An autonomous orbit determination method and system for lunar spacecraft based on DRO relay

Abstract

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

Description

Technical Field

[0001] This invention belongs to the field of spacecraft navigation and control technology, and specifically relates to an autonomous orbit determination method and system for lunar orbiting spacecraft based on DRO relay, which utilizes the special orbital characteristics of the Earth-Moon space and reduces reliance on ground telemetry and control. Background Technology

[0002] As lunar exploration activities gradually shift from the early "lunar exploration project" to "long-term stay and development on the moon," deep space tracking, telemetry, and command (STDM) and navigation technologies face unprecedented challenges. The unique lunar orbital environment, obstructed line-of-sight between the Earth and the Moon, and limited mission coverage necessitate increasingly higher demands for continuous, high-precision, and multi-target navigation and communication for future lunar scientific research, sampling, manned, and unmanned autonomous operations.

[0003] Currently, lunar spacecraft orbit determination and navigation mainly rely on the following two traditional technologies: The first method is direct tracking based on a ground-based tracking network. This involves using deep-space tracking stations on Earth's surface (such as radio telescopes and radar) to measure the distance and velocity of lunar spacecraft. However, this method has significant limitations: 1. High construction and maintenance costs: The construction of ground-based deep space tracking and control stations is extremely complex, and the operation and maintenance costs are enormous.

[0004] 2. Severe resource bottleneck: Ground station resources are limited, making it difficult to meet the tracking needs of the lunar constellation system, which will operate in parallel with multiple tasks and for a long period of time in the future.

[0005] 3. Limited coverage: Due to the Earth-Moon geometry and the Moon's own obstruction, ground stations cannot continuously track spacecraft operating on the far side of the Moon or in some polar orbits.

[0006] The second approach is to establish a lunar satellite navigation system similar to Earth's GNSS. Currently, several lunar communication and navigation constellation plans have been proposed internationally (such as ESA's Moonlight project and NASA's LCRNS project), attempting to deploy navigation satellites in lunar orbit. However, unlike near-Earth GNSS systems, the construction of a lunar navigation system faces unique challenges: 1. The near-Earth model cannot be directly applied: High-precision ephemeris data for near-Earth GNSS satellites primarily relies on globally distributed ground monitoring stations for all-weather tracking and data injection. However, it is difficult to deploy a high-precision ground monitoring network on a large scale on the lunar surface.

[0007] 2. High Challenges in Autonomous Orbit Determination: Due to the lack of continuous ground support, lunar orbiting satellites must possess high-precision autonomous orbit determination capabilities. Existing autonomous navigation technologies typically rely on inter-satellite links, but in the absence of a highly stable spatial reference, the overall orbital error of the constellation will diverge over time.

[0008] 3. High cost of orbit maintenance: Conventional lunar orbits are severely affected by the non-spherical gravitational perturbation of the Moon. Maintaining the orbit requires a large amount of fuel and is difficult to use as a long-term stable navigation reference.

[0009] Therefore, how to reduce dependence on ground-based telemetry and control resources and find an autonomous orbit determination method with natural stability that can provide a high-precision space reference for other lunar orbiting spacecraft has become a pressing technical challenge in building a lunar communication and navigation constellation system. Summary of the Invention

[0010] In response to the problems in existing technologies, such as the heavy reliance on ground-based telemetry and control resources for lunar exploration missions, the difficulty in maintaining traditional lunar navigation references, and the inability of single-satellite relay orbit determination accuracy to meet the requirements of high-precision missions, this invention provides an autonomous orbit determination method and system for lunar spacecraft based on DRO relay, so as to at least solve the problems of difficulty in maintaining navigation references for a long time, bottleneck of ground-based telemetry and control resources, and limited orbit determination accuracy of single links.

[0011] According to one aspect of the present invention, an autonomous orbit determination method for a lunar orbiting spacecraft based on DRO relay is provided, comprising: A two-layer navigation constellation architecture consisting of a navigation reference layer and a user service layer is constructed, wherein the navigation reference layer is deployed with relay navigation stars operating in a long-distance retrograde orbit in the Earth-Moon space, and the user service layer is deployed with lunar user stars operating in lunar orbit; By using ground-based telemetry and control stations to conduct telemetry and control only on the relay navigation satellite, its precise ephemeris in the inertial frame can be calculated and used as a dynamic measurement base station in lunar orbit that does not rely on ground station tracking. The lunar user satellite is controlled to receive navigation signals broadcast by the relay navigation satellite, while simultaneously acquiring inter-satellite pseudorange observations and inter-satellite carrier phase observations; Using the precise ephemeris of the relay navigation satellite as a constraint, and combining the inter-satellite pseudorange observations and inter-satellite carrier phase observations, the orbital state parameters of the lunar user satellite are calculated using the built-in high-precision dynamic model and hierarchical parameter estimation algorithm without relying on ground stations to measure the distance to the lunar user satellite.

[0012] As a further technical solution, the orbital parameters of the relay navigation star are generated based on the target shooting algorithm and are configured to satisfy the periodic resonance characteristics and the Earth-Moon distance characteristics. At the same time, the relay navigation star is configured under these orbital parameters to maintain long-term orbital stability by utilizing the coupling effect of the gravitational forces of the Earth, Moon, and Earth.

