Method and apparatus for determining orbit of detector, electronic device and storage medium

Abstract

The present invention discloses an orbit determination method and device for a detector, an electronic device, and a storage medium. Among them, the determination method includes: obtaining a basic star detection orbit from the Earth to a target star and deep space network orbit measurement data, initializing the orbit measurement information based on the basic star detection orbit and the deep space network orbit measurement data, after completing the initialization processing of the orbit measurement information, performing numerical integration on the detector to obtain the operating trajectory parameters of the detector in the station coordinate system, and determining the orbit parameters of the detector based on the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector. The present invention solves the technical problem in the related art that the orbit parameters of the detector cannot be accurately determined.

Description

Technical Field

[0001] The present invention relates to the field of space technology, and in particular, to an orbital determination method and apparatus for a detector, an electronic device, and a storage medium. Background Art

[0002] In the project of exploring celestial bodies (for example, Mars exploration), the success of the project is marked by an orbiter orbiting the celestial body, a detection device (such as a Mars rover for detecting Mars) landing on the surface of the celestial body, and carrying out scientific exploration. Among them, the engineering goals to be achieved include: (1) breaking through key technologies such as celestial body braking capture, entry / descent / landing, long-term autonomous management, long-distance TT&C communication, and surface patrol of the celestial body, realizing orbiting exploration and surface patrol exploration of the celestial body, obtaining scientific data on Mars exploration of the celestial body, and achieving a leap in the field of deep space exploration; (2) building an independent basic engineering system for deep space exploration, including design, manufacturing, testing, implementation of flight missions, scientific research, project management, and talent teams, and promoting the sustainable development of deep space exploration activities.

[0003] The scientific goals to be achieved are: through orbiting exploration, carrying out global and comprehensive exploration of the celestial body, and through surface patrol exploration, carrying out high-precision and high-resolution fine exploration of key areas on the surface of the celestial body.

[0004] In the related art, in the celestial body exploration mission, there are still the following problems: (1) As the distance between the detector and the Earth increases, the sensitivity of the ground-based measurement data to the orbit gradually decreases, and the geometric certainty of the orbit becomes increasingly poor; (2) The dynamic constraints of the detector orbit gradually weaken, the perturbation magnitude of the solar radiation pressure relative to the planetary gravity increases, and the requirements for the modeling accuracy of the solar radiation pressure in orbit determination are also more stringent; (3) The weakening of the dynamic constraints necessarily requires using a longer orbit measurement arc to carry out orbit determination, and will also face more interference force effects caused by on-board actions; (4) As the detector leaves the Earth's gravitational range, the commonly used Earth space-time reference system is no longer applicable, and the transmission environment of the space signal between the Earth and the celestial body to be detected is more complex.

[0005] In response to the above problems, no effective solution has been proposed yet. Summary of the Invention

[0006] Embodiments of the present invention provide an orbital determination method and apparatus for a detector, an electronic device, and a storage medium, so as to at least solve the technical problem in the related art that the orbital parameters of the detector cannot be accurately determined.

[0007] According to one aspect of the embodiments of the present invention, there is provided a method for determining the orbit of a detector, including: obtaining a basic stellar detection orbit from the Earth to a target star and deep space network orbit measurement data; initializing the orbit measurement information based on the basic stellar detection orbit and the deep space network orbit measurement data; after completing the initialization processing of the orbit measurement information, performing numerical integration on the detector to obtain the operating trajectory parameters of the detector in the station coordinate system; and determining the orbit parameters of the detector based on the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector.

[0008] Optionally, the step of initializing the orbit measurement information based on the basic stellar detection orbit and the deep space network orbit measurement data includes: initializing the dynamic model, initializing the data type of the orbit measurement data, and initializing the orbit measurement arc segment based on the basic stellar detection orbit and the deep space network orbit measurement data, so as to complete the initialization processing of the orbit measurement information, where the dynamic model is used to analyze the change of dynamic parameters in each flight stage during the trajectory process of the detector flying from the Earth to the target star.

[0009] Optionally, after obtaining the basic stellar detection orbit from the Earth to the target star and the deep space network orbit measurement data, it further includes: obtaining the coordinate positions of each measurement station, and performing geodetic coordinate system conversion and station horizon coordinate system conversion based on the coordinate positions of the measurement stations; obtaining the current coordinate position of the detector, and performing conversion on the coordinate reference system used by the detector in each flight stage based on the current coordinate position of the detector.

[0010] Optionally, the step of performing numerical integration on the detector to obtain the operating trajectory parameters of the detector in the station coordinate system includes: analyzing the current position, current speed, solar radiation pressure coefficient, and attitude control jet parameters of the detector in each flight stage from the Earth to the target star; performing numerical integration on the current position, current speed, solar radiation pressure coefficient, and attitude control jet parameters to calculate the position and speed of the detector in the station coordinate system, so as to obtain the operating trajectory parameters.

[0011] Optionally, the step of determining the orbital parameters of the detector based on the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector includes: performing a first-order Taylor expansion on the observation equation of the detector to determine the correlation relationship between the orbiter, the ascender, and the systematic error of the observation, and establishing a system of linear regression equations; using the system of linear regression equations to calculate the theoretical observable, the observation partial derivative, and the state partial derivative of the detector; substituting the theoretical observable, the observation partial derivative, and the state partial derivative into the expanded observation equation to calculate the observation data of the detector; and determining the orbital parameters of the detector in combination with the observation data, the weight parameter of the observation data, the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector.

[0012] Optionally, after determining the orbital parameters of the detector based on the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector, it further includes: obtaining the data residuals of calculating the orbital parameters of the detector during each flight stage of the detector, where the data residuals include at least one of the following: ranging residual, velocity measurement residual, time delay residual, and time delay rate residual; obtaining the to-be-traveled orbital parameters of the orbital parameters; and evaluating the to-be-traveled orbital parameters by using the overlapping arc segment method in combination with the data residuals to obtain an orbital evaluation result, where the orbital evaluation result is used to indicate whether the accuracy of the orbital parameters of the detector is within a preset accuracy error range.

[0013] Optionally, the target celestial body includes: Mars.

[0014] According to another aspect of the embodiments of the present invention, there is also provided an orbital determination device for a detector, including: an acquisition unit, configured to acquire the celestial body basic detection orbit from the Earth to the target celestial body and the deep space network orbit measurement data; a processing unit, configured to perform initialization processing on the orbit measurement information based on the celestial body basic detection orbit and the deep space network orbit measurement data; an integration unit, configured to perform numerical integration on the detector after completing the initialization processing of the orbit measurement information to obtain the operating trajectory parameters of the detector in the station coordinate system; and a determination unit, configured to determine the orbital parameters of the detector based on the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector.

[0015] Optionally, the processing unit includes: a first initialization module, configured to perform dynamic model initialization, data type initialization of the orbit measurement data, and orbit measurement arc segment initialization based on the celestial body basic detection orbit and the deep space network orbit measurement data to complete the initialization processing of the orbit measurement information, where the dynamic model is used to analyze the change of the dynamic parameters of the detector during each flight stage in the trajectory process of flying from the Earth to the target celestial body.

