requirements. Specifically, this invention is the first to transplant this algorithm to the field of lunar orbit applications, which is unprecedented. Considering the special characteristics of the lunar gravitational field, where the acceleration measures of various gravitational perturbations are of similar order and cannot be approximated by the J2 term alone, the J22 term perturbation acceleration is added to the algorithm, thereby significantly improving the computational accuracy and numerical stability of the algorithm.

The invention relates to a lunar landing orbit initial value design method, device and equipment based on velocity curved surface parameters. The method comprises the following steps: determining a value range of each variable according to engineering constraints and constants; calculating a first velocity curved surface parameter equation meeting perigee height constraints; calculating a second velocity curved surface parameter equation meeting the perilunar height constraint, carrying out simultaneous solution, and constructing a third velocity curved surface parameter equation of the general earth-moon orbit; performing traversal assignment on each variable according to the variable value range; substituting each assignment into a third velocity curved surface parameter equation for calculation to obtain a tangential velocity; the tangential speed meeting the tolerance is corrected, and the corrected tangential speed is obtained; obtaining a corresponding variable initial value set based on the corrected tangential speed; variable initial values in the variable initial value set are screened through engineering constraints, and variable initial values meeting conditions are obtained; the design precision and reliability of the track can be improved.

The present invention discloses a rapid iterative guidance method for lunar orbit rendezvous and remote guidance, comprising: obtaining an initial orbital state corresponding to an initial moment and an orbital target state corresponding to a terminal moment; establishing a nonlinear equation group of design variables and terminal orbital states; transforming the nonlinear equation group of design variables and terminal orbital states based on the near-circular orbit element representation of the terminal orbital state and the orbital target state; determining iterative initial values ​​and convergence conditions; iteratively solving the nonlinear equation group of the transformed design variables and terminal orbital states based on the determined iterative initial values ​​and convergence conditions until the convergence conditions are met, obtaining the final design variables, and completing the guidance. The method of the present invention has high iterative efficiency and fast calculation speed, is applicable to any number of orbit changes, and the components of each orbit change can be arbitrarily specified. It is particularly applicable to the needs of on-orbit real-time generation of lunar orbit rendezvous and remote guidance orbit change strategies and reconstruction of remote guidance strategies under fault conditions.

The application provides a steady-state measuring device and method for extraterrestrial space radiation heat flow measurement, comprising: a mounting base made of thermal insulation material; a first heat flow meter arranged in a first area of the mounting base; a first measuring component and a second measuring component, the second measuring component being stacked on the surface of the first measuring component, and the diameter of the second measuring component being smaller than that of the first measuring component; the surface layer of the first measuring component being provided with a first compensation sheet, and the surface layer of the second measuring component being provided with a first temperature-sensing disc; a second heat flow meter arranged in a second area of the mounting base; a third measuring component and a fourth measuring component, the fourth measuring component being stacked on the surface of the third measuring component, and the diameter of the fourth measuring component being smaller than that of the third measuring component; the surface layer of the third measuring component being provided with a second compensation sheet, and the surface layer of the fourth measuring component being provided with a second temperature-sensing disc; and the application can simultaneously measure the lunar orbit solar radiation, the lunar reflection and the lunar infrared radiation heat flow.

The method for in-situ production of methane on the moon comprises the following steps: step S1, inputting star soil of a C-type near-earth asteroid and moon water into a carbon gasification reaction system to generate a mixed gas of carbon monoxide and hydrogen; s2, inputting the carbon monoxide and hydrogen mixed gas generated in the step S1 into a methanation reaction system to generate methane and water mixed gas; s3, methane and water in the methane and water mixed gas generated in the step S2 are separated, and the methane is stored; and S4, circulating the unreacted carbon monoxide and hydrogen to the methanation reaction system. According to the method, the complete process of in-situ production of methane in the lunar orbit is achieved, carbon resources in the C-shaped near-earth asteroid star soil are organically combined with the lunar water-ice resource development technology, and the blank of production of methane propellants through lunar in-situ resources in the prior art is filled.

