A lunar-earth space user aircraft navigation and positioning system

Abstract

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

Description

Technical Field

[0001] The present invention relates to the technical field of satellite navigation, and in particular to a navigation and positioning system for a user aircraft in the Earth-Moon space. Background Art

[0002] When positioning using satellite navigation technology, the accuracy of the navigation ephemeris orbit parameters directly affects the user positioning accuracy. The orbit determination accuracy of navigation satellites and the expression accuracy of ephemeris parameters are the two main factors affecting the accuracy of navigation ephemeris orbit parameters, and the expression accuracy of ephemeris parameters is closely related to the broadcast ephemeris expression model (navigation ephemeris parameter model). Different satellite navigation systems have different broadcast ephemeris expression models. Traditional satellite navigation systems such as GPS, GALILEO, and BD use a 16- or 18-parameter navigation ephemeris parameter model based on satellite orbital elements. For users in the near-Earth space, traditional satellite navigation technology can meet all high-precision navigation and positioning requirements above the Earth's surface. However, for Earth-Moon system navigation, only using satellite navigation systems orbiting the Earth, the geometric configuration of user relative to navigation satellites for positioning is poor, and the positioning accuracy decays severely, making it difficult to meet the requirements. Therefore, the 16-parameter or 18-parameter navigation ephemeris parameter model used by Earth orbit navigation satellites is not applicable to libration point navigation satellites. Summary of the Invention

[0003] The purpose of the present invention is to provide a navigation and positioning system for a user aircraft in the Earth-Moon space, which can improve the navigation and positioning accuracy and is more suitable for libration point navigation satellites.

[0004] To achieve the above object, the present invention provides the following solutions:

[0005] A navigation and positioning system for a user aircraft in the Earth-Moon space, comprising:

[0006] A user aircraft in the Earth-Moon space and a navigation satellite system connected to the user aircraft in the Earth-Moon space, the navigation satellite system comprising a plurality of navigation satellites; each navigation satellite is equipped with an on-board data processing payload; the navigation satellite is a libration point navigation satellite or a circumlunar navigation satellite; inter-satellite ranging measurements are sent between each navigation satellite;

[0007] The on-board data processing payload carried by each navigation satellite includes:

[0008] A self-observation quantity correction module, configured to correct the antenna phase center of the navigation satellite according to the satellite attitude of the navigation satellite to obtain the antenna phase center correction quantity of the navigation satellite, and correct the inter-satellite ranging measurement sent by the target navigation satellite received by the navigation satellite according to the antenna phase center correction quantity of the navigation satellite to obtain the corrected inter-satellite ranging observation quantity corresponding to the target navigation satellite sent by the navigation satellite; the target navigation satellite is a navigation satellite other than the navigation satellite in the navigation satellite system;

[0009] An error correction module, configured to perform error correction on the inter-satellite ranging amount sent by the target navigation satellite received by the navigation satellite based on the inter-satellite ranging correction observation amount sent by the target navigation satellite corresponding to the navigation satellite and the inter-satellite ranging correction observation amount sent by the navigation satellite corresponding to the target navigation satellite, so as to obtain the corrected inter-satellite ranging amount sent by the target navigation satellite received by the navigation satellite;

[0010] An inter-satellite distance observation equation construction module, configured to obtain the inter-satellite distance observation equation of the navigation satellite to the target navigation satellite according to the corrected inter-satellite ranging amount sent by the target navigation satellite received by the navigation satellite and the corrected inter-satellite ranging amount sent by the navigation satellite received by the target navigation satellite;

[0011] An estimation module, configured to obtain the optimal estimations of the dynamic navigation ephemeris parameters of the navigation satellite and the correction amount of the inter-satellite measurement system deviation parameters based on the inter-satellite distance observation equation of the navigation satellite to the target navigation satellite and the dynamic equation of the navigation satellite by using an optimal parameter estimation method;

[0012] The Earth-Moon space user aircraft includes:

[0013] An inter-satellite ranging information determination module, configured to obtain the inter-satellite ranging information between the Earth-Moon space user aircraft and each navigation satellite according to the optimal estimations of the dynamic navigation ephemeris parameters of each navigation satellite and the correction amount of the inter-satellite measurement system deviation parameters;