[0016] Optionally, the determining device further includes: a first conversion module, configured to, after obtaining the basic detection orbit of the star from the Earth to the target star and the deep space network orbit measurement data, obtain the coordinate positions of each measurement station, and perform geodetic coordinate system conversion and measurement station horizon coordinate system conversion based on the coordinate positions of the measurement stations; a second conversion module, configured to obtain the current coordinate position of the detector, and perform coordinate reference system conversion on the coordinate reference systems used by the detector in each flight stage based on the current coordinate position of the detector.

[0017] Optionally, the integration unit includes: a first analysis module, configured to analyze the current position, current speed, solar radiation pressure coefficient, and attitude control jet parameter of the detector during each flight stage from the Earth to the target star; a first calculation module, configured to perform numerical integration on the current position, current speed, solar radiation pressure coefficient, and attitude control jet parameter, and calculate the position and speed of the detector in the measurement station coordinate system to obtain the operation trajectory parameters.

[0018] Optionally, the determining unit includes: a first determining module, configured to perform a first-order Taylor expansion on the observation equation of the detector, determine the correlation relationship between the system errors of the orbiter, the ascender, and the observation, and establish a linear regression equation set; a second calculation module, configured to use the linear regression equation set to calculate the theoretical observation quantity, observation partial derivative, and state partial derivative of the detector; a third calculation module, configured to substitute the theoretical observation quantity, observation partial derivative, and state partial derivative into the expanded observation equation to calculate the observation data of the detector; a second determining module, configured to combine the observation data, the weight parameter of the observation data, the operation trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector to determine the orbit parameters of the detector.

[0019] Optionally, the determining device further includes: a first obtaining module, configured to, after determining the orbit parameters of the detector based on the operation trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector, obtain the data residuals of calculating the orbit parameters of the detector during each flight stage, where the data residuals include at least one of the following: ranging residual, velocity measurement residual, time delay residual, time delay rate residual; a second obtaining module, configured to obtain the to-be-traveled orbit parameters of the orbit parameters; a first evaluation module, configured to combine the data residuals and evaluate the to-be-traveled orbit parameters by the overlapping arc segment method to obtain an orbit evaluation result, where the orbit evaluation result is used to indicate whether the accuracy of the orbit parameters of the detector is within a preset accuracy error range.

[0020] Optionally, the target star includes: Mars.

[0021] According to another aspect of the embodiments of the present invention, an electronic device is further provided, including: a processor; and a memory for storing executable instructions of the processor; wherein, the processor is configured to execute the orbit determination method of the above-mentioned detector by executing the executable instructions.

[0022] According to another aspect of the embodiments of the present invention, a computer-readable storage medium is further provided. The computer-readable storage medium includes a stored computer program, wherein when the computer program runs, it controls the device where the computer-readable storage medium is located to execute the orbit determination method of the above-mentioned detector.

[0023] In the present disclosure, the basic star detection orbit from the Earth to the target star and the deep space network orbit measurement data are obtained. Based on the basic star detection orbit and the deep space network orbit measurement data, the orbit measurement information is initialized. After the initialization process of the orbit measurement information is completed, numerical integration is performed on the detector to obtain the running trajectory parameters of the detector in the measurement station coordinate system. Based on the running trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector, the orbit parameters of the detector are determined. In the present application, the initialization process of the orbit measurement information can be completed according to the basic star detection orbit from the Earth to the target star and the deep space network orbit measurement data, and then numerical integration is performed on the detector to obtain the running trajectory parameters of the detector in the measurement station coordinate system. Thus, the orbit parameters of the detector can be accurately determined according to the running trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector, thereby solving the technical problem in the related art that the orbit parameters of the detector cannot be accurately determined. Description of the Drawings

[0024] The drawings described herein are used to provide a further understanding of the present invention, and constitute a part of this application. The schematic embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0025] Figure 1 is a flowchart of an optional orbit determination method of a detector according to an embodiment of the present invention;

[0026] Figure 2 is a schematic diagram of an optional coordinate system conversion process according to an embodiment of the present invention;

[0027] Figure 3 is a schematic diagram of an optional tracking and measurement of a Mars detector according to an embodiment of the present invention;

[0028] Figure 4 is a flowchart of an optional statistical orbit determination of a Mars detector according to an embodiment of the present invention;

[0029] Figure 5It is a schematic diagram of determining ephemeris error for an optional cruise - segment orbit according to an embodiment of the present invention;

[0030] Figure 6 It is a schematic diagram of determining ephemeris error for an optional orbit during the circum - Mars - flyby segment according to an embodiment of the present invention;

[0031] Figure 7 It is a schematic diagram of an optional time - system conversion process according to an embodiment of the present invention;

[0032] Figure 8 It is a schematic diagram of an optional orbit - determination device for a detector according to an embodiment of the present invention. Detailed implementation manners

[0033] In order to enable those skilled in the art to better understand the solution of the present invention, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

[0034] It should be noted that the terms "first", "second", etc. in the specification and claims of the present invention and the above - mentioned drawings are used to distinguish similar objects, and do not necessarily need to describe a specific order or sequence. It should be understood that such used data can be interchanged under appropriate circumstances so that the embodiments of the present invention described here can be implemented in an order different from those illustrated or described here. In addition, the terms "include" and "have" and any variations thereof are intended to cover non - exclusive inclusion. For example, a process, method, system, product or device that includes a series of steps or units does not necessarily have to be limited to those clearly listed steps or units, but may include other steps or units not clearly listed or inherent to these process, method, product or device.

[0035] The following embodiments of the present invention can be applied to the orbit determination task of a detector, and specifically can be applied to the trajectory tracking system of the detector. The present invention can design a spatio-temporal reference system suitable for deep space exploration according to the orbit characteristics of various celestial bodies (illustrated by the Mars celestial body in this application), design a dynamic model according to the orbit characteristics of celestial body detection, design an observation model according to the tracking characteristics of the deep space network, design a numerical integration algorithm according to the time coordinate system and the dynamic model, and design a statistical orbit determination method according to various models. The present invention can also complete the initialization settings such as spatio-temporal reference system conversion, dynamic model, data type, orbit arc segment, parameter solution, etc. according to the deep space network orbit measurement data and the celestial body detection orbit, and then perform numerical integration on the detector, construct an observation equation, and perform a first-order Taylor expansion on the equation to obtain a linear differential equation, complete the calculation of theoretical observation quantities, 0-C residuals, observation partial derivatives and state partial derivatives, and substitute them into the expanded equation. According to the weighted least squares principle, perform differential iteration to complete the solution of the orbit parameters and other model parameters of the celestial body detector.

[0036] The present invention will be described in detail below in conjunction with each embodiment.