[0013] As a further technical solution, the periodic resonance characteristic is that the orbital period and the lunar orbital period are in a resonant proportional relationship, and the Earth-Moon distance characteristic is that the distance between the perigee of the orbit and the center of the Moon is approximately 40,000 kilometers.

[0014] As a further technical solution, the lunar user satellite is configured to operate in a lunar elliptical frozen orbit, and the orbital parameters of the lunar user satellite satisfy dynamic constraints: , in, For orbital eccentricity, This represents the track inclination angle.

[0015] As a further technical solution, the hierarchical parameter estimation algorithm is performed for observation geometry configurations with only a single relay navigation satellite, and includes: An observation equation containing only the inter-satellite pseudorange observations is constructed, a first measurement noise weight at the meter level is set, and the initial orbit is obtained by solving the least squares method. Based on the initial orbit, an observation equation incorporating the inter-satellite carrier phase observations is constructed, a second measurement noise weight at the centimeter level is set, the search range of integer ambiguity is limited by the initial orbit, ambiguity is fixed, and the target orbit with sub-meter accuracy is obtained by calculation.

[0016] As a further technical solution, the dynamic model is configured as follows: For the aforementioned relay navigation satellite: the central celestial body is Earth, taking into account the non-spherical gravitational perturbation of Earth, the gravitational perturbation of the three bodies of the Moon, and the solar radiation pressure perturbation; For the aforementioned lunar user star: the central celestial body is the Moon, taking into account the non-spherical gravitational perturbation of the Moon and the gravitational perturbation of the Earth's three-body structure.

[0017] According to one aspect of the present invention, an autonomous orbit determination system for a lunar orbiter is provided, the system being constructed based on a two-layer navigation constellation architecture, comprising: The space dynamic reference node corresponds to a relay navigation satellite operating in a long-distance retrograde orbit in the Earth-Moon space. This node is equipped with a reference maintenance and broadcasting module, which is used to maintain its own precise ephemeris based on ground telemetry and control, and to broadcast navigation signals into lunar orbit. The autonomous orbit-determining user node corresponds to the lunar user satellite operating in lunar orbit. This node is configured with: The dual-mode observation module is used to simultaneously output inter-satellite pseudorange observation values ​​and inter-satellite carrier phase observation values ​​after receiving the navigation signal; The single-link orbit determination processor is configured to calculate an initial orbit using the inter-satellite pseudorange observations, and to search for and fix the integer ambiguity of the inter-satellite carrier phase observations using the initial orbit, thereby calculating its own orbit parameters with sub-meter accuracy based on the single-link measurement data.

[0018] As a further technical solution, the orbital parameters of the relay navigation star are generated based on the target shooting algorithm and are configured to satisfy the periodic resonance characteristics and the Earth-Moon distance characteristics. At the same time, the relay navigation star is configured under these orbital parameters to maintain long-term orbital stability by utilizing the coupling effect of the gravitational forces of the Earth, Moon, and Earth.

[0019] According to one aspect of the present invention, a computer-readable storage medium is provided having a computer program stored thereon, which, when executed by a processor, implements the described autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay.

[0020] According to one aspect of the present invention, an autonomous orbit determination device for a lunar orbiting spacecraft is provided, comprising a processor and a memory, wherein the memory stores a computer program, and the processor executes the computer program to implement the aforementioned autonomous orbit determination method for a lunar orbiting spacecraft based on DRO relay.

[0021] Compared with the prior art, the beneficial effects of the present invention are as follows: (1) A highly stable and low-cost space navigation reference was constructed. This invention utilizes the DRO orbit as a navigation reference. The relay navigation satellite is configured with orbital parameters that satisfy periodic resonance characteristics and Earth-Moon distance characteristics, and the long-term stability of the orbit is maintained by utilizing the coupling effect of the Earth-Moon gravitational forces. Compared with conventional Halo or NRHO relay orbits, the DRO orbit has natural long-term stability in the Earth-Moon gravitational field, which significantly reduces the fuel consumption and ground control frequency required for orbit maintenance, enabling it to serve as a stable "space lighthouse" to provide continuous service to lunar orbit.

[0022] (2) A highly efficient telemetry and control mode of "long baseline maintenance + short link transmission" has been achieved. This invention only requires the ground station to maintain the accuracy of the DRO satellite (long baseline), while any user satellite in lunar orbit only needs to receive the DRO signal (short link) to determine its orbit. This architecture avoids the resource congestion problem of the ground station directly tracking a large number of lunar orbit satellites, and greatly improves the overall efficiency of the lunar exploration system.

[0023] (3) It breaks through the bottleneck of high-precision orbit determination under single-satellite relay conditions. Although conventional carrier phase measurement technology is used, this invention combines the DRO high-precision benchmark with the hierarchical constraint algorithm. For the observation geometry configuration with only a single relay navigation satellite, the initial orbit is first calculated using pseudorange observations, and then the initial orbit is used to limit the search range of integer ambiguity. This solves the problem of fixing integer ambiguity under harsh geometric conditions such as single satellite, short arc segment, and no ground assistance. Thus, sub-meter level orbit determination accuracy is achieved at the system level, which can effectively support high-precision tasks such as lunar landing and rendezvous and docking.

