Design method for transferring a lunar orbiter to a geosynchronous orbit
By designing a three-pulse transfer framework and optimizing orbit control parameters, the orbit design problem of transferring the lunar probe to the Earth-Moon resonance orbit was solved, achieving efficient use of remaining propellant and meeting engineering constraints, thus improving the feasibility and adaptability of the mission.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- DEEP SPACE EXPLORATION LABORATORY
- Filing Date
- 2026-05-15
- Publication Date
- 2026-07-14
AI Technical Summary
Existing technologies lack a systematic orbit design approach to transfer lunar probes to a lunar resonance orbit, fail to make full use of the probe's remaining propellant, and fail to effectively consider engineering constraints and mission requirements.
A three-pulse transfer framework was designed, and key events were optimized through grid search and traversal methods. Combined with remaining propellant and telemetry distance, orbit control parameters and objectives were optimized. The adaptability and flight conditions of the transfer mission were analyzed to ensure that the orbit design meets complex engineering constraints and minimizes fuel consumption.
It enables efficient transfer of remaining propellant from the lunar orbiter, meets complex engineering constraints, improves the effectiveness and feasibility of the design, and provides technical support for lunar space missions.
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Figure CN122197109B_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of Earth-Moon space orbit design technology, and relates to a design method for transferring a lunar orbiter to an Earth-Moon resonance orbit. Background Technology
[0002] A lunar resonance orbit refers to an orbit in which the orbital period of a probe is in a specific integer ratio to the lunar orbital period. Such orbits are highly stable, requiring only small velocity increments to maintain their annual orbital position. Spacecraft operating in this orbit can achieve wide-area coverage and patrol of multiple nodes in space, including the Earth, the Moon, and the lunar libration point, with less fuel. This has significant application value for future missions such as resource development, manned lunar landings, and space science exploration.
[0003] Extensive research has been conducted on Earth-Moon resonance orbits. In 2011, the IBEX spacecraft moved from its original science orbit to a 3:1 lunar resonance orbit to study the interaction between the solar wind and the interstellar medium. The TESS spacecraft, launched in 2018, entered a 2:1 lunar resonance mission orbit through five maneuvers to support scientific observations.
[0004] Currently, there is no systematic research on the complete orbit design process for transferring from a lunar orbit to a lunar resonance orbit. This transfer orbit aims to make full use of the lunar orbiter that has completed its mission and has remaining propellant to conduct verification of new Earth-Moon space orbits, which can provide technical support for the future construction of the Earth-Moon space economic zone and Earth-Moon space situational awareness. Summary of the Invention
[0005] This invention provides a design method for transferring a lunar probe to a lunar resonance orbit. By calculating the probe's remaining propellant and maximum telemetry distance, the target orbit is determined to be a lunar space resonance orbit. A three-pulse transfer framework is designed, and the design is optimized for critical pulse events using grid search and traversal methods to obtain a transfer orbit that meets engineering constraints such as propellant, safety, and distance. Simultaneously, based on long-term recursive orbit results, it is determined whether the resonance orbit needs to maintain maneuvers to ensure mission requirements. Furthermore, the invention analyzes contingency situations such as potential delays in the first pulse execution time, the probe's adaptability to the transfer mission, and the probe's flight characteristics.
[0006] The technical solution of this invention includes:
[0007] A design method for transferring a lunar orbiter to a lunar resonance orbit includes:
[0008] Step S110: Determine the remaining propellant of the lunar orbiter;
[0009] Step S120: Determine the maximum distance for satellite-to-ground telemetry and control;
[0010] Step S130: Based on the remaining propellant of the lunar orbiter and the farthest distance of the satellite-to-ground telemetry and control system, design the transfer framework and transfer orbit control strategy. The transfer framework refers to the complete path architecture from the lunar orbit, through three pulse maneuvers, finally entering the resonance orbit and finally entering the atmosphere for destruction or escaping the Earth-Moon space. The transfer orbit control strategy refers to the specific provisions of the orbit control parameters and orbit control targets for each pulse in order to realize the above transfer framework.
[0011] Step S140: Optimize track control parameters and track control objectives;
[0012] Step S150: Analyze the maneuverability of maintaining the resonant orbit;
[0013] Step S160: Analyze the transfer task status.