This invention discloses a lunar orbit space solar power station system, belonging to the field of spacecraft structure and energy system technology. Addressing the problems of traditional solar power generation failure in the lunar polar shadow regions and the inability of existing space power station structures to adapt to the lunar orbit environment, a novel system is proposed, comprising: a double-layered box-type central truss made of carbon fiber composite material; 12 thin-film solar subarrays, achieving independent solar tracking via a first conductive rotary joint; 5 laser transmitting antenna subarrays, achieving precise pointing via a second conductive rotary joint; and a rotary connection module. The system adopts a tree-like topological multibody structure, allowing independent attitude adjustment of the solar energy collection and laser transmitting modules, achieving decoupling between solar tracking and laser pointing. This invention offers advantages such as modular on-orbit assembly, lightweight and high rigidity, and flexible vibration suppression, providing continuous wireless energy transmission to the lunar polar regions.

The invention relates to a periodic earth-moon connection transfer system and method based on an earth-moon round trip 3: 1 resonance orbit, and belongs to the technical field of aerospace. The system mainly comprises a near-earth track connection station arranged on a near-earth track, an earth-moon transfer aircraft arranged on an earth-moon 3: 1 resonance track and used for transporting goods between the near-earth track and the connection station, a moon track connection station arranged on a moon track and a standardized goods cabin. The characteristic that the earth-moon 3: 1 resonance orbit is periodically close to the earth and the moon is utilized, long-term in-orbit and low-energy-consumption operation of the transfer aircraft is achieved, and a periodic flight-based earth-moon two-way transportation system is constructed through modular design and the reusable aircraft. According to the system, the earth-moon transportation cost can be greatly reduced, the transportation efficiency and flexibility are improved, and an economical and efficient logistics solution is provided for lunar exploration and development.

The invention provides a lunar surface continuous positioning method and system based on factor graph optimization under the condition that lunar navigation resources are limited. The method comprises the following steps: acquiring pseudo-range observation and Doppler frequency shift observation of each visible lunar orbit satellite and speed information provided by an IMU (Inertial Measurement Unit), and constructing a navigation state vector of a lunar mobile terminal; constructing a factor graph positioning model; combining the local cost functions to construct a global cost function so as to construct an optimization objective function; and iteratively solving and optimizing the objective function to obtain an optimal estimation result. According to the method, the speed deviation output by an inertial navigation system is introduced into a positioning model as a state quantity, and joint estimation and online correction are performed on the inertial navigation speed deviation through satellite pseudo-range and Doppler frequency shift observation, so that the long-term stability and precision of a positioning solution of the lunar mobile terminal are improved; speed information provided by an IMU serves as external input of a Doppler observation model, so that Doppler frequency shift observation directly participates in user position estimation under the condition that explicit extension of user speed state dimensions is not needed.

The invention discloses a high-precision lunar orbit satellite laser ranging echo signal grading rapid search method, and belongs to the technical field of deep space laser ranging. The method comprises the following steps: a ground end laser emission system tracks and points to a target position according to a guide file; transmitting a long-pulse-width laser signal, and if a tracking target point echo is not monitored, performing target point airspace search based on an Archimedes spiral search model; after an echo signal of a tracking target point is received at the current position, Archimedes spiral search is stopped, and the current position serves as a correction original point to continue to track and forecast the trend of an orbit guide file; and if the quality of the echo signal of the tracking target point reaches a predetermined identification threshold, emitting a narrow pulse width laser signal, entering an accurate measurement mode, and obtaining high-precision laser measurement data. According to the method, the searching time of the lunar orbit satellite angle reflection echo can be effectively shortened, and the success rate of capturing the satellite angle reflection laser echo signal in a complex and deep earth-moon space environment is improved.

The invention discloses a lunar orbit satellite laser ranging data error correction and precision analysis method, and belongs to the technical field of deep space measurement, and the method comprises the steps: setting an internal coincidence error and an external coincidence error of a whole laser ranging process; performing calibration and modeling processing on system errors and random errors in the internal coincidence errors, correcting system delay in real time by adopting a calibration target, and synthesizing a random error standard deviation to evaluate data precision; establishing a physical model for correcting the relativistic effect, atmospheric refraction, earth tide, observation station offset, corner reflector offset and the like in the external coincidence error; and finally, through comparison with an external reference value, a residual error root-mean-square error is calculated, and data accuracy is evaluated. According to the method, systematic correction and precision quantitative analysis of the laser ranging full-link error of the lunar orbit satellite are realized, the ranging data quality is remarkably improved, and reliable data support is provided for high-precision lunar orbit determination.