[0024] (4) An optimized polar coverage solution is provided. In a preferred embodiment of the present invention, a lunar elliptical frozen orbit technology is incorporated, wherein the orbital parameters of the lunar user satellite satisfy dynamic constraints. The DRO orbit provides wide-area background field coverage, while the lunar elliptical frozen orbit utilizes dynamic freezing characteristics to achieve high-frequency coverage of the lunar polar regions with low maintenance costs. The combination of the two forms a complementary and complete lunar navigation constellation configuration. Attached Figure Description

[0025] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the accompanying drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the accompanying drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0026] Figure 1 A flowchart illustrating an autonomous orbit determination method for a lunar orbiting spacecraft based on DRO relay, provided in an embodiment of the present invention; Figure 2 This is a schematic diagram of the lunar constellation autonomous orbit determination system provided in an embodiment of the present invention; Figure 3 A schematic diagram of the lunar DRO orbital family provided for embodiments of the present invention; Figure 4 This is a schematic diagram of a CLFO lunar orbiter provided in an embodiment of the present invention; Figure 5 This is a schematic diagram of the residual distribution of precise orbit determination during the pseudorange navigation stage provided in an embodiment of the present invention; Figure 6 This is a schematic diagram of the precise orbit determination residual distribution during the carrier phase navigation stage provided in an embodiment of the present invention; Figure 7 This is a schematic diagram illustrating lunar surface navigation and positioning accuracy under different operating conditions, provided for embodiments of the present invention. Detailed Implementation

[0027] The terms “comprising” and “having”, and any variations thereof, in the specification, claims, and accompanying drawings of this invention are intended to cover a non-exclusive inclusion, such as a process, method, system, product, or apparatus that includes a series of steps or units, not necessarily limited to those explicitly listed, but may include other steps or units not explicitly listed or inherent to such processes, methods, products, or apparatus.

[0028] This invention aims to provide an autonomous orbit determination method and system for lunar orbiting spacecraft based on DRO relay, in order to solve the following main technical problems: (1) Solve the problem of long-term maintenance of navigation reference: overcome the shortcomings of traditional lunar orbits (such as Halo and NRHO) that require frequent orbit maintenance and have complex dynamic environments, and build a space reference with low maintenance cost; (2) Solve the bottleneck problem of ground-based telemetry and control resources: reduce the dependence of lunar orbiting spacecraft on all-weather tracking by the ground telemetry and control network, and realize the efficient telemetry and control architecture of "single station manages single satellite, single satellite manages the whole moon"; (3) Solve the problem of limited orbit determination accuracy of single link: Under the limited observation conditions of lack of ground observation assistance and only a single relay satellite, solve the coupling problem between orbit state and measurement deviation, and realize sub-meter level high-precision autonomous orbit determination.

[0029] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention. In addition, the technical features of the various embodiments or individual embodiments provided by the present invention can be arbitrarily combined to form new technical solutions. Such combinations are not bound by the order of steps and / or structural composition patterns, but must be based on the ability of those skilled in the art to implement them. When the combination of technical solutions is contradictory or cannot be implemented, it should be considered that such a combination of technical solutions does not exist and is not within the scope of protection claimed by the present invention.

[0030] This invention provides an autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay, such as... Figure 1 As shown, it includes the following steps: Step S1: Construct a two-layer navigation constellation architecture; construct a lunar orbital space navigation system consisting of a navigation reference layer and a user service layer; wherein, the navigation reference layer is deployed with relay navigation stars operating in a distant retrograde orbit (DRO) in Earth-Moon space, and the user service layer is deployed with lunar orbital user stars operating in lunar orbit; Step S2: Establish a high-precision single-link benchmark; use ground control stations to perform telemetry and control only on the DRO relay navigation satellite, and calculate its precise ephemeris in the inertial frame; utilize the long-term stability of the DRO orbit to use the relay navigation satellite as a dynamic measurement base station for tracking user satellites in lunar orbit without relying on ground stations. Step S3: Acquire inter-satellite observation data with enhanced carrier phase; control the lunar user satellite to receive navigation signals broadcast by the relay navigation satellite, and simultaneously collect inter-satellite pseudorange observations and inter-satellite carrier phase observations; Step S4: Short-segment high-precision autonomous orbit determination calculation; Using the precise ephemeris of the relay navigation satellite as a constraint, combined with the inter-satellite pseudorange observations and inter-satellite carrier phase observations, the orbital state parameters of the lunar user satellite are calculated using the built-in high-precision dynamic model and hierarchical parameter estimation algorithm without relying on ground stations to measure the distance to the lunar user satellite.

[0031] Optionally, in step S2, the tracking and measurement duration of the relay navigation satellite by the ground telemetry and control station is configured to be approximately 7 days; in step S4, the duration of the inter-satellite observation data arc used for orbit determination is configured to be approximately 0.5 days.

[0032] Optionally, the orbital parameters of the relay navigation satellite are generated based on a target-shooting algorithm and configured to satisfy the following spatial geometric constraints to maintain baseline stability: the orbital period is approximately 7 days, exhibiting a resonant proportional relationship with the lunar orbital period, i.e., periodic resonance characteristics; the distance parameter between the orbital perigee and the center of the moon... The distance is approximately 40,000 kilometers, which corresponds to the Earth-Moon distance. The relay navigation satellite is configured with these orbital parameters, utilizing the gravitational coupling of the Earth, Moon, and Earth to maintain long-term orbital stability.

[0033] To further reduce the interference of user satellite orbit maintenance on the continuity of autonomous orbit determination, preferably, the lunar user satellite is configured to operate in a lunar elliptical frozen orbit (CLFO); its orbital parameters satisfy the following dynamic constraints: , where e is the orbital eccentricity and i is the orbital inclination.