[0014] Beneficial effects:
[0015] This invention discloses a design method for transferring a lunar orbiter to a lunar resonant orbit. It fully utilizes the remaining propellant of the lunar orbiter, designs the lunar orbit using a three-pulse method, and optimizes the design for key events through grid search and traversal methods to obtain a transfer orbit that meets complex engineering constraints and has low fuel consumption. Considering practical engineering aspects, the method analyzes factors such as delayed initial transfer time, probe adaptability, and probe flight behavior to further improve the effectiveness and feasibility of the design, providing technical support for future lunar-lunar space economic zone construction and lunar-lunar space situational awareness. Attached Figure Description
[0016] Figure 1 This is a schematic flowchart illustrating a design method for transferring a lunar probe to a lunar resonance orbit according to the present invention.
[0017] Figure 2 This is a diagram of the transfer trajectory in a geocentric rotating system.
[0018] Figure 3 This is a diagram of the migration trajectory within the geocentric ecliptic system.
[0019] Figure 4 This is a curve showing the change in the distance between the probe's orbit and the Earth's center over time.
[0020] Figure 5 This represents the closest distance between the detector and the GEO satellite. Detailed Implementation
[0021] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the disclosure to those skilled in the art.
[0022] This invention provides a design method for transferring a lunar orbiter to a lunar resonance orbit. This method makes full use of a lunar orbiter that has completed its mission and has remaining propellant to conduct verification of a new lunar orbit. The method achieves a complete orbit design process by transferring the lunar orbiter to a lunar resonance orbit and finally entering the atmosphere for destruction or escaping from the lunar space through three pulses. The method also includes analysis of the delayed transfer start time, maintenance of the resonance orbit, and the flight status of the probe.
[0023] Figure 1 A schematic flowchart of the method is shown. (e.g.) Figure 1 As shown, the method includes:
[0024] Step S110: Determine the remaining propellant of the lunar orbiter. This step may specifically include:
[0025] Step S110-1: Estimate the mass of remaining propellant using the pressure-volume-temperature (PVT) method.
[0026] According to the ideal gas equation, the tank pressure, tank temperature, and propellant mass have the following relationship:
[0027] ,
[0028] In the formula, The initial pressure of the tank after filling is completed, in MPa; The pressure inside the tank after a certain operating condition is completed, in MPa; The total volume of the storage tank is expressed in liters (L). The initial temperature of the storage tank, in K; The tank temperature after a certain operating condition is completed, in K; This refers to the amount of propellant added, expressed in kg. The remaining propellant is expressed in kg. Propellant density, unit: , For temperature The propellant density below, For temperature The propellant density below.
[0029] The mass of the remaining propellant can be calculated using the above formulas.
[0030] Step S110-2: Estimating the remaining propellant using the accounting method: By statistically analyzing the propellant consumption during each orbit control mission of the lunar probe, the mass of the remaining propellant is estimated.
[0031] Step S110-3: Calculate the remaining available propellant and velocity increment: Based on the propellant PVT method and accounting method estimation results, obtain the remaining available propellant amount according to the maximum envelope, and estimate the remaining velocity increment by combining the relationship between thruster specific impulse and thrust magnitude with the remaining fuel mass.
[0032] In one embodiment, considering a propulsion extrusion efficiency of 98.5%, pipeline residue of 0.1 kg, and subsequent attitude control usage of 0.1 kg, the remaining usable propellant amount is obtained based on the propellant PVT method and the accounting method, using the maximum envelope.
[0033] The remaining velocity increment is estimated by combining the relationship between the thruster specific impulse and the thrust magnitude with the remaining fuel mass, and a certain margin is considered as a constraint on the available velocity increment, which serves as the input for the transfer design.
[0034] Step S120: Determine the maximum distance for satellite-to-ground telemetry and control. This step may include:
[0035] Step S120-1: Analyze uplink capability: Calculate uplink capability based on telemetry data and corresponding satellite-to-ground distance, antenna angle, and ground station.
[0036] Uplink computational constraints: Onboard carrier acquisition, carrier tracking, and remote demodulation must meet margin requirements; to ensure the tonal regulation of the relay ranging, the onboard received signal level must be better than a certain value. The uplink distance of the selected station is calculated.
[0037] Step S120-2: Analyze bidirectional link capabilities.
[0038] The computational constraints for the two-way link are that the ground station acquisition and telemetry demodulation meet the margin requirements. The satellite-to-ground telemetry and control distance of the selected stations is calculated.