[0014] A position and velocity determination module, configured to perform numerical integration on the dynamic equation of each navigation satellite according to the dynamic navigation ephemeris parameters of each navigation satellite to obtain the position and velocity of each navigation satellite;

[0015] A user aircraft orbit determination observation equation determination module, configured to obtain a user aircraft orbit determination observation equation according to the positions and velocities of each navigation satellite, the optimal estimated values of the inter-satellite ranging system deviation parameter correction amounts of each navigation satellite, and the ranging information between the Earth-Moon space user aircraft and each navigation satellite;

[0016] A user aircraft state transition equation determination module, configured to obtain a user aircraft state transition equation according to the dynamic equation of the Earth-Moon space user aircraft;

[0017] A positioning module, configured to implement navigation positioning of the Earth-Moon space user aircraft through iterative processing by using a parameter optimization estimation method according to the user aircraft state transition equation and the user aircraft orbit determination observation equation.

[0018] According to the specific embodiments provided by the present invention, the present invention discloses the following technical effects:

[0019] The present invention proposes a method for a translunar space user aircraft navigation and positioning system, which utilizes libration point navigation satellites to autonomously generate and broadcast the optimal estimates of satellite dynamic state parameters (the optimal estimate of the position correction amount of the navigation satellite, the optimal estimate of the velocity correction amount, and the optimal estimate of the dynamic model parameter correction amount) and the optimal estimate of the clock error parameter correction amount at a specific moment as dynamic navigation ephemeris parameters. The translunar space user receives the navigation signal and the dynamic navigation ephemeris parameters to achieve navigation and positioning, avoiding the problem that the traditional navigation ephemeris parameter model is not applicable to libration point navigation satellites. Brief Description of the Drawings

[0020] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the embodiments. Obviously, the drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can be obtained based on these drawings.

[0021] Figure 1 It is a schematic diagram for deploying libration point navigation satellites;

[0022] Figure 2 It is a specific working flowchart of the translunar space user aircraft navigation and positioning system provided by the embodiment of the present invention. Detailed Embodiments

[0023] The following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, rather than all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0024] In order to make the above objects, features, and advantages of the present invention more obvious and understandable, the present invention will be further described in detail below with reference to the drawings and specific embodiments.

[0025] The embodiment of the present invention provides a translunar space user aircraft navigation and positioning system, including:

[0026] A translunar space user aircraft and a navigation satellite system connected to the translunar space user aircraft. The navigation satellite system includes a plurality of navigation satellites; each navigation satellite is equipped with an on-board data processing payload; the navigation satellite is a libration point navigation satellite or a circumlunar navigation satellite; the navigation satellites send inter-satellite ranging measurements to each other. The specific working process of the translunar space user aircraft navigation and positioning system is as Figure 2 shown, and the on-board data processing payloads carried by each of the navigation satellites all include:

[0027] The self-observed quantity correction module is used to correct the antenna phase center of the navigation satellite according to the satellite attitude of the navigation satellite to obtain the antenna phase center correction quantity of the navigation satellite, and correct the inter-satellite ranging quantity sent by the target navigation satellite received by the navigation satellite according to the antenna phase center correction quantity of the navigation satellite, so as to obtain the inter-satellite ranging corrected observed quantity sent by the target navigation satellite corresponding to the navigation satellite; the target navigation satellite is a navigation satellite other than the navigation satellite in the navigation satellite system.

[0028] The error correction module is used to perform error correction on the inter-satellite ranging quantity sent by the target navigation satellite received by the navigation satellite based on the inter-satellite ranging corrected observed quantity sent by the target navigation satellite corresponding to the navigation satellite and the inter-satellite ranging corrected observed quantity sent by the navigation satellite corresponding to the target navigation satellite, so as to obtain the corrected inter-satellite ranging quantity sent by the target navigation satellite received by the navigation satellite.

[0029] The inter-satellite distance observation equation construction module is used to obtain the inter-satellite distance observation equation of the navigation satellite to the target navigation satellite according to the corrected inter-satellite ranging quantity sent by the target navigation satellite received by the navigation satellite and the corrected inter-satellite ranging quantity sent by the navigation satellite received by the target navigation satellite.