[0034] To balance freezing stability and lunar polar coverage, the specific orbital parameters of the lunar user satellite can be configured as follows: orbital inclination i approximately 120 degrees, eccentricity e approximately 0.76, orbital period approximately 24 hours, and perigee argument... It is approximately 90 degrees.

[0035] Preferably, for short-arc observation conditions of single-satellite relay, the following coarse-to-fine orbit determination strategy is adopted: Initial orbit determination sub-step: Construct observation equations that contain only the inter-satellite pseudorange observations, set a first measurement noise weight at the meter level (e.g., 10 meters), and obtain an initial orbit that meets the linearization requirements by solving the least squares method; Precise sperm determination steps: Based on the initial orbit, construct an observation equation that includes the inter-satellite carrier phase observation values, set a second measurement noise weight at the centimeter level (e.g., 1 cm); use the initial orbit to limit the search range of integer ambiguity, complete the ambiguity fixation, and calculate to obtain the target orbit with sub-meter accuracy.

[0036] Preferably, the dynamic model is configured as follows: For the aforementioned relay navigation satellite: the central celestial body is Earth, taking into account the non-spherical gravitational perturbation of Earth, the gravitational perturbation of the three bodies of the Moon, and the solar radiation pressure perturbation; For the aforementioned lunar user star: the central celestial body is the Moon, taking into account the non-spherical gravitational perturbation of the Moon and the gravitational perturbation of the Earth's three-body system; The relay navigation satellite and the lunar user satellite both adopt the solar radiation pressure perturbation model, and the surface-to-mass ratio parameters are configured as 0.02 and 0.01, respectively.

[0037] To make the objectives, technical solutions, and advantages of this invention clearer, the technical solutions of this invention will be further described in detail below with reference to the accompanying drawings, mathematical models, and specific simulation parameters.

[0038] This embodiment constructs as follows: Figure 2 The illustrated two-layer navigation constellation architecture includes relay navigation satellites (DRO satellites) located in the navigation reference layer and lunar user satellites located in the user service layer.

[0039] I. Dynamic Environment and Orbit Parameter Settings 1. Dynamics Model Configuration To realistically simulate the complex gravitational environment of the Earth-Moon space, a high-precision dynamic model was established in the orbit determination solver in this embodiment. This model comprehensively considers the gravity of the central celestial body, multi-body perturbations, non-spherical gravity, solar radiation pressure, and relativistic effects. Specific force model configuration parameters are shown in Table 1.

[0040] Table 1 Detailed Configuration of Dynamic Model .

[0041] 2. Orbital Parameter Design and Ephemeris The simulation epoch is set to January 1, 2030 (UTC).

[0042] DRO Navigation Satellite: Under a high-precision dynamic model, the distance parameters between the DRO orbital perigee and the Moon are adjusted. By combining this with a target-shooting algorithm, the DRO orbital family under different resonance ratio conditions can be obtained, such as... Figure 3 As shown. Based on comprehensive factors such as long-term orbital stability and navigation signal broadcasting distance, this embodiment selects the perigee and lunar distance parameters. The period is approximately 7 days. The orbit remains stable in the Earth-Moon rotating coordinate system during the integration time.

[0043] Lunar User Satellite: This embodiment selects a CLFO frozen orbit, and the orbital parameters meet the following requirements. ,like Figure 4 As shown.

[0044] The specific initial orbital element configuration is shown in Table 2.

[0045] Table 2 Initial Roots of Simulated Track .

[0046] II. Measurement Model and Observation Equations This embodiment involves two types of measurement links, and the following mathematical models are established for each: 1. Ground-DRO telemetry and control link (two-way Doppler) The ground station measures the velocity of the DRO satellite using a two-way integral Doppler model. Its observation equations can be expressed as: , in, and These represent the two-way optical travel time distance between the ground station and the satellite at the end and start times of the observation count, respectively, and T is the counting interval (set to 60 seconds in this embodiment). For measuring noise (in this embodiment, it is set to white noise of 0.1 mm / s), This represents the rate of change of distance.

[0047] 2. DRO - User Inter-Satellite Link (Pseudorange and Carrier Phase) The user satellite receives navigation signals broadcast by the DRO satellite, which include two types of observations: pseudorange and carrier phase.

[0048] Pseudorange observation equation: , Carrier phase observation equation: , in, Let c be the geometric distance between satellites and c be the speed of light. and Clock biases for user satellites and DRO satellites, respectively. Where is the carrier wavelength, and N is the integer ambiguity. and These are the measurement noise terms for pseudorange and carrier phase, respectively.

[0049] Noise parameter settings: pseudorange noise standard deviation Carrier phase noise standard deviation .

[0050] III. Simulation Strategy and Orbit Determination Process To verify the autonomous orbit determination performance, the following simulation strategy was designed: 1. Design of observation arc segment Baseline Maintenance Phase: Ground tracking stations will conduct long-term tracking of the DRO navigation satellite for a total duration of 7 days. Considering station visibility, the continuous tracking time per day will be set at 8 hours (8h / day), with no ground observations during the remaining time. Observation data will be Doppler data integrated over 60 seconds.

[0051] Autonomous orbit determination segment: The lunar user satellite receives signals from the DRO navigation satellite. To verify the rapid orbit determination capability, only a short arc segment of data (0.5 days, 12 hours) is used for orbit determination calculation. The observation data consists of 1Hz pseudorange and carrier phase observations.