[0037] Embodiment 1

[0038] According to an embodiment of the present invention, there is provided an embodiment of an orbit determination method for a detector. It should be noted that the steps shown in the flowchart of the accompanying drawings can be executed in a computer system such as a set of computer executable instructions, and although the logical order is shown in the flowchart, in some cases, the steps shown or described can be executed in a different order than here.

[0039] Figure 1 is a flowchart of an optional orbit determination method for a detector according to an embodiment of the present invention, as Figure 1 shown, the method includes the following steps:

[0040] Step S102, obtain the basic celestial body detection orbit from the Earth to the target celestial body and the deep space network orbit measurement data.

[0041] Step S104, based on the basic celestial body detection orbit and the deep space network orbit measurement data, perform initialization processing on the orbit measurement information.

[0042] Step S106, after completing the initialization processing of the orbit measurement information, perform numerical integration on the detector to obtain the operating trajectory parameters of the detector in the measurement station coordinate system.

[0043] Step S108, based on the operating trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector, determine the orbit parameters of the detector.

[0044] Through the above steps, the basic detection orbit of the star from the Earth to the target star and the deep space network orbit measurement data can be obtained. Based on the basic detection orbit of the star and the deep space network orbit measurement data, the orbit measurement information is initialized. After the initialization of the orbit measurement information is completed, numerical integration is performed on the detector to obtain the trajectory parameters of the detector in the station coordinate system. Based on the trajectory parameters, the station coordinate position, and the initial coordinate position of the detector, the orbit parameters of the detector are determined. In the embodiment of the present invention, the initialization of the orbit measurement information can be completed according to the basic detection orbit of the star from the Earth to the target star and the deep space network orbit measurement data, and then numerical integration is performed on the detector to obtain the trajectory parameters of the detector in the station coordinate system. Thus, the orbit parameters of the detector can be accurately determined according to the trajectory parameters, the station coordinate position, and the initial coordinate position of the detector, thereby solving the technical problem in the related art that the orbit parameters of the detector cannot be accurately determined.

[0045] The embodiments of the present invention will be described in detail below in conjunction with the above steps.

[0046] In the embodiment of the present invention, the detection of various stars can be realized. Optionally, the target stars include: Mars, Venus, etc. In this embodiment, Mars is used as an illustrative option for description.

[0047] Step S102, obtain the basic detection orbit of the star from the Earth to the target star and the deep space network orbit measurement data.

[0048] In the embodiment of the present invention, the deep space network orbit measurement data includes: UXB (Unified X-Band) ranging, velocity measurement, and VLBI (Very Long Base Interferometer) time delay, time delay rate, SBI (Same Beam Interferometry) time delay, time delay rate, and inter-satellite measurement (microwave radar, lidar) ranging, velocity measurement data, etc.

[0049] In this embodiment, the initialization of the orbit measurement information can be completed according to the deep space network orbit measurement data and the Mars detection orbit (i.e., the basic detection orbit of the star from the Earth to the target star).

[0050] Optionally, after obtaining the basic detection orbit of the star from the Earth to the target star and the deep space network orbit measurement data, it further includes: obtaining the coordinate positions of each station, and performing geodetic coordinate system conversion and station horizon coordinate system conversion based on the coordinate positions of the stations; obtaining the current coordinate position of the detector, and performing coordinate reference system conversion on the coordinate reference systems used by the detector in each flight stage based on the current coordinate position of the detector.

[0051] In an embodiment of the present invention, taking a Mars probe as an example, the coordinate systems involved in the orbit calculation of the probe include but are not limited to: Earth-fixed coordinate system, geodetic coordinate system, station horizon coordinate system, geocentric inertial coordinate system, heliocentric inertial coordinate system, solar system barycentric inertial coordinate system (referred to as the barycentric coordinate system), martian-centered inertial coordinate system, Mars-fixed coordinate system, and orbit coordinate system. Figure 2 is a schematic diagram of an optional coordinate system conversion process according to an embodiment of the present invention, as Figure 2 shown, including: Earth-fixed coordinate system, geodetic coordinate system, station horizon coordinate system, geocentric inertial coordinate system, heliocentric inertial coordinate system, barycentric coordinate system, martian-centered inertial coordinate system, Mars-fixed coordinate system. Among them, steps 1 and 2 (i.e., from the Earth-fixed coordinate system to the geodetic coordinate system, from the geodetic coordinate system to the station horizon coordinate system) involve coordinate conversion of the station, and steps 3 to 6 (i.e., from the Earth-fixed coordinate system to the geocentric inertial coordinate system, from the geocentric inertial coordinate system to the martian-centered inertial coordinate system, from the martian-centered inertial coordinate system to the Mars-fixed coordinate system, from the geocentric inertial coordinate system to the heliocentric inertial coordinate system) involve coordinate conversion of the probe. It also includes: step 7 from the heliocentric inertial coordinate system to the barycentric coordinate system, step 8 from the geocentric inertial coordinate system to the barycentric coordinate system, and step 9 from the martian-centered inertial coordinate system to the barycentric coordinate system.

[0052] In this embodiment, the geocentric inertial coordinate system is defined as: the coordinate origin is the center of the Earth, the xy coordinate plane is close to the mean equator at the epoch J2000.0, the x-axis is close to pointing to the mean vernal equinox of this epoch, and the other inertial coordinate systems are all translations of the coordinate origin, with the reference plane and pointing remaining unchanged.

[0053] In this embodiment, the coordinate positions of each station can be obtained first, so that the geodetic coordinate system conversion and the station horizon coordinate system conversion can be carried out based on the coordinate positions of the stations. The current coordinate position of the probe can also be obtained, so that the coordinate reference systems used by the probe in each flight stage can be converted based on the current coordinate position of the probe. In this embodiment, the flight stages may include: low-Earth flight stage, cruise flight stage, and orbiting flight stage around the celestial body, etc.

[0054] Step S104, based on the basic detection orbit of the celestial body and the orbit measurement data of the deep space network, perform initialization processing on the orbit measurement information.

[0055] Optionally, the step of performing initialization processing on the orbit measurement information based on the basic detection orbit of the celestial body and the orbit measurement data of the deep space network includes: based on the basic detection orbit of the celestial body and the orbit measurement data of the deep space network, perform initialization of the dynamic model, initialization of the data type of the orbit measurement data, and initialization of the orbit measurement arc segment, and complete the initialization processing of the orbit measurement information. Among them, the dynamic model is used to analyze the change of the dynamic parameters of the probe in each flight stage during the trajectory process from the Earth to the target celestial body.

[0056] In an embodiment of the present invention, initialization settings such as spatio-temporal reference frame conversion, dynamic model, data type, orbit measurement arc segment, and parameter solution (i.e., initializing the orbit measurement information) can be completed according to the basic detection orbit of the celestial body and the orbit measurement data of the deep space network. Specifically: Based on the basic detection orbit of the celestial body and the orbit measurement data of the deep space network, dynamic model initialization, data type initialization of the orbit measurement data, and orbit measurement arc segment initialization can be performed, so as to complete the initialization processing of the orbit measurement information. The dynamic model in this embodiment is used to analyze the change of dynamic parameters in each flight stage during the trajectory process of the detector flying from the Earth to the target celestial body.