[0030] The estimation module is used to obtain the optimal estimations of the dynamic navigation ephemeris parameters of the navigation satellite and the correction quantity of the inter-satellite measurement system deviation parameters based on the inter-satellite distance observation equation of the navigation satellite to the target navigation satellite and the dynamic equation of the navigation satellite by using the optimal parameter estimation method. The optimal estimation of the dynamic navigation ephemeris parameter correction quantity specifically includes: the optimal estimation of the position correction quantity of the navigation satellite, the optimal estimation of the velocity correction quantity, the optimal estimation of the dynamic model parameter correction quantity, and the optimal estimation of the clock error parameter correction quantity. The dynamic model parameter is initially considered as 1 light pressure model parameter. The signal modulated with the optimal estimations of the dynamic navigation ephemeris parameters of the navigation satellite and the correction quantity of the inter-satellite measurement system deviation parameters is the navigation signal.

[0031] The Earth-Moon space user aircraft includes:

[0032] The inter-satellite ranging information determination module is used to obtain the inter-satellite ranging information between the Earth-Moon space user aircraft and each navigation satellite according to the optimal estimations of the dynamic navigation ephemeris parameters of each navigation satellite and the correction quantity of the inter-satellite measurement system deviation parameters.

[0033] A position and velocity determination module is used to numerically integrate the dynamic equations of each of the navigation satellites based on the dynamic ephemeris parameters of each of the navigation satellites to obtain the positions and velocities of each of the navigation satellites. Specifically, satellite orbit numerical integration software can be used, and a fifth-order Runge-Kutta integrator can be adopted. Using a second-order polynomial clock error model, the clock error of the navigation satellite at any time can be calculated.

[0034] A user aircraft orbit determination observation equation determination module is used to obtain a user aircraft orbit determination observation equation based on the positions and velocities of each of the navigation satellites, the optimal estimated value of the correction amount of the inter-satellite ranging system deviation parameter of each of the navigation satellites, and the ranging information between the Earth-Moon space user aircraft and each of the navigation satellites. The user aircraft orbit determination observation equation also needs to perform measurement system errors such as relativity, antenna phase center, and ranging system deviation on the inter-satellite ranging observation quantity, and normalize the measurement time scale, which is the same as the previous navigation satellite processing steps.

[0035] A user aircraft state transition equation determination module is used to obtain a user aircraft state transition equation based on the dynamic equation of the Earth-Moon space user aircraft.

[0036] A positioning module is used to, based on the user aircraft state transition equation and the user aircraft orbit determination observation equation, adopt a parameter optimization estimation method and perform iterative processing to obtain improved user aircraft position state quantities, velocity state quantities, clock error information, and parameter estimation error statistical results, so as to realize the navigation and positioning of the Earth-Moon space user aircraft. The Earth-Moon space user aircraft cycles in the above order to achieve continuous navigation and positioning.

[0037] In practical applications, the self-observation quantity correction module includes:

[0038] A self-observation quantity correction unit is used to calculate the corrected inter-satellite ranging observation quantity δρ sent by the target navigation satellite j corresponding to the navigation satellite i according to the formula

[0039] where δρ i,ant , δρ i,ant is the correction value of the inter-satellite ranging observation quantity for the antenna phase center of the navigation satellite i. Among them, ρ ij (t) is the inter-satellite ranging quantity sent by the target navigation satellite j received by the navigation satellite i at time t, is the position vector of the navigation satellite i in the inertial coordinate system, is the position vector of the navigation satellite j in the inertial coordinate system, is the correction amount of the antenna phase center of the navigation satellite i.