[0052] 2. Initial Error Settings (Critical Configuration) To simulate a real engineering environment, artificial errors were applied to the initial orbit values ​​of the user satellite before orbit determination began: Position error: applied in three axial directions The prior position deviation is km.

[0053] This setting is used to verify the algorithm's convergence ability when the initial deviation is large.

[0054] 3. Hierarchical solution strategy The processor performs the following two steps of calculation: Step A (Initial Orbit Determination): Using only pseudorange observations The least squares method was used to reduce the orbital position error from 10km to about 30m, providing a reliable initial value for fixing the ambiguity.

[0055] Step B (Precise Orbit Determination): Introduce carrier phase observations Using the orbit obtained in step A as a priori constraint, sub-meter orbit convergence is achieved using carrier phase observations. Specifically, the carrier phase cycle deviation corresponding to an initial position error of 30 meters is within a finite range. By searching for integer solutions within this range and combining them with the minimization criterion of detection quantities (such as the sum of squared residuals), the correct integer ambiguity N can be quickly locked, thereby improving the ranging accuracy from the meter level of pseudorange to the millimeter / centimeter level of carrier.

[0056] IV. Analysis of Implementation Results Based on the above configuration, numerical simulations were performed. This paper evaluates the orbit determination accuracy by assessing the difference (absolute error) between the reconstructed orbit and the nominal orbit. The results are as follows: Table 3. Numerical results of orbit determination simulation for DRO navigation satellite and lunar user satellite. .

[0057] The results show that the DRO navigation satellite orbit determination accuracy of long-term Doppler telemetry tracking (7 days of data) by the ground station is in the meter range, with an accuracy of 8.901m in the pseudorange navigation phase and 4.306m in the carrier phase navigation phase.

[0058] Under the joint orbit determination framework adopted in this study, based on navigation signals broadcast by the DRO navigation satellite and 0.5 days of inter-satellite pseudorange measurement data, the average autonomous orbit determination accuracy of the lunar user satellite is approximately 30 meters. However, with 0.5 days of inter-satellite carrier measurement data, the autonomous orbit determination error of the lunar user satellite is better than 0.5 meters, achieving sub-meter level accuracy. The results indicate that the inter-satellite ranging link formed by the DRO navigation satellite and the lunar user satellite has the potential to achieve high-precision autonomous navigation in a lunar environment, especially with carrier phase involvement, which can improve the orbit determination accuracy of the lunar user satellite by more than an order of magnitude.

[0059] Figure 5 and Figure 6 The weighted residual distributions of two types of data before precise orbit determination (Prefit), after precise orbit determination (Postfit), and at the reference level are presented, reflecting the error evolution during the orbit determination process. Here, "Ranging" represents the pseudorange / carrier phase ranging data between the DRO navigation satellite and the lunar user satellite, "InstantRangeRate" represents the Doppler telemetry tracking data of the DRO navigation satellite from the ground station, N represents the data volume of each type of observation, and Wmean and WRMS represent the weighted average and weighted root mean square value of the residuals, respectively. It can be observed that the orbit determination process significantly reduces the residuals of both types of observations. Specifically, the Wmean±WRMS of the inter-satellite navigation ranging and Doppler telemetry data are consistent with the reference ephemeris, with a deviation of less than 0.01, reaching the accuracy level of the reference ephemeris. This indicates that the joint orbit determination results in the simulation experiment statistically achieve the accuracy of the reference orbit, verifying the applicability and reliability of the dynamic model, measurement model, and the autonomous orbit determination method of the lunar constellation.

[0060] V. Performance Evaluation of Lunar Navigation Service To further verify the service capability of the high-precision satellite orbit determined by the present invention for lunar users, this embodiment further evaluates the accuracy of lunar navigation and positioning.

[0061] 1. Evaluation Model Based on the above calculation of the lunar user satellite orbit error ( ) and ranging error ( Construct a single-point positioning error model on the lunar surface: , Among them, GDOP (Geometric Precision Factor) reflects the geometric configuration quality of the satellite for lunar users.

[0062] 2. Evaluation Results According to the sub-meter level orbital accuracy obtained in Example 1 ( ), combined with ranging noise ( The lunar surface positioning accuracy under different GDOP values ​​was simulated (m).

[0063] like Figure 7 As shown, under poor geometric configuration (GDOP=10), if the satellite orbital error is large (such as 30m for traditional pseudorange orbit determination), the lunar surface positioning error will be as high as 300m or more.

[0064] After applying this invention, the satellite orbital error is reduced to within 0.5m, and even under the unfavorable configuration of GDOP=10, the lunar surface positioning accuracy can be maintained at around 5 meters; when the geometry is good (GDOP=2), the lunar surface positioning accuracy can reach the level of 1 meter (m level).

[0065] The results show that a single ground station can achieve orbit determination accuracy on the order of ten meters using 7 days of telemetry and control data from the DRO navigation satellite. Furthermore, the navigation and ranging link between the DRO navigation satellite and the lunar constellation, in the initial pseudorange navigation phase, can achieve autonomous orbit determination accuracy on the order of 30 meters for a 0.5-day link segment. When carrier phase navigation becomes available, the autonomous orbit determination accuracy of the lunar constellation satellites can be improved to the sub-meter level, thus enabling meter-level navigation and positioning accuracy for lunar surface operations. The full-process autonomous orbit determination analysis method and numerical analysis results for the lunar constellation based on DRO navigation signals established in this study can provide theoretical support and technical reference for the design and mission implementation of future lunar communication and navigation constellation systems.

