32 results about "Lunar orbit" patented technology

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 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 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.

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 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.

The invention relates to an autonomous navigation positioning navigation method oriented to a lunar surface and earth-moon space. The method comprises the following steps: time synchronization of a lunar surface navigation base station and a lunar orbit satellite is carried out; the lunar orbit satellite and each lunar surface navigation base station send navigation signals of pseudo code spread spectrum of the station; the served object receives navigation signals from the lunar orbit satellite and each lunar surface navigation base station; performing data fusion processing; and calculating the time, position and speed of the receiver based on the multi-ball intersection measurement principle, calculating the deviation between the current position of the served object and a planned trajectory, and planning trajectory tracking. According to the method, the satellite and the base station are combined, the satellite is used for providing time service and determining the position of the lunar navigation base station, the base station broadcasts a navigation message, and then a receiver of a served object can realize high-precision positioning and navigation of the position of the receiver based on a multi-ball intersection principle.

The application belongs to the technical field of satellite navigation and positioning, and provides a navigation and positioning method based on a lunar orbit receiver, so as to solve problems such as time synchronization error and positioning accuracy of the lunar orbit receiver in the positioning process. The application introduces a method for bidirectional time synchronization with ground communication equipment, so that the lunar orbit receiver can accurately calculate the clock difference with the GPS system, thereby overcoming the time synchronization error caused by signal delay in the prior art, and effectively improving the positioning accuracy. Meanwhile, the lunar orbit receiver also receives the GPS satellite broadcast signal, and combines the clock difference to perform real-time positioning calculation, so as to ensure accurate positioning in the lunar orbit.

This invention discloses a method and system for autonomous orbit determination of a lunar spacecraft based on DRO relay. The method first constructs a two-layer navigation architecture consisting of a long-distance retrograde orbit (DRO) navigation satellite and a lunar orbit user satellite. Ground-based tracking stations track the DRO navigation satellite to maintain its high-precision ephemeris, using it as a space reference node. The DRO navigation satellite broadcasts navigation signals to the lunar orbit user satellite, which receives the signals and acquires pseudorange and carrier phase observation data for the inter-satellite link. Finally, based on the high-precision ephemeris and observation data, the lunar orbit user satellite uses a weighted estimation algorithm to calculate its own orbital state. This invention leverages the long-term stability and high coverage of the DRO orbit, significantly reducing the dependence of lunar exploration missions on ground-based tracking and control resources. Simultaneously, by fusing carrier phase observations, it achieves sub-meter-level high-precision autonomous orbit determination for lunar spacecraft, providing a high-precision space reference for lunar surface positioning and navigation.

A lunar orbit satellite angular momentum prediction method, device and electronic equipment. The method comprises: acquiring satellite historical angular momentum and attitude quaternion telemetry data, dividing the angular momentum into multiple periods with the rapid change process as a node, and performing polynomial fitting on the angular momentum and attitude quaternion in each period to complete and smooth the missing data due to the control constraint. Based on the completed historical angular momentum and attitude quaternion, the historical angular momentum increment of the satellite in the system is calculated. Based on the satellite ephemeris, the historical angle between the body axis with the largest solar pressure area of the satellite and the sun position vector in the system is calculated. Using a preset order sine and cosine model, the historical angle and its corresponding historical angular momentum increment are fitted to obtain the fitting coefficients of the three-axis directions in the system, and an angular momentum increment fitting model is constructed. The model is used to predict the angular momentum prediction value of the satellite in the system, and the accurate prediction of the angular momentum is realized.

The present invention provides a multi-source fusion positioning method and system for the lunar South Pole that integrates lunar orbit satellites, inertial navigation, and wireless communications. The method comprises constructing a multi-source fusion positioning measurement model using lunar orbit satellite pseudorange measurements and wireless communication time difference of arrival (TDOA) measurements, and calculating position information using the least squares method; using the system error of the INS system as the error state, and the difference between the position calculated by the INS system and the least squares method as the measurement error state; constructing a loosely coupled model based on the error state Kalman filter (ESKF) to calculate the error state and correct the INS's calculated position, and updating the current INS internal information to obtain the accurate position of the lunar rover. The present invention effectively solves the problems of limited coverage and low positioning accuracy caused by the insufficient number of lunar orbit satellites; reduces the impact of INS error accumulation on positioning accuracy, and ultimately obtains the precise position of the lunar rover; and significantly improves the overall positioning accuracy and reliability of the system.

The invention belongs to the technical field of satellite navigation and positioning, and provides a navigation and positioning method based on a lunar orbit receiver, so as to solve the problems of time synchronization error, positioning precision and the like of the lunar orbit receiver in the positioning process. And the lunar orbit receiver can accurately calculate the clock error with the GPS system, so that the time synchronization error caused by signal delay in the existing method is overcome, and the positioning accuracy is effectively improved. Meanwhile, the lunar orbit receiver can also receive GPS satellite broadcast signals, real-time positioning calculation is carried out in combination with clock correction, and it is ensured that the accurate position is obtained in the lunar orbit.

The present application relates to the technical field of lunar orbit image analysis, in particular to a crater detection method based on orbit camera images, a training data set is constructed based on lunar orbit camera images, the training data set contains crater edge heat map labels, offset direction labels and radius labels based on multi-interval coding; a deep learning network is trained using the data set, the network output layer is set to three branches, respectively used for predicting edge heat maps, offset directions and radius multi-intervals; input the image to be detected into the network, obtain the prediction result and filter the effective edge pixel set based on the heat map, decode the direction and radius to generate the crater center candidate point; the mean shift clustering algorithm is used to separate the candidate points into instances, and the crater detection result is obtained, the combination of edge prediction and multi-interval coding improves the detection rate and positioning accuracy of weak edges, small scales and adjacent craters.

The invention discloses a method for measuring the expansion rate of a lunar orbit by using an earth-moon light clock. The method comprises the following steps: acquiring relative frequency deviation between a ground light clock and a lunar surface light clock obtained by continuous measurement of an earth-moon bidirectional time-frequency transmission link; the earth-moon two-way time-frequency transmission link is established between a lunar surface observation station and a ground observation station and is used for realizing optical clock time-frequency signal comparison and laser ranging, and the lunar surface observation station and the ground observation station are both provided with optical clocks superior to 10-20 orders of magnitude; establishing a physical relation model between the relative frequency deviation change rate and the lunar orbit expansion rate based on a generalized relativistic gravitational red shift theory; and on the basis of the physical relation model, the lunar orbit expansion rate is calculated according to the measured relative frequency deviation change rate. According to the invention, direct, continuous and high-precision measurement of lunar orbit expansion is realized, the limitation of the traditional moon laser ranging technology is overcome, and a new generation of measurement means is provided for earth-moon system dynamics research and basic physical examination.