[0040] In practical applications, the error correction module performs antenna phase center correction, relativistic correction, and ranging system deviation correction on the received inter-satellite ranging observables, specifically including:

[0041] An error correction unit, which is used to calculate the corrected inter-satellite ranging quantity sent by the target navigation satellite j received by the navigation satellite i at time t according to the formula

[0042]

[0043] where ρ (t) is the inter-satellite ranging quantity sent by the target navigation satellite j received by the navigation satellite i at time t, δρ ij is the inter-satellite ranging correction observable sent by the target navigation satellite j corresponding to the navigation satellite i, δρ i,ant is the inter-satellite ranging correction observable sent by the target navigation satellite i corresponding to the navigation satellite j, δρ j,ant is the relativistic correction of the inter-satellite ranging observable, δt rel is the satellite clock error of the navigation satellite i, δt i is the satellite clock error of the navigation satellite j, δb j is the inter-satellite ranging system deviation of the navigation satellite i, δb i is the inter-satellite ranging system deviation of the navigation satellite j. j

[0044] In practical applications, the corrected two-way inter-satellite ranging data is combined to form an inter-satellite distance observation equation formula, and the inter-satellite distance observation equation is:

[0045] where is the corrected inter-satellite ranging quantity sent by the target navigation satellite j received by the navigation satellite i, is the corrected inter-satellite ranging quantity sent by the target navigation satellite i received by the navigation satellite j, is the inter-satellite distance observable after eliminating the clock error.

[0046] In practical applications, the earth-moon space user aircraft navigation and positioning system further includes: a dynamic equation construction module for navigation satellites:

[0047] The dynamic equation construction module for navigation satellites includes:

[0048] A navigation satellite dynamic differential equation construction unit, which is used to construct the dynamic differential equation of the navigation satellite; in the inertial coordinate system, comprehensively considering the earth, the sun and the moon, the major planets, and the solar radiation pressure perturbation, the dynamic differential equation of the navigation satellite is constructed.

[0049] The dynamic differential equation of the navigation satellite and the variational equation corresponding to the dynamic differential equation of the navigation satellite are processed by using a numerical integration method to obtain the instantaneous position of the satellite of the navigation satellite and the state transition matrix;

[0050] The dynamic equation of the navigation satellite is obtained according to the instantaneous position of the satellite of the navigation satellite and the state transition matrix.

[0051] In practical applications, each of the navigation satellites carries an inter-satellite link payload, and communication between the navigation satellites is carried out through the inter-satellite link payload.

[0052] In practical applications, each of the navigation satellites carries a navigation signal generation payload, a narrow-beam navigation signal generation and transmission payload for external communication. The narrow-beam navigation signal generation and transmission payload is a narrow-beam antenna. The narrow-beam antenna can be used to transmit navigation signals modulated with the optimal estimates of the dynamic ephemeris parameters of the navigation satellite, the correction amount of the inter-satellite measurement system deviation parameters, and information such as the satellite mass to space vehicles in a specified range of the Earth-Moon space. Considering the limited number of space vehicles in the Earth-Moon system, the navigation signals can be transmitted point-to-point in the Ka or V band or transmitted using a narrow-beam antenna with relatively concentrated energy. The transmission method adopts a sequential cyclic transmission method for multiple Earth-Moon space vehicles. Therefore, it is necessary to pre-inject a beam cyclic transmission control strategy into the libration point vehicle. The Earth-Moon space user vehicle carries a navigation signal receiver and an atomic frequency standard. The receiver is a Ka inter-satellite link device for regularly receiving the navigation signals transmitted by the navigation satellite.

[0053] In practical applications, the navigation satellite carries an on-board atomic clock to provide a consistent time reference for the inter-satellite link payload and the navigation signal generation payload. The navigation satellite is powered by solar energy and needs to be orbit-controlled regularly to keep the orbit near the theoretical position.

[0054] In practical applications, such as Figure 1As shown in the figure, navigation satellites moving around the libration points are deployed near the four libration points L2, L3, L4, and L5 in the Earth-Moon space. The orbits of the libration point navigation satellites are Halo periodic orbits or Lissajous quasi-periodic orbits. Navigation satellites orbiting the Moon can also be deployed in lunar orbits, and the orbits of the lunar navigation satellites are elliptical orbits. The four navigation satellites complete two-way inter-satellite ranging and inter-satellite data exchange every 30 minutes; after each navigation satellite obtains the inter-satellite ranging data, the correction amount of the phase center of each antenna is calculated using the satellite attitude information; after correction, information such as the inter-satellite ranging observation amount, satellite clock error parameters, satellite mass, and antenna phase center correction is sent to other satellites through the inter-satellite link device. After completing a 30-minute measurement and communication cycle, each libration point satellite can obtain the corrected inter-satellite ranging observation amounts, clock error parameters, satellite masses, and information related to antenna phase center correction measurements between all libration point satellites. There are 5 Earth-Moon libration points. Together with the satellite in the lunar orbit, under ideal conditions, a single satellite can obtain 5 observation amounts, and an optimal estimate can be obtained by synthesizing the 5 observation amounts.