[0066] Based on the same inventive concept as the foregoing embodiments, this invention also provides an autonomous orbit determination system for lunar spacecraft based on DRO relay. The system is constructed based on a two-layer navigation constellation architecture, including a space dynamic reference node and an autonomous orbit determination user node.

[0067] The space dynamic reference node corresponds to a relay navigation satellite operating in a long-distance retrograde orbit in the Earth-Moon space. This node is equipped with a reference maintenance and broadcasting module, which is used to maintain its own precise ephemeris based on ground telemetry and control, and to broadcast navigation signals into lunar orbit.

[0068] Specifically, the ground control station only tracks and controls the relay navigation satellite, calculating its precise ephemeris in the inertial frame. In this embodiment, the ground control station tracks and measures the relay navigation satellite for 7 days, with 8 hours of continuous tracking per day, and the observation data is Doppler data. The calculated precise ephemeris position accuracy of the relay navigation satellite reaches the meter level.

[0069] In a preferred embodiment, the orbital parameters of the relay navigation satellite are generated based on a target-shooting algorithm and configured to satisfy periodic resonance characteristics and Earth-Moon distance characteristics. Specifically, the perigee of the orbit is approximately 40,000 kilometers from the center of the Moon, and the orbital period is in a resonant proportional relationship with the Moon's orbital period around the Earth (approximately 7 days). With these orbital parameters, the relay navigation satellite maintains long-term orbital stability by utilizing the gravitational coupling effect of the Earth, Moon, and Sun, significantly reducing fuel consumption and the frequency of ground control required for orbit maintenance.

[0070] The autonomous orbit determination user node corresponds to a lunar user satellite operating in lunar orbit. This node is equipped with a dual-mode observation module and a single-link orbit determination processor.

[0071] In a preferred embodiment, the lunar user satellite is configured to operate in a lunar elliptical frozen orbit, with orbital parameters satisfying dynamic constraints. Specifically, the orbital inclination is approximately 120 degrees, the eccentricity is approximately 0.76, the orbital period is approximately 24 hours, and the perigee argument is approximately 90 degrees. This orbit utilizes lunar non-spherical gravitational perturbation to achieve orbit freezing, reducing interference with the continuity of autonomous orbit determination while simultaneously enabling high-frequency coverage of the lunar polar regions.

[0072] The dual-mode observation module is used to simultaneously output inter-satellite pseudorange observation values ​​and inter-satellite carrier phase observation values ​​after receiving the navigation signal.

[0073] In this embodiment, the measurement noise standard deviation of the pseudorange observation is set to 10 meters, and the measurement noise standard deviation of the carrier phase observation is set to 1 centimeter.

[0074] The single-link orbit determination processor is configured to execute a hierarchical solution strategy. Specifically, the processor performs the following two steps of solution: The first step is to calculate the initial orbit using the inter-satellite pseudorange observations. An observation equation containing only inter-satellite pseudorange observations is constructed, and a first measurement noise weight (10 meters) at the meter level is set. The initial orbit is then calculated using the least squares method. In this embodiment, with an initial position error of ±10 km, the orbital position error can be converged to approximately 30 meters.

[0075] The second step involves calculating a target orbit with sub-meter accuracy based on the initial orbit and the inter-satellite carrier phase observations. An observation equation incorporating the inter-satellite carrier phase observations is constructed, and a second measurement noise weight (1 cm) at the centimeter level is set. The initial orbit is used to limit the search range for integer ambiguities, thus fixing the ambiguities and calculating the target orbit with sub-meter accuracy. In this embodiment, the three-dimensional position and orbit determination accuracy of the lunar user satellite is better than 0.5 meters.

[0076] The single-link orbit determination processor incorporates a high-precision dynamic model: For the relay navigation satellite: the central celestial body is Earth, and the non-spherical gravitational perturbation of Earth (20th order Earth gravity field model), the gravitational perturbation of the three bodies of the Moon, and the solar radiation pressure perturbation (area-to-mass ratio 0.02) are considered. For lunar user stars: the central celestial body is the Moon, considering the non-spherical gravitational perturbation of the Moon (100th order lunar gravity field model) and the gravitational perturbation of the Earth's three-body system.

[0077] Based on the same inventive concept as any of the foregoing embodiments, this embodiment of the invention also provides a computer-readable storage medium storing a computer program thereon, which, when executed by a processor, implements the aforementioned autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay.

[0078] Specifically, the computer-readable storage medium can be any tangible medium that contains or stores a program, including but not limited to: read-only memory (ROM), random access memory (RAM), magnetic disk, optical disk, flash memory, solid-state drive (SSD), etc. The computer program is written in one or more programming languages, including high-level programming languages ​​(such as C, C++, Python, Java, etc.) and / or assembly language.