[0057] Step S106, after completing the initialization processing of the orbit measurement information, perform numerical integration on the detector to obtain the running trajectory parameters of the detector in the measurement station coordinate system.

[0058] Optionally, the step of performing numerical integration on the detector to obtain the running trajectory parameters of the detector in the measurement station coordinate system includes: analyzing the current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters of the detector during each flight stage from the Earth to the target celestial body; performing numerical integration on the current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters, and calculating the position and velocity of the detector in the measurement station coordinate system to obtain the running trajectory parameters.

[0059] In an embodiment of the present invention, after completing the initialization processing of the orbit measurement information, numerical integration can be performed on the detector to obtain the running trajectory parameters of the detector in the measurement station coordinate system. In this embodiment, for the Mars exploration orbit, the numerical integration method can obtain the position and velocity of the detector at a specified moment and the corresponding partial derivatives. Considering the calculation accuracy and efficiency, the multi-step integration method can be used for numerical integration. The KSG (Krogh-Shampine-Gardon) fixed-step integration method can be used when there is no attitude control jet, and the Runge-Kutta single-step method can be used when attitude control jets are included. Specifically: The current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters of the detector during each flight stage from the Earth to the target celestial body can be analyzed, and numerical integration is performed on the current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters to calculate the position and velocity of the detector in the measurement station coordinate system to obtain the running trajectory parameters.

[0060] In this embodiment, the basic process of determining the orbit of the Mars detector is to perform differential correction on a set of parameter estimates from the observation model so that the weighted sum of squares of the differences between the measured observation data and the corresponding quantities calculated by the model is minimized. The observable quantity (i.e., the obtained running trajectory parameters) is determined by geometric relationships and can be expressed as the following formula (1):

[0061]

[0062] where t represents the time stamp of the observed data, and δt represents the time constant systematic error, the position and velocity of the aircraft corresponding to t = t + δt in the station coordinate system, represents the dynamic parameters, which include variables related to the initial state parameters of the aircraft, gravitational field coefficients, damping coefficients, etc., represents the coordinate quantity of the station, b represents the observation constant systematic error, RF c including the correction values of the observed data caused by atmospheric refraction, transponder delay, antenna pedestal error correction, etc., represents the model parameters to be estimated.

[0063] In this embodiment, the Mars exploration process includes UXB measurements: ranging, velocity measurement, and angle measurement, as well as the time delay, time delay rate, right ascension, and declination of VLBI. Figure 3 is a schematic diagram of an optional tracking and measurement of a Mars probe according to an embodiment of the present invention, as Figure 3 shown, including: the launch vehicle TT&C system composed of Chinese satellites and launch vehicles, the near-Earth space TT&C network, the international link station, the VLBI orbit determination subsystem, the antenna array, the deep space TT&C network (Deep Space Station 1, Deep Space Station 2), where the Earth-Mars transfer section involves the antenna array and the deep space TT&C network (Deep Space Station 1, Deep Space Station 2), and the Mars-orbiting working section involves the VLBI orbit determination subsystem, the antenna array, and the deep space TT&C network (Deep Space Station 1, Deep Space Station 2). In the calculation of theoretical observable quantities, the near-Earth orbit is calculated in the geocentric inertial system, and other types of orbits are calculated in the barycentric coordinate system. Among them, the near-Earth orbit refers to an orbit where the spacecraft is at a relatively low altitude above the ground. Generally, a near-circular orbit with an altitude below 2000 kilometers can be called a near-Earth orbit.

[0064] Step S108, based on the operating trajectory parameters, the coordinate position of the station, and the initial coordinate position of the detector, determine the orbit parameters of the detector.

[0065] Optionally, the step of determining the orbit parameters of the detector based on the operating trajectory parameters, the coordinate position of the station, and the initial coordinate position of the detector includes: performing a first-order Taylor expansion on the observation equation of the detector to determine the correlation relationship between the orbiter, the ascender, and the systematic error of the observation, and establishing a linear regression equation system; using the linear regression equation system to calculate the theoretical observable quantity, the observation partial derivative, and the state partial derivative of the detector; substituting the theoretical observable quantity, the observation partial derivative, and the state partial derivative into the expanded observation equation to calculate the observed data of the detector; combining the observed data, the weight parameter of the observed data, the operating trajectory parameters, the coordinate position of the station, and the initial coordinate position of the detector to determine the orbit parameters of the detector.

[0066] In an embodiment of the present invention, the orbital parameters of the detector can be determined based on the running track parameters, the coordinate positions of the measuring stations, and the initial coordinate position of the detector. Specifically: perform a first-order Taylor expansion on the observation equation (Formula (1)):

[0067]

[0068] Among them, q is the actual observed data, q c is the calculated observed value at the initial state epoch according to Formula (2), σ is the parameter to be estimated, and e is the observed white noise.

[0069] In this embodiment, Formula (2) is a first-order Taylor series expansion formula near the predicted observed value, which determines the model of the deviation between the actual observed value and the calculated value, establishes the relationship between the systematic errors of the orbiter, the ascender, and the observation in the observed data residuals, and establishes the required linear regression equations (that is, perform a first-order Taylor expansion on the observation equation of the detector, determine the correlation relationship between the systematic errors of the orbiter, the ascender, and the observation, and establish a linear regression equation).

[0070] In the determination of the Mars exploration orbit, the parameter σ to be estimated includes the position and velocity at time t, the observation systematic error b, and the dynamic parameter p. Then the total estimated parameter is expressed as the following Formula (3):

[0071]

[0072] The observation equation is further expressed as Formula (4):

[0073] y = q - q c = FΔσ + e (4);

[0074] Among them, F can be expressed as:

[0075]

[0076] Its linear unbiased minimum variance estimate can be expressed as the following Formula (6):

[0077]

[0078] Among them, R -1 represents the weight of the observed data.

[0079] In this embodiment, a linear regression equation system can be used to calculate the theoretical observable quantity, the observation partial derivative, and the state partial derivative of the detector. Substitute the theoretical observable quantity, the observation partial derivative, and the state partial derivative into the expanded observation equation to calculate the observation data of the detector. Then, combine the observation data, the weight parameter of the observation data, the running trajectory parameter, the coordinate position of the measuring station, and the initial coordinate position of the detector to determine the orbit parameter of the detector (that is, after obtaining the initial position of the Mars detector, the position of the measuring station, and the observable quantity, the weighted least squares improvement of the detector orbit and other parameters can be carried out to achieve accurate solution).