[0039] Step S120-3: Calculate the maximum telemetry and control distance: Considering uplink capability and bidirectional link capability, determine the maximum distance for satellite-to-ground telemetry and control. This process can be solved iteratively using known link budget techniques, which will not be elaborated here.
[0040] Step S130: Based on the remaining propellant of the lunar orbiter and the farthest distance of the satellite-to-ground telemetry and control system, design the transfer framework and transfer orbit control strategy. The transfer framework refers to the complete path structure from the lunar orbit, through three pulse maneuvers, finally entering the resonance orbit and finally entering the atmosphere for destruction or escaping the Earth-Moon space. The transfer orbit control strategy refers to the specific provisions of the orbit control parameters and orbit control targets for each pulse in order to realize the above transfer framework.
[0041] This step may specifically include:
[0042] Step S130-1: Define the design inputs and constraints for the transfer trajectory;
[0043] The design inputs for the transfer orbit include the remaining propellant on the lunar probe and the maximum distance for satellite-to-ground tracking and control.
[0044] The transfer orbit design constraints include: transfer start time constraint (the lunar probe will transfer after completing its predetermined mission); total velocity increment constraint (the transfer orbit with the optimal velocity increment); flight distance constraint (not exceeding the maximum telemetry and control distance); safety constraint (flying outside of geosynchronous orbit to avoid collision risks); and a complete orbit design (the probe will eventually re-enter the atmosphere and burn up or escape from Earth-Moon space). The total velocity increment constraint is the remaining velocity increment from step S110-3.
[0045] Step S130-2: Design the three-pulse transfer framework
[0046] The orbital transfer process is as follows: the lunar orbiter maneuvers and escapes in near-lunar space, enters a lunar resonance orbit, and eventually enters the atmosphere to be destroyed or escapes from the lunar space.
[0047] The three-pulse transfer framework is as follows: When the lunar probe reaches the perigee of its lunar orbit, the first pulse is applied, with the pulse direction along the positive tangent of the orbit. After maneuvering, the probe leaves the lunar orbit and, with the help of lunar gravity, lowers its perigee. When the probe reaches the target perigee, the second pulse is applied, with the pulse direction along the tangent of the orbit. After maneuvering, the probe enters a lunar resonance orbit. After the mission is completed, a third pulse is applied at the apogee, with the pulse direction along the negative tangent of the orbit, to lower the perigee of the orbit and eventually burn up in the atmosphere; or a third pulse is applied at the perigee, with the pulse direction along the positive tangent of the orbit, to raise the apogee of the orbit and eventually escape from the lunar space.
[0048] Step S130-3: Design a three-pulse track control strategy.
[0049] The track control parameters for the first pulse are:
[0050] First variable: the start time of the transfer, i.e. the moment when the near-lunar pulse is applied;
[0051] Second variable: the positive velocity increment of the perigee along the orbital tangential direction;
[0052] The goal of track control is:
[0053] First objective: the perigee radius upon reaching perigee;
[0054] Second objective: The angle between the probe, Earth, and the Moon at perigee.
[0055] The track control parameters for the second pulse are:
[0056] Third variable: velocity increment along the orbital tangential at perigee;
[0057] The goal of track control is:
[0058] Third objective: After multiple orbits, the perigee probe will reach the Earth-Moon angle.
[0059] The orbital control parameters for the third pulse's atmospheric reentry and destruction are:
[0060] Fourth variable: the velocity increment at the apogee along the tangential direction of the orbit in the opposite direction;
[0061] The goal of track control is:
[0062] Fourth objective: Reduce the perigee radius;
[0063] The orbital control parameters for the third pulse to escape Earth-Moon space are:
[0064] Fifth variable: The magnitude of the velocity increment along the positive tangential direction of the orbit at perigee;
[0065] The goal of track control is:
[0066] Fifth objective: Increase the apogee radius;
[0067] With three pulses, the number of variables is the same as the number of targets. The orbital control speed increment can be obtained by shooting at targets.
[0068] For the Earth-Moon resonance orbit in this mission, ideally, the angle between the perigee probe and the Earth and Moon for both orbit control targets should have an ideal nominal value.