[0055] In practical applications, the optimal parameter estimation methods include optimal parameter estimation methods such as Kalman filtering and sequential least squares.

[0056] In practical applications, the satellite dynamic differential equation is:

[0057]

[0058]

[0059]

[0060] is the acceleration of satellite i, p are the dynamic model parameters such as the position, velocity, and light pressure parameters of satellite i at the initial moment, are the perturbation forces on the satellite by the Earth, the Moon, the Sun, the major planets, and the solar light pressure, respectively.

[0061] Integrating the above differential equation and the corresponding variational equation using numerical integration methods, the satellite position at the observation moment and its partial derivatives with respect to the dynamic parameters can be obtained. Using these partial derivatives, the satellite state transition matrix can be formed and the state transition equation can be generated:

[0062]

[0063]

[0064] where is the correction amount of the dynamic parameters, Φ t / t0 is the state transition matrix, and ΔZ0 is the correction amount of the dynamic parameters at time 0 (initial moment); For the satellite position correction at time 0, is the satellite velocity correction at any time (time 0, t), is the partial derivative of the satellite position vector at any time with respect to the satellite position vector at the initial time.

[0065] By synthesizing the state transition equation and the inter-satellite ranging observation equation, a dynamic equation is obtained. Optimal parameter estimation methods such as Kalman filtering or least squares are used to process the dynamic equation and the inter-satellite ranging observation equation to obtain the optimal estimates of the dynamic navigation ephemeris parameters and the correction amounts of the inter-satellite measurement system deviation parameters.

[0066] In practical applications, the process of constructing the dynamic equation of the Earth-Moon space user aircraft needs to consider the gravitational forces of the sun, the moon, the Earth, and the major planets, as well as the solar radiation pressure perturbation force of the user aircraft, which is similar to the process of constructing the dynamic equation of the navigation satellite.

[0067] In practical applications, the user aircraft state transition equation determination module is specifically:

[0068] The state transition matrix of the Earth-Moon space user aircraft can be obtained by integrating the variational equation of the dynamic equation of the Earth-Moon space user aircraft using numerical integration methods, and the user aircraft state transition equation can be calculated using the state transition matrix.

[0069] The present invention has the following technical effects:

[0070] The present invention proposes a method for the Earth-Moon space user to receive navigation signals and dynamic navigation ephemeris parameters to achieve navigation and positioning by using the libration point navigation satellite to autonomously generate and broadcast the satellite dynamic state parameters and clock difference parameters at a specific time as the dynamic navigation ephemeris parameters, avoiding the problem that the traditional navigation ephemeris parameter model is not applicable to the libration point navigation satellite.

[0071] The present invention effectively solves the problem that the 16-parameter or 18-parameter navigation ephemeris parameter model adopted by the traditional satellite navigation system cannot be applied to non-elliptical orbit satellites such as libration point navigation satellites, providing an alternative solution for the construction of the Earth-Moon space satellite navigation system.

[0072] The dynamic navigation ephemeris parameters proposed by the present invention only include very few parameters such as the reference time, satellite position, velocity, mass, and dynamic model parameters. On the premise of ensuring the orbit accuracy of the navigation ephemeris, the data volume of the navigation ephemeris parameters to be transmitted is greatly reduced, the orbit prediction accuracy of the dynamic model parameters is high, and the data transmission frequency can be further reduced.

[0073] The present invention takes into account the characteristics of the on-board data processing capabilities of libration point navigation satellites and Earth-Moon space vehicles, transfers some data processing tasks to the Earth-Moon space vehicle, reduces the need for the on-board data processing capabilities of libration point navigation satellites, and has certain universality for the ephemeris parameter design of navigation satellites with similar usage scenarios.