[0080] Figure 4 It is a flowchart of an optional statistical orbit determination of a Mars detector according to an embodiment of the present invention. As Figure 4 shown, it includes the following steps: First, perform program initialization, read in the initial orbit state, the initial value setting of ephemeris calculation, and the time range setting, etc. When the maximum number of allowed iterations is zero, set the last iteration calculation flag, and then calculate the detector ephemeris to record the ephemeris file. Otherwise, directly calculate the detector ephemeris to record the ephemeris file. Then, obtain the observation data from the data file, perform ephemeris table interpolation to calculate the state value at the observation moment, and retain the interpolation parameters. Calculate the observed value and the observation partial derivative (if necessary, solve the light travel time). Then, determine the weight of the observed value, calculate the observation residual and accumulate the normal matrix. Output the residual to the file in the last iteration. Judge whether all the observation data has been processed. If not, re-obtain the observation data from the data file. Otherwise, perform the inverse of the normal matrix, calculate the correction amount of the solution parameter and the corrected solution parameter, and then calculate the RMS and various statistical quantities. Judge whether the iteration diverges. In the case of iteration divergence, output the iteration divergence information and exit the program. In the case of no iteration divergence, judge whether this iteration is the last iteration. In the case of the last iteration, output the solution variable result and various calculation information and exit the program. In the case of not the last iteration, judge whether the iteration converges. In the case of convergence, obtain the solution parameter before the correction of this iteration, and set the next iteration as the last iteration flag. Then, increment the iteration counter by one. If the next iteration is the maximum number of allowed iterations, set the last iteration flag and re-perform the step of calculating the detector ephemeris to record the ephemeris file. In the case of non-convergence, directly increment the iteration counter by one. If the next iteration is the maximum number of allowed iterations, set the last iteration flag and re-perform the step of calculating the detector ephemeris to record the ephemeris file.

[0081] Optionally, after determining the orbital parameters of the detector based on the running track parameters, the station coordinate positions, and the initial coordinate positions of the detector, it further includes: obtaining the data residuals for calculating the orbital parameters of the detector during each flight stage, where the data residuals include at least one of the following: ranging residuals, velocity measurement residuals, time delay residuals, and time delay rate residuals; obtaining the to-be-traveled orbital parameters among the orbital parameters; combining the data residuals, and evaluating the to-be-traveled orbital parameters by the overlapping arc segment method to obtain an orbital evaluation result, where the orbital evaluation result is used to indicate whether the accuracy of the orbital parameters of the detector is within the preset accuracy error range.

[0082] In the embodiments of the present invention, the statistical orbit determination strategy varies according to the measurement and tracking data and different flight stages. During the detector's orbit injection, cruise stage, and parking stage at the initial stage of Mars orbit, the UXB and VLBI tracking intensities are high, and it is basically in continuous tracking. After entering the relay orbit and parking orbit, basically only UXB tracking is required. Specifically, as shown in Table 1:

[0083] Table 1

[0084]

[0085]

[0086] In this embodiment, the data residuals can be obtained through orbit determination calculation (that is, obtaining the data residuals for calculating the orbital parameters of the detector during each flight stage). Among them, the data residuals can include: ranging residuals, velocity measurement residuals, time delay residuals, time delay rate residuals, etc. After obtaining the to-be-traveled orbital parameters among the orbital parameters, the to-be-traveled orbital parameters can be evaluated by the overlapping arc segment method in combination with the data residuals to obtain an orbital evaluation result (the orbital evaluation result is used to indicate whether the accuracy of the orbital parameters of the detector is within the preset accuracy error range). Among them, the orbital evaluation result includes: the orbit determination ephemeris error in the cruise stage and the orbit determination ephemeris error in the Mars orbit stage, etc.

[0087] Figure 5 It is a schematic diagram of an optional orbit determination ephemeris error in the cruise stage according to the embodiments of the present invention. As Figure 5 shown, with t (hour) as the abscissa (ranging from 0 to 150, with each 50 hours as a unit), and δR (m) (ranging from -2000 to -500) and δV (m / s) (ranging from -4 to 4) as the ordinates respectively, an axis representing the position and velocity deviation of the predicted orbit in the RTN direction is formed. Among them, (a) represents the velocity deviation of the δr, δt, and δn curves, and (b) represents the velocity deviation of the δrd, δtd, and δnd curves.

[0088] Figure 6It is a schematic diagram for determining ephemeris errors of the ring fire section orbit according to an embodiment of the present invention. As shown in Figure 6 the figure, with t (hour) as the abscissa (ranging from 0 to 50, with each unit being 5 hours), and δR (m) (ranging from -400 to 600) and δV (m / s) (ranging from -0.2 to 0.2) as the ordinates respectively, coordinate axes representing the position and velocity deviations of the predicted orbit in the RTN direction are formed. Among them, (a) represents the velocity deviations of the three curves of δr, δt, and δn, and (b) represents the velocity deviations of the three curves of δrd, δtd, and δnd.

[0089] The following will be described in detail in combination with another optional specific embodiment.

[0090] The embodiment of the present invention can use multi-target joint measurement of the lunar orbit for modeling, and proposes a method applicable to multi-target joint orbit determination. It can complete the joint orbit determination of multiple targets based on the ranging, velocity measurement of the lunar probe UXB, VLBI time delay, time delay rate, SBI time delay, time delay rate, and inter-satellite measurement (microwave radar, lidar) ranging, velocity measurement data and the initial orbit, and obtain the precise orbit and relative orbit of the lunar multi-target probe. This embodiment can be based on the multi-target joint measurement of the lunar exploration project, establish the observation models of UXB ranging, velocity measurement, VLBI time delay, time delay rate, SBI time delay, time delay rate, and inter-satellite measurement (microwave radar, lidar) ranging, velocity measurement data, design the statistical orbit determination algorithm for the combination of the orbiter and the ascender, calibrate the data accuracy of the same-beam measurement time delay, time delay rate, and inter-satellite measurement ranging, velocity measurement using the reference orbit, and introduce the same-beam measurement in orbit determination, which can significantly improve the relative orbit accuracy of the two targets. Especially in the case of short arc segment measurement, it is necessary to introduce the time delay, time delay rate, and inter-satellite ranging of the same-beam measurement simultaneously to achieve the convergence standard of statistical orbit determination calculation.

[0091] The method proposed in this embodiment is applicable to the manned rendezvous and docking process of subsequent manned lunar landing and the orbit determination calculation of unmanned rendezvous and docking of other extraterrestrial celestial bodies. Moreover, the orbit determination method of the Mars probe in this embodiment can be used not only for single-target and double-target, but also for the joint orbit determination calculation of multiple lunar probes.

[0092] The following will be specifically described in combination with the following steps.

[0093] Step 1: Time system and its conversion.

[0094] The time systems involved in the detector orbit calculation include: Universal Time (UT), International Atomic Time (TAI), Terrestrial Dynamical Time (TT), Barycentric Dynamical Time (TDB), and Coordinated Universal Time (UTC).