[0069] Step S140: Optimize track control parameters and track control objectives. This step may include:
[0070] Step S140-1: Search transfer start time
[0071] The geocentric distance at perigee varies for the probe under different transfer start times and velocity increments, and this distance is used to characterize the distance constraint. A grid search method is used to recursively calculate the trajectory for both the transfer start time and velocity increment. The transfer start time variable is discretely assigned values at each perigee, and the search optimizes a set of transfer start times that satisfy the distance constraint. The transfer trajectory is then designed based on these transfer start times.
[0072] Step S140-2: Traverse the perigee radius of the target
[0073] The process involves iterating through the target's perigee radius, using the nominal value of the target's angle, obtaining the velocity increment required to enter the resonance orbit through target firing, and recursively calculating the change in the satellite-to-ground distance. The optimal target perigee radius is then selected. The optimal velocity increment and the satellite-to-ground distance change are then assessed to determine if they meet the distance and safety constraints. If they do, the orbital control parameter (the velocity increment value used for target firing calculation in the current iteration step) can be used to design the orbital transfer with the target. If not, the target's angle is further optimized through iteration.
[0074] Step S140-3: Traverse the target angles
[0075] The target angles are traversed, and the target perigee radius is the optimal target perigee radius selected in step S140-2. The velocity increment required to enter the resonance orbit is obtained by firing the target, and the orbit is recursively calculated to obtain the change in the satellite-to-ground distance. The traversal results are analyzed, and a set of target angles is selected to optimize the velocity increment and satisfy the distance constraint and safety constraint for the change in satellite-to-ground distance.
[0076] In summary, by optimizing the transfer start time, target perigee radius, and target angle, a set of track control parameters and targets that satisfy the design constraints are obtained.
[0077] Step S150: Analyze the resonant orbit maintenance maneuver.
[0078] Step S150-1: After the probe enters the Earth-Moon space resonance orbit, determine whether the resonance can be maintained during the subsequent flight process under the influence of gravity and perturbation, and whether the design constraints are still met.
[0079] Step S150-2: If the subsequent flight process can maintain resonance and meet the transfer trajectory design constraints, then maintenance maneuvers are not required;
[0080] Step S150-3: If the subsequent flight process cannot maintain resonance and meet the transfer trajectory design constraints, a maintenance maneuver is required;
[0081] Step S150-4: If the subsequent flight process cannot maintain resonance, but can meet the design constraints of the transfer trajectory, then decide whether to perform a maintenance maneuver based on mission requirements.
[0082] Step S150-5: If a maneuver is to be sustained, under the pulse model, the orbital control parameters for sustaining the maneuver are:
[0083] Third variable: velocity increment along the orbital tangential at perigee;
[0084] The orbit control objective is: Third objective: After multiple orbits, the perigee probe will reach the Earth-Moon angle.
[0085] Step S160: Analyze the transfer task status. This process may include:
[0086] Step S160-1: Analyze the delay in the transfer start time.
[0087] If some uncontrollable factors cause the first pulse to be delayed, analyze the orbital transfer situation when the first pulse is applied at the next lunar perigee, including whether it can enter the resonance orbit, whether the velocity increment requirement is met, and whether the transfer orbit design constraints are met. This analysis serves as a contingency plan to ensure the success of the mission.
[0088] Step S160-2: Analyze the detector's adaptability
[0089] We need to analyze whether the operational mode of this transfer mission can draw on previous missions. In fact, this transfer process is similar to the reverse process of the Earth-Moon transfer and capture of the Moon. We need to assess whether the current probe's individual components are functioning properly, whether its attitude is stable, and whether its communication and power systems are functioning correctly, in order to determine whether it possesses the capability and conditions to carry out the aforementioned transfer mission.
[0090] Step S160-3: Analyze the probe's flight status, including changes in orbital parameters, visibility of key event stations, and shadow conditions.
[0091] The following describes the specific implementation of the present invention in further detail using a set of Earth-Moon 3:1 resonance orbits as an example.
[0092] The initial orbital parameters are shown in Table 1:
[0093] Table 1 Initial orbit (Geocentric J2000 coordinate system)
[0094]
[0095] Step S110: Determine the remaining propellant of the lunar probe.
[0096] Substitute the initial pressure P1, initial temperature T1, and propellant loading amount at the launch site, along with the current tank pressure P2 and current temperature T2 based on the latest telemetry data, into the formula:
[0097] ,
[0098] The relationship between ρ and temperature is shown in the following formula:
[0099] ,
[0100] The remaining propellant mass estimated by the PVT method is 7.80 kg.