[0074] The dynamic model parameters adopted in this application have the following advantages. First, the precise orbit dynamic model of the libration point is directly used to calculate the numerical orbit, avoiding the problems of large errors in the analytical solution model of the libration point orbit and complex model expression forms. Second, the amount of data of the dynamic navigation ephemeris parameters is relatively small, the requirement for the data transmission frequency is low, and the inter-satellite data transmission burden is reduced. At the same time, the dynamic navigation ephemeris parameters have a data structure similar to the data transmitted by the inter-satellite link for autonomous orbit determination of libration point navigation satellites based on inter-satellite measurements, which can simplify the inter-satellite data transmission protocol. Third, directly broadcasting the dynamic navigation ephemeris parameters does not require operations such as orbit prediction and ephemeris fitting at the libration point navigation satellite end, reducing the data processing load burden of the libration point orbit navigation satellite. The present invention first completely proposes a method and technical approach for serving the navigation of Earth-Moon space vehicles by broadcasting dynamic navigation ephemeris parameters through libration point navigation satellites. The method has certain universality for the ephemeris parameter design of navigation satellites.

[0075] The various embodiments in this specification are described in a progressive manner. Each embodiment focuses on the differences from other embodiments. The same or similar parts among the various embodiments can be referred to each other. For the system disclosed in the embodiments, since it corresponds to the method disclosed in the embodiments, the description is relatively simple, and the relevant parts can be referred to the description of the method part.

[0076] Specific examples are used in this article to elaborate on the principles and implementation manners of the present invention. The descriptions of the above embodiments are only used to help understand the method and its core idea of the present invention. At the same time, for those of ordinary skill in the art, there will be changes in the specific implementation manners and application scopes according to the idea of the present invention. In summary, the content of this specification should not be construed as a limitation to the present invention.

Claims

1. A lunar-earth space user aircraft navigation and positioning system, characterized in that, Comprising: A user spacecraft in the Earth-Moon space and a navigation satellite system connected to the user spacecraft in the Earth-Moon space, the navigation satellite system including a plurality of navigation satellites; each navigation satellite is equipped with an on-board data processing payload; The navigation satellites are libration point navigation satellites or lunar orbit navigation satellites; Each navigation satellite sends inter-satellite ranging measurements to each other; The on-board data processing payloads carried by each of the navigation satellites all include: A self-observation quantity correction module, configured to correct the antenna phase center of the navigation satellite according to the satellite attitude of the navigation satellite to obtain the antenna phase center correction quantity of the navigation satellite, and correct the inter-satellite ranging measurement sent by the target navigation satellite received by the navigation satellite according to the antenna phase center correction quantity of the navigation satellite, to obtain the corrected inter-satellite ranging observation quantity sent by the target navigation satellite corresponding to the navigation satellite; the target navigation satellite is a navigation satellite other than the navigation satellite in the navigation satellite system; An error correction module, configured to perform error correction on the inter-satellite ranging measurement sent by the target navigation satellite received by the navigation satellite based on the corrected inter-satellite ranging observation quantity sent by the target navigation satellite corresponding to the navigation satellite and the corrected inter-satellite ranging observation quantity sent by the navigation satellite corresponding to the target navigation satellite, to obtain the corrected inter-satellite ranging measurement sent by the target navigation satellite received by the navigation satellite; An inter-satellite distance observation equation construction module, configured to obtain the inter-satellite distance observation equation of the navigation satellite to the target navigation satellite according to the corrected inter-satellite ranging measurement sent by the target navigation satellite received by the navigation satellite and the corrected inter-satellite ranging measurement sent by the navigation satellite received by the target navigation satellite; An estimation module, configured to obtain the optimal estimates of the dynamic navigation ephemeris parameters of the navigation satellite and the correction quantity of the inter-satellite measurement system deviation parameters based on the inter-satellite distance observation equation of the navigation satellite to the target navigation satellite and the dynamic equation of the navigation satellite, by using an optimal parameter estimation method; The user spacecraft in the Earth-Moon space includes: An inter-satellite ranging information determination module, configured to obtain the inter-satellite ranging information between the user spacecraft in the Earth-Moon space and each of the navigation satellites according to the optimal estimates of the dynamic navigation ephemeris parameters of each of the navigation satellites and the correction quantity of the inter-satellite measurement system deviation parameters; A position and velocity determination module, configured to perform numerical integration on the dynamic equations of each of the navigation satellites according to the dynamic navigation ephemeris parameters of each of the navigation satellites, to obtain the positions and velocities of each of the navigation satellites; A user spacecraft orbit determination observation equation determination module, configured to obtain the user spacecraft orbit determination observation equation according to the positions and velocities of each of the navigation satellites, the optimal estimated values of the inter-satellite ranging system deviation parameters of each of the navigation satellites, and the ranging information between the user spacecraft in the Earth-Moon space and each of the navigation satellites; A user spacecraft state transition equation determination module, configured to obtain the user spacecraft state transition equation according to the dynamic equation of the user spacecraft in the Earth-Moon space; A positioning module, which is used to realize the navigation and positioning of a user aircraft in the Earth-Moon space by using a parameter optimization estimation method through iterative processing according to the user aircraft state transition equation and the user aircraft orbit determination observation equation.