[0095] Figure 7It is a schematic diagram of an optional time system conversion process according to an embodiment of the present invention, as Figure 7 shown, and can be specifically described as follows:

[0096] (1) Given Beijing Time (BJT), Coordinated Universal Time (UTC) is calculated through time zone conversion, as shown in the following formula:

[0097] BJT = UTC + 8h;

[0098] (2) From Coordinated Universal Time, the corresponding International Atomic Time (TAI) is calculated through leap seconds, as shown in the following formula:

[0099] TAI = UTC + ΔT;

[0100] Among them, ΔT can be set according to specific circumstances. For example, ΔT = 38s.

[0101] (3) From Coordinated Universal Time, the long-term term and long-period term UT1R of Universal Time (UT1) are calculated through linear interpolation. Since the short-period term has little influence, it can be directly considered as (UT1) here;

[0102] (4) Geocentric Dynamical Time (TT) is calculated from International Atomic Time (TAI), as shown in the following formula:

[0103] TT = TAI + 32.184s;

[0104] (5) Conversion to Barycentric Dynamical Time (TDB) according to Geocentric Dynamical Time, as shown in the following formula:

[0105]

[0106] Among them, are respectively the position and velocity vectors of celestial body j relative to point i in the barycentric reference frame of the solar system, which are functions of TDB and can be obtained from the DE (Development Ephemerides) ephemeris. The superscript and subscript C represent the barycenter of the solar system, S represents the sun, B represents the barycenter of the Earth-Moon system, E represents the Earth, M represents the moon, J represents Jupiter, Sa represents Saturn, A represents the position of the atomic clock on the Earth where the time reading is obtained, μ S , μ J , μ Sa respectively represent the gravitational constants of the sun, Jupiter, and Saturn, and the conversion accuracy can ensure an accuracy of about μs. If a one-way time delay observation value needs to be calculated in actual work and an accuracy of ns level is required, a higher-precision calculation form can be adopted.

[0107] Step 2: Coordinate system and its conversion.

[0108] The coordinate systems involved in the orbital calculation of a Mars probe include: the Earth-fixed coordinate system, the geodetic coordinate system, the station horizon coordinate system, the geocentric inertial coordinate system, the heliocentric inertial coordinate system, the barycentric inertial coordinate system of the solar system (abbreviated as the barycentric coordinate system), the martian-centered inertial coordinate system, the Mars-fixed coordinate system, and the orbital coordinate system. The coordinate system conversion process involves the coordinate conversion of the station and the coordinate conversion of the probe. Among them, the geocentric inertial coordinate system is defined as: the coordinate origin is the center of the Earth, the xy coordinate plane is close to the mean equator at the epoch J2000.0, and the x-axis is close to pointing to the mean vernal equinox of this epoch. The other inertial coordinate systems are all translations of the coordinate origin, with the reference plane and direction unchanged.

[0109] Step 3: Perturbation model.

[0110] The dynamic model is used to analyze the changes in the dynamic parameters of the probe during each flight stage in the process of flying from the Earth to the target celestial body. The specific model and parameter table are shown in Table 2:

[0111] Table 2

[0112]

[0113]

[0114] Step 4: Numerical integration.

[0115] For the Mars exploration orbit, the numerical integration method can be used to obtain the position, velocity, and corresponding partial derivatives of the probe at a specified time. The numerical integration can adopt the multi-step method to calculate the integration accuracy and efficiency. When there is no attitude control jet, the KSG (Krogh-Shampine-Gardon) fixed-step integration method is used, and when there is an attitude control jet, the Runge-Kutta single-step method is used for calculation.

[0116] Step 5: Observation model.

[0117] The basic process of determining the orbit of a Mars probe is to perform differential correction on the estimation of a set of parameters from the observation model so that the weighted sum of squares of the difference between the measured observation data and the corresponding quantity calculated by this model is minimized. Among them, the observed quantity is determined by the geometric relationship and can be expressed as (that is, the observation formula can be expressed as):

[0118]

[0119] Among them, t represents the time mark of the observation data, δt represents the time constant systematic error, the position and velocity of the aircraft corresponding to t = t + δt in the station coordinate system, represents the dynamic parameters, and these parameters include variables related to the initial state parameters of the aircraft, gravitational field coefficients, damping coefficients, etc., represents the coordinate quantity of the station, b represents the observation constant systematic error, RFc including the correction values of the observed data caused by atmospheric refraction, transponder time delay, antenna pedestal error correction, etc. b, δt represent the model parameters to be estimated.

[0120] During the Mars exploration process, it includes UXB (Unified X-band) measurements: ranging, velocity measurement, and angle measurement, as well as the time delay, time delay rate, right ascension, and declination of VLBI (Very Long Baseline Interferometry). In the calculation of theoretical observable quantities, the low Earth orbit is calculated in the geocentric inertial system, and other types of orbits are calculated in the barycentric coordinate system.

[0121] Step 6: Statistical orbit determination method.

[0122] The Taylor expansion of the observation formula is:

[0123]

[0124] where q is the actual observed data, q c is the calculated observed value according to the observation formula at the initial state epoch, σ is the parameter to be estimated, and e is the observation white noise.

[0125] In this embodiment, the observation formula is the first-order Taylor series expansion around the predicted observed value, which determines the model of the deviation between the actual observed value and the calculated value, establishes the relationship between the orbital vehicle, the ascender, and the systematic error of the observation in the observation data residuals, and establishes the required linear regression equation system.

[0126] In the determination of the Mars exploration orbit, the parameter σ to be estimated includes the position and velocity at time t, the observation system error b, and the dynamic parameter p. Then the total estimated parameter is expressed as:

[0127]

[0128] The observation equation is further expressed as:

[0129] y = q - q c = FΔσ + e;

[0130]

[0131] Its linear unbiased minimum variance estimate can be expressed as:

[0132]

[0133] where R -1 represents the weight of the observed data.

[0134] In this embodiment, after obtaining the initial position of the Mars probe, the station position, and the observables, the least squares improvement with weights can be performed on the probe orbit and other parameters to achieve accurate solution.

[0135] Step 7: Data verification and orbit accuracy analysis.

[0136] In the embodiment of the present invention, the statistical orbit determination strategy varies according to the measurement and tracking data and different flight phases. During the probe's orbit insertion, cruise phase, and parking phase at the initial stage of Mars orbit, the UXB and VLBI tracking intensities are high and are basically in continuous tracking. After entering the relay orbit and parking orbit, basically only UXB tracking is required.

[0137] In this embodiment, the data residuals can be calculated through orbit determination. Among them, the data residuals can include: ranging residuals, velocity residuals, time delay residuals, time delay rate residuals, etc. The accuracy of Mars probe orbit determination can be evaluated using the overlapping arc method. Among them, the orbit accuracy in the cruise phase is as Figure 5 shown, and the orbit accuracy in the Mars orbit phase is as Figure 6 shown.

[0138] In the embodiment of the present invention, the initialization settings such as space-time reference system conversion, dynamic model, data type, orbit determination arc segment, and parameter solution can be completed according to the deep space network orbit measurement data and the Mars exploration orbit. Then, numerical integration is performed on the probe, an observation equation is constructed, and the equation is expanded by the first-order Taylor expansion to obtain a linear differential equation. The theoretical observables, 0-C residuals, observation partial derivatives, and state partial derivatives are calculated and substituted into the expanded equation. According to the weighted least squares principle, differential iteration is performed to solve the orbit parameters and other model parameters of the Mars probe, so that the orbit parameters of the probe can be accurately determined, and the orbit determination of the probe is completed.