[0079] When the computer program is executed by the processor, the following steps are performed: A two-layer navigation constellation architecture consisting of a navigation reference layer and a user service layer is constructed. The navigation reference layer is deployed with relay navigation stars operating in a long-distance retrograde orbit in the Earth-Moon space, and the user service layer is deployed with lunar user stars operating in lunar orbit. By using ground-based telemetry and control stations to conduct telemetry and control only on the relay navigation satellite, its precise ephemeris in the inertial frame can be calculated and used as a dynamic measurement base station in lunar orbit that does not rely on ground station tracking. The lunar user satellite is controlled to receive navigation signals broadcast by the relay navigation satellite, while simultaneously acquiring inter-satellite pseudorange observations and inter-satellite carrier phase observations; Using the precise ephemeris of the relay navigation satellite as a constraint, and combining the inter-satellite pseudorange observations and inter-satellite carrier phase observations, the orbital state parameters of the lunar user satellite are calculated using the built-in high-precision dynamic model and hierarchical parameter estimation algorithm without relying on ground stations to measure the distance to the lunar user satellite.

[0080] As a preferred embodiment, the hierarchical parameter estimation algorithm is executed for observation geometry configurations with only a single relay navigation satellite, and includes: constructing an observation equation containing only inter-satellite pseudorange observations, setting a first measurement noise weight at the meter level, and obtaining an initial orbit by solving using the least squares method; based on the initial orbit, constructing an observation equation containing inter-satellite carrier phase observations, setting a second measurement noise weight at the centimeter level, using the initial orbit to limit the search range of integer ambiguity, completing ambiguity fixation, and solving to obtain a target orbit with sub-meter accuracy.

[0081] Based on the same inventive concept as any of the foregoing embodiments, this embodiment of the invention also provides an autonomous orbit determination device for a lunar orbiting spacecraft, including a processor and a memory. The memory stores a computer program, and when the processor executes the computer program, it implements the aforementioned autonomous orbit determination method for a lunar orbiting spacecraft based on DRO relay.

[0082] Specifically, the device may be a dedicated orbit determination computer on a spacecraft, part of an integrated electronic unit, or a ground test device.

[0083] Processor: Employs a radiation-hardened high-performance processor, such as an aerospace-grade processor based on the SPARC V8 architecture (e.g., ATMEL AT697F) or an ARM Cortex-R series processor (e.g., Cortex-R52). The processor's clock speed is no less than 100MHz, and it supports floating-point arithmetic units to meet the computational requirements of real-time orbit determination.

[0084] Memory: This includes program memory and data memory. Program memory uses non-volatile memory (such as PROM, EEPROM, or Flash) to store the computer program that implements the orbit determination method. Data memory uses SRAM or SDRAM to store intermediate data, observation data, and orbital state parameters during program execution.

[0085] When the processor executes a computer program in memory, it implements the following functional modules: Reference Maintenance Module: Utilizes data from ground-based telemetry and control stations that only monitor and control the relay navigation satellite to calculate the precise ephemeris of the relay navigation satellite in the inertial frame.

[0086] Observation data acquisition module: controls the lunar user satellite to receive navigation signals broadcast by the relay navigation satellite, and simultaneously acquires inter-satellite pseudorange observations and inter-satellite carrier phase observations.

[0087] The hierarchical orbit determination module, constrained by the precise ephemeris of the relay navigation satellite, combines inter-satellite pseudorange observations and inter-satellite carrier phase observations, and utilizes a built-in high-precision dynamic model and hierarchical parameter estimation algorithm to calculate the orbital state parameters of the lunar user satellite. When executing this module for observational geometry configurations with only a single relay navigation satellite, the specific steps are as follows: Initial orbit determination submodule: Constructs observation equations that contain only inter-satellite pseudorange observations, sets a first measurement noise weight at the meter level, and obtains the initial orbit by solving the least squares method; Precision Accuracy Submodule: Based on the initial orbit, an observation equation including inter-satellite carrier phase observations is constructed, a second measurement noise weight at the centimeter level is set, the search range of integer ambiguity is limited by the initial orbit, ambiguity is fixed, and the target orbit with sub-meter accuracy is obtained by calculation.

[0088] Dynamics model configuration module: Built-in high-precision dynamics models, including geocentric dynamics models for relay navigation satellites and lunar-centric dynamics models for lunar user satellites.

[0089] After the device is powered on, the processor loads the orbit determination program from the program memory and initializes it. During operation: The system receives telemetry and control data from the ground station for the relay navigation satellite via the telemetry and control interface and calculates the ephemeris of the relay navigation satellite. The pseudorange and carrier phase observation values ​​output by the dual-mode observation module are received through the navigation receiver interface; Run the graded orbit determination calculation program to calculate its own orbit status in real time; The orbit determination results are output to the spacecraft platform or other payloads via a data interface. As needed, the orbit determination results are transmitted via telemetry data through the telemetry and control interface.

[0090] In summary, this invention discloses an autonomous orbit determination method and system for lunar orbiting spacecraft based on DRO relay, relating to the field of deep space exploration and navigation technology. The method first constructs a two-layer navigation architecture consisting of a long-distance retrograde orbit (DRO) navigation satellite and a lunar orbiting user satellite. Ground-based tracking stations track the DRO navigation satellite to maintain its high-precision ephemeris, using it as a space reference node. The DRO navigation satellite broadcasts navigation signals to the lunar orbiting user satellite, which receives the signals and acquires pseudorange and carrier phase observation data for the inter-satellite link. Finally, based on the high-precision ephemeris and observation data, the lunar orbiting user satellite uses a weighted estimation algorithm to calculate its own orbital state. This invention utilizes the long-term stability and high coverage characteristics of the DRO orbit, significantly reducing the dependence of lunar exploration missions on ground-based tracking and control resources. Simultaneously, by fusing carrier phase observations, it achieves sub-meter-level high-precision autonomous orbit determination for lunar orbiting spacecraft, providing a high-precision space reference for lunar surface positioning and navigation.