[0101] Based on the statistical analysis of past orbit control data, the estimated remaining propellant mass using the accounting method is 7.60 kg.
[0102] Based on the comprehensive estimation results, considering the maximum envelope, a propulsion extrusion efficiency of 98.5%, a pipeline residue of 0.1 kg, and a subsequent attitude control usage of 0.1 kg, the usable propellant mass is approximately 7.0 kg.
[0103] Step S120: Determine the farthest telemetry and control distance.
[0104] The maximum distance for satellite-to-ground telemetry and control is determined by the combined performance of the uplink and bidirectional links. Using an 18m station, the maximum satellite-to-ground telemetry and control distance can reach 450,000 kilometers.
[0105] Step S130: Based on the remaining propellant of the lunar probe and the farthest distance for satellite-to-ground telemetry and control, design the transfer framework and transfer orbit control strategy.
[0106] Step S140: Optimize track control parameters and track control objectives.
[0107] By searching the grid for the transfer start time and velocity increment, early March is optimally selected as the transfer start time. The radius and angle of the target's perigee are then optimized to satisfy the constraints.
[0108] The initial trajectory was pushed back to early March, which served as the initial time for the transfer maneuver.
[0109] The first pulse maneuver took place on March 9, with a pulse size of 107 m / s. At this time, the probe reached the lunar perigee of its lunar orbit and will escape the moon after the maneuver.
[0110] The second pulse maneuver took place on April 20, with a pulse size of 10 m / s. At this time, the probe was at its perigee. After the maneuver, the probe entered a 3:1 Earth-Moon resonance orbit, meaning that for every one orbit the Moon takes, the probe takes three.
[0111] The third pulse maneuver will be executed at an opportune time, depending on the mission lifespan requirements of the resonant orbit.
[0112] The transfer maneuver yielded the probe's flight trajectory diagram as follows: Figure 2 , Figure 3 and Figure 4 As shown. Figure 2The diagram shows the transfer trajectory of the probe in the geocentric rotating system. The coordinates have been dimensionless, with the unit being the Earth-Moon distance. The Earth's coordinates are (0.01215, 0), and the Moon's coordinates are (0.98785, 0). When the probe reaches the perigee of its lunar orbit, it applies the first pulse. After maneuvering, the probe leaves the lunar orbit and, with the help of lunar gravity, lowers to its perigee. Upon reaching perigee, a second pulse is applied, and after maneuvering, the probe enters a 3:1 Earth-Moon resonance orbit. After the mission concludes, a third pulse is applied at perigee to escape Earth-Moon space. Figure 3 The diagram shows the transfer trajectory of the probe in the Earth-centered ecliptic system. The coordinates have been dimensionless, with the unit being the Earth-Moon distance. The origin of the coordinate system is Earth, and the Moon's orbit is approximately circular. When the probe reaches its perigee in its lunar orbit, it applies its first pulse. After maneuvering, the probe leaves the lunar orbit and, with the help of lunar gravity, lowers its perigee. Upon reaching perigee, a second pulse is applied, and after maneuvering, the probe enters a 3:1 Earth-Moon resonance orbit. Throughout the entire flight, the apogee remains within the Moon's orbit. Figure 4 As shown, this is a curve showing the change in the distance between the probe's orbit and the Earth's center over time. The horizontal axis represents the number of days the probe has been in orbit, and the vertical axis represents the distance between the probe and the Earth's center. The zero point on the horizontal axis is the moment when the first pulse was applied, which shows the descent of the perigee assisted by the Moon's gravity. 41 days later, the second pulse was applied, and the probe entered a 3:1 Earth-Moon resonance orbit. The 6378km dashed line represents the Earth's radius, and the 42000km dashed line represents the altitude of the geosynchronous orbit zone, used to determine whether there is a risk of collision with the geosynchronous orbit zone.
[0113] Based on the design inputs and constraints: the designed transfer trajectory meets the constraints of transfer start time, propellant reserve, and telemetry distance; however, the minimum geocentric distance of the resonant trajectory is close to that of the geosynchronous orbit zone. Analyze whether this trajectory will collide with the geosynchronous orbit zone.