[0139] Embodiment 2

[0140] An orbit determination device for a probe provided in this embodiment includes multiple implementation units, and each implementation unit corresponds to each implementation step in Embodiment 1 above.

[0141] Figure 8 is a schematic diagram of an optional orbit determination device for a probe according to an embodiment of the present invention, as Figure 8 shown. The determination device may include: an acquisition unit 80, a processing unit 82, an integration unit 84, and a determination unit 86, where

[0142] The acquisition unit 80 is used to acquire the basic detection orbit of the star body from the Earth to the target star body and the deep space network orbit measurement data;

[0143] A processing unit 82, configured to initialize the orbit determination information based on the basic detection orbit of the celestial body and the orbit determination data of the Deep Space Network;

[0144] An integration unit 84, configured to perform numerical integration on the detector after the initialization processing of the orbit determination information is completed, so as to obtain the running trajectory parameters of the detector in the station coordinate system;

[0145] A determination unit 86, configured to determine the orbit parameters of the detector based on the running trajectory parameters, the coordinate position of the station, and the initial coordinate position of the detector.

[0146] The above determination device can obtain the basic detection orbit of the celestial body from the Earth to the target celestial body and the orbit determination data of the Deep Space Network through the acquisition unit 80, initialize the orbit determination information through the processing unit 82 based on the basic detection orbit of the celestial body and the orbit determination data of the Deep Space Network, perform numerical integration on the detector through the integration unit 84 after the initialization processing of the orbit determination information is completed, so as to obtain the running trajectory parameters of the detector in the station coordinate system, and determine the orbit parameters of the detector through the determination unit 86 based on the running trajectory parameters, the coordinate position of the station, and the initial coordinate position of the detector. In the embodiment of the present invention, the initialization processing of the orbit determination information can be completed according to the basic detection orbit of the celestial body from the Earth to the target celestial body and the orbit determination data of the Deep Space Network, and then numerical integration is performed on the detector to obtain the running trajectory parameters of the detector in the station coordinate system, so that the orbit parameters of the detector can be accurately determined according to the running trajectory parameters, the coordinate position of the station, and the initial coordinate position of the detector, thereby solving the technical problem in the related art that the orbit parameters of the detector cannot be accurately determined.

[0147] Optionally, the processing unit includes: a first initialization module, configured to perform dynamic model initialization, data type initialization of orbit determination data, and orbit determination arc segment initialization based on the basic detection orbit of the celestial body and the orbit determination data of the Deep Space Network, so as to complete the initialization processing of the orbit determination information, where the dynamic model is used to analyze the change of dynamic parameters in each flight stage during the trajectory process of the detector flying from the Earth to the target celestial body.

[0148] Optionally, the determination device further includes: a first conversion module, configured to obtain the coordinate positions of each station after obtaining the basic detection orbit of the celestial body from the Earth to the target celestial body and the orbit determination data of the Deep Space Network, and perform geodetic coordinate system conversion and station horizon system conversion based on the coordinate positions of the stations; a second conversion module, configured to obtain the current coordinate position of the detector, and perform conversion on the coordinate reference system used by the detector in each flight stage based on the current coordinate position of the detector.

[0149] Optionally, the integration unit includes: a first analysis module configured to analyze the current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters of the detector during each flight phase from the Earth to the target star; a first calculation module configured to perform numerical integration on the current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters to calculate the position and velocity of the detector in the station coordinate system and obtain the operating trajectory parameters.

[0150] Optionally, the determination unit includes: a first determination module configured to perform a first-order Taylor expansion on the observation equation of the detector to determine the correlation between the systematic errors of the orbiter, the ascender, and the observation, and establish a linear regression equation set; a second calculation module configured to use the linear regression equation set to calculate the theoretical observation, observation partial derivative, and state partial derivative of the detector; a third calculation module configured to substitute the theoretical observation, observation partial derivative, and state partial derivative into the expanded observation equation to calculate the observation data of the detector; a second determination module configured to determine the orbit parameters of the detector in combination with the observation data, the weight parameter of the observation data, the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector.

[0151] Optionally, the determination device further includes: a first acquisition module configured to, after determining the orbit parameters of the detector based on the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector, acquire the data residuals for calculating the orbit parameters of the detector during each flight phase, where the data residuals include at least one of the following: ranging residual, velocity measurement residual, time delay residual, and time delay rate residual; a second acquisition module configured to acquire the to-be-traveled orbit parameters of the orbit parameters; a first evaluation module configured to evaluate the to-be-traveled orbit parameters by using the overlapping arc segment method in combination with the data residuals to obtain an orbit evaluation result, where the orbit evaluation result is used to indicate whether the accuracy of the orbit parameters of the detector is within a preset accuracy error range.

[0152] Optionally, the target star includes: Mars.

[0153] The above determination device may further include a processor and a memory. The above acquisition unit 80, processing unit 82, integration unit 84, determination unit 86, etc. are all stored in the memory as program units, and the processor executes the above program units stored in the memory to implement corresponding functions.

[0154] The above processor includes a kernel, and the kernel retrieves the corresponding program unit from the memory. One or more kernels may be provided, and the orbit parameters of the detector are determined by adjusting the kernel parameters.

[0155] The above-mentioned memory may include non-permanent memory in the form of computer-readable media, random access memory (RAM) and / or non-volatile memory, such as read-only memory (ROM) or flash RAM, and the memory includes at least one memory chip.

[0156] The present application also provides a computer program product, which is adapted to execute a program initialized with the following method steps when executed on a data processing device: obtaining a basic detection orbit of a star from the Earth to a target star and deep space network orbit measurement data, initializing the orbit measurement information based on the basic detection orbit of the star and the deep space network orbit measurement data, performing numerical integration on the detector after completing the initialization process of the orbit measurement information to obtain the operating trajectory parameters of the detector in the station coordinate system, and determining the orbit parameters of the detector based on the operating trajectory parameters, the station coordinate position, and the initial coordinate position of the detector.

[0157] According to another aspect of the embodiments of the present invention, there is also provided an electronic device, including: a processor; and a memory for storing executable instructions of the processor; wherein the processor is configured to execute the above-mentioned orbit determination method of the detector by executing the executable instructions.

[0158] According to another aspect of the embodiments of the present invention, there is also provided a computer-readable storage medium, which includes a stored computer program, wherein when the computer program runs, it controls the device where the computer-readable storage medium is located to execute the above-mentioned orbit determination method of the detector.

[0159] The serial numbers of the above-mentioned embodiments of the present invention are only for description and do not represent the advantages and disadvantages of the embodiments.