2. The lunar-earth space user aircraft navigation and positioning system according to claim 1, characterized in that, The self-observation quantity correction module includes: The self-observed quantity correction unit is used to calculate the inter-satellite ranging correction observed quantity δρ transmitted by the target navigation satellite j corresponding to the navigation satellite i according to the formula where ρ i,ant (t) is the inter-satellite ranging quantity transmitted by the target navigation satellite j received by the navigation satellite i at time t, ij is the position vector of the navigation satellite i in the inertial coordinate system, is the position vector of the navigation satellite j in the inertial coordinate system, and is the antenna phase center correction quantity of the navigation satellite i.

3. The lunar-earth space user aircraft navigation and positioning system according to claim 1, characterized in that, The error correction module includes: An error correction unit, which is used according to the formula Calculate the corrected inter-satellite ranging measurement sent by the target navigation satellite j received by the navigation satellite i at time t where ρ ij (t) is the inter-satellite ranging measurement sent by the target navigation satellite j received by the navigation satellite i at time t, δρ i,ant is the corrected inter-satellite ranging measurement observed sent by the target navigation satellite j corresponding to the navigation satellite i, δρ j,ant is the corrected inter-satellite ranging measurement observed sent by the target navigation satellite i corresponding to the navigation satellite j, δρ rel is the relativistic correction of the inter-satellite ranging measurement, δt i is the satellite clock error of the navigation satellite i, δt j is the satellite clock error of the navigation satellite j, δb i is the inter-satellite ranging system deviation of the navigation satellite i, δb j is the inter-satellite ranging system deviation of the navigation satellite j.

4. The lunar-earth space user aircraft navigation and positioning system according to claim 1, characterized in that, The inter-satellite distance observation equation is: Wherein, is the corrected inter-satellite ranging measurement sent by the target navigation satellite j received by the navigation satellite i, is the corrected inter-satellite ranging measurement sent by the target navigation satellite i received by the navigation satellite j, is the inter-satellite distance observation measurement after eliminating clock errors.

5. The lunar-earth space user aircraft navigation and positioning system according to claim 1, wherein It also includes: A dynamic equation construction module for navigation satellites: The dynamic equation construction module for navigation satellites includes: A navigation satellite dynamic differential equation construction unit, which is used to construct the dynamic differential equation of the navigation satellite; Using a numerical integration method to process the dynamic differential equation of the navigation satellite and the variational equation corresponding to the dynamic differential equation of the navigation satellite, to obtain the satellite instantaneous position and state transition matrix of the navigation satellite; Obtain the dynamic equation of the navigation satellite according to the satellite instantaneous position and state transition matrix of the navigation satellite.

6. The lunar-earth space user aircraft navigation and positioning system according to claim 1, characterized in that, Each of the navigation satellites is equipped with an inter-satellite link payload, and the navigation satellites communicate with each other through the inter-satellite link payload.

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