[0114] The analysis specifically examined the closest distances between the resonant orbit and 1031 GEO satellites. The calculation method involved extracting time ranges where the geocentric distance from the resonant orbit was between 41,000 and 43,000 km, and calculating whether the probe collided with the geosynchronous orbit zone within this timeframe. For example... Figure 5 As shown, with Earth as the origin, the diagram displays the trajectories of 1031 GEO satellites and the probe's trajectory, along with their positional relationships, when the probe's resonant orbit distance from the Earth's center is between 41,000 and 43,000 km. Calculations show that the minimum closest distance between the probe and the GEO satellites is 1081.28 km, eliminating the risk of collision.
[0115] Step S150: Analyze the resonant orbit maintenance maneuver.
[0116] The above-mentioned transfer trajectory recursion can meet the design constraints but cannot maintain the 3:1 resonance configuration. Whether to maintain resonance depends on the mission requirements.
[0117] Step S160: Analyze the transfer task status.
[0118] If the transfer start time is delayed by one day, it can still reach the target orbit, but the speed increment will increase by 30 m / s, which will still meet the design input and constraints.
[0119] The detector exhibits good adaptability.
[0120] The aforementioned transfer orbit and pulse key events were all visible at domestic stations; the shadow duration did not exceed 1.7 hours.
[0121] Although the invention has been described with respect to a limited number of embodiments, those skilled in the art will understand from the foregoing description that other embodiments are conceivable within the scope of the invention described herein. Furthermore, it should be noted that the language used in this specification has been chosen primarily for readability and instructional purposes, and not for the purpose of explaining or limiting the subject matter of the invention.
Claims
1. A design method for transferring a lunar orbiter to a lunar resonance orbit, characterized in that, include: Step S110: Determine the remaining propellant of the lunar orbiter; Step S120: Determine the maximum distance for satellite-to-ground telemetry and control; Step S130: Based on the remaining propellant of the lunar orbiter and the farthest distance of the satellite-to-ground telemetry and control system, design the transfer framework and transfer orbit control strategy. The transfer framework refers to the complete path architecture from the lunar orbit, through three pulse maneuvers, finally entering the resonance orbit and finally entering the atmosphere for destruction or escaping the Earth-Moon space. The transfer orbit control strategy refers to the specific provisions of the orbit control parameters and orbit control targets for each pulse in order to realize the above transfer framework. Step S140: Determine the track control parameters and track control objectives; Step S150: Analyze the maneuverability of maintaining the resonant orbit; Step S160: Analyze the transfer task status; Step S140 includes: Step S140-1: Search for the transfer start time: The distance between the geocenter and the perigee of the detector is different under different transfer start times and velocity increments. The distance between the geocenter and the perigee is used to characterize the distance constraint, and a set of transfer start times that satisfy the distance constraint is searched. Step S140-2: Traverse the target perigee radius: Traverse the target perigee radius, take the nominal value of the target angle, obtain the velocity increment required to enter the resonance orbit by firing the target, and obtain the change of the star-to-ground distance by recursively calculating the orbit. Select the optimal target perigee radius from these values. Determine whether the velocity increment is optimal and whether the change of the star-to-ground distance meets the distance constraint and safety constraint. If they meet the constraints, use the velocity increment value used for firing the target in the current traversal step and the target perigee radius to design the orbit transfer. If they do not meet the constraints, continue to traverse and optimize the target angle. Step S140-3: Traverse the target angles: Take the target angles and traverse them. The target perigee radius is the optimal target perigee radius selected in step S140-2. The velocity increment required to enter the resonance orbit is obtained by shooting the target. The orbit is recursively calculated to obtain the change in the star-ground distance. Analyze the traversal results and select a set of target angles that make the velocity increment optimal and the change in the star-ground distance satisfy the distance constraint and safety constraint.
2. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 1, characterized in that, Step S110 includes: Step S110-1: Estimate the mass of remaining propellant using the pressure-volume-temperature method; Step S110-2: Estimating the remaining propellant using the accounting method: By statistically analyzing the propellant consumption during each orbit control mission of the lunar orbiter, the mass of the remaining propellant is estimated. Step S110-3: Calculate the remaining available propellant and velocity increment: Based on the estimation results of the pressure-volume-temperature method and the accounting method, the remaining available propellant amount is obtained by the maximum envelope, and the remaining velocity increment is estimated by combining the relationship between the thruster specific impulse and the thrust magnitude with the remaining fuel mass.
3. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 2, characterized in that, Step S120 includes: Step S120-1: Analyze uplink capabilities; Step S120-2: Analyze bidirectional link capability; Step S120-3: Calculate the maximum telemetry and control distance: Based on the combined uplink capability and bidirectional link capability, determine the maximum distance for satellite-to-ground telemetry and control.
4. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 3, characterized in that, Step S130 includes: Step S130-1: Define the design inputs and constraints for the transfer trajectory; Step S130-2: Design the three-pulse transfer framework; Step S130-3: Design a three-pulse transfer track control strategy.
5. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 4, characterized in that, The design inputs for the transfer orbit include the remaining propellant on the lunar orbiter and the maximum distance for satellite-to-ground telemetry and control. The constraints of the transfer orbit design include: transfer start time constraint, the transfer is designed after the lunar probe completes its predetermined mission; total velocity increment constraint, the transfer orbit with the optimal velocity increment is designed; flight distance constraint, not exceeding the farthest telemetry and control distance; safety constraint, flying outside the geosynchronous orbit to avoid the risk of collision; and complete orbit design, the probe will eventually enter the atmosphere and be destroyed or escape from the Earth-Moon space.
6. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 4, characterized in that, The three-pulse transfer framework is as follows: When the lunar probe reaches the perigee of the lunar orbit, the first pulse is applied, with the pulse direction along the positive tangential direction of the orbit. After the maneuver, the probe leaves the lunar orbit and lowers its perigee with the help of lunar gravity. When the probe reaches the target perigee, the second pulse is applied, with the pulse direction along the tangential direction of the orbit. After the maneuver, the probe enters the Earth-Moon resonance orbit. After the mission is completed, a third pulse is applied at the apogee, with the pulse direction reversed along the orbital tangent to lower the orbital perigee and eventually burn up in the atmosphere; or a third pulse is applied at the perigee, with the pulse direction forward along the orbital tangent to raise the orbital apogee and eventually escape the Earth-Moon space.
7. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 4, characterized in that, The three-pulse orbit control strategy includes: The orbit control parameters for the first pulse are: First variable: transfer start time, i.e., the moment the perilune pulse is applied; Second variable: the velocity increment along the positive tangential direction of the orbit at the perilune; The orbit control objectives are: First objective: the perigee radius upon reaching perigee; Second objective: the angle between the probe, Earth, and the Moon upon reaching perigee. The orbit control parameters for the second pulse are: the third variable: the velocity increment along the orbital tangential at perigee; the orbit control objective is: the third objective: the perigee angle between the probe, Earth, and Moon after multiple orbits. The orbital control parameters for the third pulse's atmospheric destruction are: Fourth variable: Velocity increment in the opposite direction along the orbital tangential at apogee; Orbital control objective: Fourth objective: Reduce perigee radius; The orbital control parameters for the third pulse to escape the Earth-Moon space are: the fifth variable: the magnitude of the velocity increment along the positive tangential direction of the orbit at perigee; the orbital control objective is: the fifth objective: to raise the apogee radius.
8. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 7, characterized in that, Step S150 includes: Step S150-1: After the probe enters the Earth-Moon space resonance orbit, determine whether the resonance can be maintained during the subsequent flight process under the influence of gravity and perturbation, and whether the design constraints are still met. Step S150-2: If the subsequent flight process can maintain resonance and meet the transfer trajectory design constraints, then maintenance maneuvers are not required; Step S150-3: If the subsequent flight process cannot maintain resonance and meet the transfer trajectory design constraints, a maintenance maneuver is required; Step S150-4: If the subsequent flight process cannot maintain resonance, but can meet the design constraints of the transfer trajectory, then decide whether to perform a maintenance maneuver based on mission requirements. Step S150-5: If a maneuver is to be sustained, under the pulse model, the orbital control parameters for sustaining the maneuver are: The third variable: the velocity increment along the orbital tangent at perigee; the orbital control objective is: the third objective: the angle between the perigee probe, Earth, and Moon after multiple orbits.
9. The design method for transferring a lunar probe to a lunar resonance orbit according to claim 8, characterized in that, Step S160 includes: Step S160-1: Analyze the delayed transfer start time: If the first pulse is delayed, analyze the orbital transfer situation when the first pulse is applied at the next lunar perigee; Step S160-2: Analyze the detector's adaptability; Step S160-3: Analyze the probe's flight status.