[0160] In the above embodiments of the present invention, the descriptions of the various embodiments have their own emphases. For the parts not detailed in a certain embodiment, reference may be made to the relevant descriptions of other embodiments.

[0161] In several embodiments provided by the present application, it should be understood that the disclosed technical content can be implemented in other ways. Among them, the device embodiments described above are only illustrative. For example, the division of the units can be a logical function division, and there can be other division methods in actual implementation. For example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the displayed or discussed couplings or direct couplings or communication connections to each other can be through some interfaces, and the indirect couplings or communication connections of the units or modules can be in an electrical or other form.

[0162] The unit described as a separation component may or may not be physically separated. The component shown as a unit may or may not be a physical unit, that is, it may be located in one place or distributed over multiple units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0163] In addition, each functional unit in various embodiments of the present invention can be integrated in a processing unit, or each unit can exist physically alone, or two or more units can be integrated in one unit. The above-mentioned integrated unit can be implemented in the form of hardware or in the form of a software functional unit.

[0164] If the above-mentioned integrated unit is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in a computer-readable storage medium. Based on such an understanding, the technical solution of the present invention, in essence, or the part that makes a contribution to the prior art, or all or part of this technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions for causing a computer device (which can be a personal computer, a server, or a network device, etc.) to execute all or part of the steps of the methods described in various embodiments of the present invention. The foregoing storage medium includes: USB flash drives, read-only memories (ROMs), random access memories (RAMs), mobile hard disks, magnetic disks, or optical discs and other various media that can store program codes.

[0165] The above is only the preferred embodiment of the present invention. It should be noted that for those of ordinary skill in the art, without departing from the principle of the present invention, several improvements and refinements can be made, and these improvements and refinements should also be regarded as the protection scope of the present invention.

Claims

1. A method for determining the orbit of a detector, characterized in that, Including: Obtaining the basic detection orbit of the star from the Earth to the target star and the deep space network orbit measurement data; Based on the basic detection orbit of the star and the deep space network orbit measurement data, initializing the orbit measurement information; After completing the initialization processing of the orbit measurement information, performing numerical integration on the detector to obtain the operating trajectory parameters of the detector in the measurement station coordinate system, including: analyzing the current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters of the detector during each flight stage from the Earth to the target star; performing numerical integration on the current position, current velocity, solar radiation pressure coefficient, and attitude control jet parameters to calculate the position and velocity of the detector in the measurement station coordinate system, and obtaining the operating trajectory parameters; Based on the operating trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector, determining the orbit parameters of the detector, including: performing a first-order Taylor expansion on the observation equation of the detector to determine the correlation relationship between the systematic errors of the orbiter, the ascender, and the observation, and establishing a linear regression equation set; using the linear regression equation set to calculate the theoretical observation quantity, observation partial derivative, and state partial derivative of the detector; substituting the theoretical observation quantity, observation partial derivative, and state partial derivative into the expanded observation equation to calculate the observation data of the detector; combining the observation data, the weight parameter of the observation data, the operating trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector to determine the orbit parameters of the detector.

2. The orbit determination method according to claim 1, characterized in that The step of initializing the orbit measurement information based on the basic detection orbit of the star and the deep space network orbit measurement data includes: Based on the basic detection orbit of the star and the deep space network orbit measurement data, performing initialization of the dynamic model, initialization of the data type of the orbit measurement data, and initialization of the orbit measurement arc segment to complete the initialization processing of the orbit measurement information, where the dynamic model is used to analyze the change of the dynamic parameters of the detector during each flight stage in the trajectory process from the Earth to the target star.

3. The orbit determination method according to claim 1, characterized in that After obtaining the basic detection orbit of the star from the Earth to the target star and the deep space network orbit measurement data, it further includes: Obtaining the coordinate positions of each measurement station, and based on the coordinate positions of the measurement stations, performing geodetic coordinate system conversion and measurement station horizon coordinate system conversion; Obtaining the current coordinate position of the detector, and based on the current coordinate position of the detector, performing conversion of the coordinate reference system used by the detector during each flight stage.

4. The orbit determination method according to claim 1, characterized in that, After determining the orbit parameters of the detector based on the operating trajectory parameters, the measurement station coordinate position, and the initial coordinate position of the detector, it further includes: Obtaining the data residuals for calculating the orbit parameters of the detector during each flight stage, where the data residuals include at least one of the following: ranging residual, velocity measurement residual, time delay residual, time delay rate residual; Obtaining the orbit parameters to be traveled in the orbit parameters; Combined with the data residuals, the parameters of the to-be-traveled orbit are evaluated by the overlapping arc segment method to obtain an orbit evaluation result, where the orbit evaluation result is used to indicate whether the accuracy of the orbit parameters of the detector is within a preset accuracy error range.

5. The orbit determination method according to any one of claims 1 to 4, characterized in that The target celestial body includes: Mars.

6. An orbital determination device for a detector, characterized in that, Including: An acquisition unit, configured to acquire the basic detection orbit of the celestial body from the Earth to the target celestial body and the deep space network orbit measurement data; A processing unit, configured to perform initialization processing on the orbit measurement information based on the basic detection orbit of the celestial body and the deep space network orbit measurement data; An integration unit, configured to perform numerical integration on the detector after completing the initialization processing of the orbit measurement information to obtain the operation trajectory parameters of the detector in the station coordinate system; A determination unit, configured to determine the orbit parameters of the detector based on the operation trajectory parameters, the station coordinate position, and the initial coordinate position of the detector; The integration unit includes: a first analysis module, configured to analyze the current position, current speed, solar radiation pressure coefficient, and attitude control jet parameter of the detector at each flight stage from the Earth to the target celestial body; a first calculation module, configured to perform numerical integration on the current position, current speed, solar radiation pressure coefficient, and attitude control jet parameter to calculate the position and speed of the detector in the station coordinate system, and obtain the operation trajectory parameters; The determination unit includes: a first determination module, configured to perform a first-order Taylor expansion on the observation equation of the detector, determine the correlation relationship between the systematic errors of the orbiter, the ascender, and the observation, and establish a linear regression equation set; a second calculation module, configured to use the linear regression equation set to calculate the theoretical observation quantity, observation partial derivative, and state partial derivative of the detector; a third calculation module, configured to substitute the theoretical observation quantity, observation partial derivative, and state partial derivative into the expanded observation equation to calculate the observation data of the detector; a second determination module, configured to determine the orbit parameters of the detector in combination with the observation data, the weight parameter of the observation data, the operation trajectory parameters, the station coordinate position, and the initial coordinate position of the detector.

7. An electronic device, characterized in that, Including: A processor; And A memory, configured to store the executable instructions of the processor; Wherein, the processor is configured to execute the orbit determination method of the detector according to any one of claims 1 to 5 by executing the executable instructions.

8. A computer-readable storage medium, characterized in that, The computer-readable storage medium includes a stored computer program, wherein when the computer program runs, it controls the device where the computer-readable storage medium is located to execute the orbit determination method of the detector according to any one of claims 1 to 5.

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