[0091] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features therein; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the technical solutions of the embodiments of the present invention.

Claims

1. A method for autonomous orbit determination of a lunar orbiter based on DRO relay, characterized in that, include: A two-layer navigation constellation architecture consisting of a navigation reference layer and a user service layer is constructed, wherein the navigation reference layer is deployed with relay navigation stars operating in a long-distance retrograde orbit in the Earth-Moon space, and the user service layer is deployed with lunar user stars operating in lunar orbit; By using ground-based telemetry and control stations to conduct telemetry and control only on the relay navigation satellite, its precise ephemeris in the inertial frame can be calculated and used as a dynamic measurement base station in lunar orbit that does not rely on ground station tracking. The lunar user satellite is controlled to receive navigation signals broadcast by the relay navigation satellite, while simultaneously acquiring inter-satellite pseudorange observations and inter-satellite carrier phase observations; Using the precise ephemeris of the relay navigation satellite as a constraint, and combining the inter-satellite pseudorange observations and inter-satellite carrier phase observations, the orbital state parameters of the lunar user satellite are calculated using the built-in high-precision dynamic model and hierarchical parameter estimation algorithm without relying on ground stations to measure the distance to the lunar user satellite.

2. The DRO-relay-based autonomous orbit determination method for a circumlunar spacecraft according to claim 1, wherein, The orbital parameters of the relay navigation satellite are generated based on a target-shooting algorithm and are configured to satisfy periodic resonance characteristics and Earth-Moon distance characteristics. At the same time, the relay navigation satellite is configured under these orbital parameters to maintain long-term orbital stability by utilizing the gravitational coupling effect of the Earth-Moon tri-body system.

3. The autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay according to claim 2, characterized in that, The periodic resonance characteristic is that the orbital period and the lunar orbital period are in a resonant proportional relationship, and the Earth-Moon distance characteristic is that the distance between the perigee of the orbit and the center of the Moon is approximately 40,000 kilometers.

4. The autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay according to claim 1, characterized in that, The lunar user satellite is configured to operate in a lunar elliptical frozen orbit, and the orbital parameters of the lunar user satellite satisfy the following dynamic constraints: ， in, For orbital eccentricity, This represents the track inclination angle.

5. The autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay according to claim 1, characterized in that, The hierarchical parameter estimation algorithm is performed for observation geometry configurations with only a single relay navigation satellite, and includes: An observation equation containing only the inter-satellite pseudorange observations is constructed, a first measurement noise weight at the meter level is set, and the initial orbit is obtained by solving the least squares method. Based on the initial orbit, an observation equation incorporating the inter-satellite carrier phase observations is constructed, a second measurement noise weight at the centimeter level is set, the search range of integer ambiguity is limited by the initial orbit, ambiguity is fixed, and the target orbit with sub-meter accuracy is obtained by calculation.

6. The autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay according to claim 1, characterized in that, The dynamic model is configured as follows: For the aforementioned relay navigation satellite: the central celestial body is Earth, taking into account the non-spherical gravitational perturbation of Earth, the gravitational perturbation of the three bodies of the Moon, and the solar radiation pressure perturbation; For the aforementioned lunar user star: the central celestial body is the Moon, taking into account the non-spherical gravitational perturbation of the Moon and the gravitational perturbation of the Earth's three-body structure.

7. An autonomous orbit determination system for a lunar orbiter, characterized in that, The system is built on a two-layer navigation constellation architecture and includes: The space dynamic reference node corresponds to a relay navigation satellite operating in a long-distance retrograde orbit in the Earth-Moon space. This node is equipped with a reference maintenance and broadcasting module, which is used to maintain its own precise ephemeris based on ground telemetry and control, and to broadcast navigation signals into lunar orbit. The autonomous orbit-determining user node corresponds to the lunar user satellite operating in lunar orbit. This node is configured with: The dual-mode observation module is used to simultaneously output inter-satellite pseudorange observation values ​​and inter-satellite carrier phase observation values ​​after receiving the navigation signal; The single-link orbit determination processor is configured to calculate an initial orbit using the inter-satellite pseudorange observations, and to search for and fix the integer ambiguity of the inter-satellite carrier phase observations using the initial orbit, thereby calculating its own orbit parameters with sub-meter accuracy based on the single-link measurement data.

8. The autonomous orbit determination system for a lunar orbiter according to claim 7, characterized in that, The orbital parameters of the relay navigation satellite are generated based on a target-shooting algorithm and are configured to satisfy periodic resonance characteristics and Earth-Moon distance characteristics. At the same time, the relay navigation satellite is configured under these orbital parameters to maintain long-term orbital stability by utilizing the gravitational coupling effect of the Earth-Moon tri-body system.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that, When executed by the processor, the program implements the autonomous orbit determination method for lunar spacecraft based on DRO relay as described in any one of claims 1 to 6.

10. An autonomous orbit determination device for a lunar orbiting spacecraft, characterized in that, It includes a processor and a memory, the memory storing a computer program, and the processor executing the computer program to implement the autonomous orbit determination method for lunar orbiting spacecraft based on DRO relay as described in any one of claims 1 to 6.

Images