Method for designing rapid rendezvous and docking orbit of multi-task spacecraft

Abstract

The invention discloses a rapid rendezvous and docking orbit design method for a multi-task spacecraft, and the method comprises the steps: obtaining a phase adaptation lower limit and a phase adaptation upper limit through employing a rapid rendezvous and docking orbit transfer strategy according to a preset target aircraft orbit height and a preset active aircraft injection initial orbit; obtaining the earliest moment when the spacecraft performs orbital maneuver; analyzing a phase deviation maximum value and a phase deviation minimum value caused by the initial orbit deviation of the active aircraft at the earliest moment of orbit maneuver of the spacecraft according to the initial orbit deviation of the active aircraft in orbit injection; according to the fact that the minimum value of phase deviation caused by initial orbit deviation of the active aircraft at the earliest moment of orbit maneuver of the spacecraft is not smaller than the rapid rendezvous and docking phase adaptation lower limit, the orbit phase value at the orbit injection moment is obtained. According to the method, the orbital transfer interval is prolonged, the task requirement of an on-orbit test is met, rapid rendezvous and docking are also considered, and the adaptability of the spacecraft to the fault working condition is enhanced.

Description

Technical Field

[0001] The invention belongs to the technical field of spacecraft orbit design, and in particular relates to a rapid rendezvous and docking orbit design method for a multi-mission spacecraft. Background Art

[0002] The rendezvous and docking technology was first explored by the United States and the Soviet Union in the 1960s and gradually developed into a mature and reliable 2-5 day rendezvous and docking mode. During this period, multiple orbital controls were performed to reduce the distance between the two spacecraft and correct the speed deviation. The rendezvous and docking process is mainly divided into a long-distance guidance phase and a short-distance autonomous control phase. Among them, the long-distance guidance phase is controlled by the ground. According to the results of precise orbit determination, five orbital controls are planned to correct the phase difference and orbital plane deviation between the two targets and aim at the coplanar circular orbit about 50 km behind and below the target spacecraft. The rendezvous and docking strategy based on ground long-distance guidance is highly dependent on ground control. In the interaction between the earth and the sky, at least three orbital cycles are required between each two orbit changes for precise orbit determination, guidance strategy calculation, flight program generation and injection, and the total rendezvous and docking time is relatively long. After entering the 21st century, in order to reduce the waiting time of astronauts in the spacecraft and adapt to the increasingly active space activities, Russia took the lead in conducting rapid rendezvous and docking experiments. After orbiting the earth four times, it successfully docked with the "International Space Station". Compared with the original technology, the main difference of autonomous rapid rendezvous and docking is that the long-distance guidance stage does not rely on ground control. The active aircraft calculates the flight trajectory based on the satellite navigation equipment, and then autonomously calculates the relative relationship between the two targets based on the orbital elements of the target aircraft, and plans the orbit change strategy for rendezvous and docking. Since it does not rely on ground orbit determination data and strategy calculations, the long-distance stage adopts quasi-Hohmann orbit change, and the shortest orbit control can be performed once every half a circle, so the shortest time to complete the docking is about 3 hours.

[0003] However, with the increasing complexity of spacecraft missions, after separation from the carrier, the spacecraft needs to perform a series of platform status settings and tests before performing the rendezvous and docking mission. It is impossible to achieve orbital maneuvers and does not meet the requirements of rapid orbit change. In addition, as the flight time increases, the orbital deviation caused by the initial orbital deviation is gradually amplified, which poses a higher challenge to the orbital adaptability of rapid rendezvous and docking. If the original 2-5 day rendezvous and docking mode is adopted, the problems of long spacecraft platform status setting time and large orbital deviation can be basically solved, but the total rendezvous and docking time is too long, which does not match the demand for rapid docking in the event of failures such as the failure of the solar wing to unfold. Summary of the invention

[0004] The technical problem solved by the present invention is: to overcome the deficiencies of the prior art and provide a method for designing a rapid rendezvous and docking orbit for a multi-mission spacecraft, which not only prolongs the orbit change interval to meet the mission requirements of on-orbit testing, but also takes into account rapid rendezvous and docking, thereby enhancing the spacecraft's ability to adapt to fault conditions.

[0005] The purpose of the present invention is achieved through the following technical solutions: a method for designing a rapid rendezvous and docking orbit for a multi-task spacecraft, comprising: according to a preset target spacecraft orbit altitude and a preset active spacecraft initial orbit for orbit entry, a rapid rendezvous and docking orbit change strategy is adopted to obtain a phase adaptation lower limit and a phase adaptation upper limit; the earliest time for the spacecraft to perform an orbital maneuver is obtained; according to the initial orbital deviation of the active spacecraft for orbit entry, a maximum value and a minimum value of the phase deviation caused by the initial orbital deviation of the active spacecraft at the earliest time for the spacecraft to perform an orbital maneuver are analyzed; according to the minimum value of the phase deviation caused by the initial orbital deviation of the active spacecraft at the earliest time for the spacecraft to perform an orbital maneuver is not less than the rapid rendezvous and docking phase adaptation lower limit, an orbit phase value at the time of orbit entry is obtained.

[0006] In the rapid rendezvous and docking orbit design method of the above-mentioned multi-mission spacecraft, the rapid rendezvous and docking orbit change strategy includes: the long-distance orbit change mainly considers the principle of optimal propellant, and adopts the long-distance guidance method of quasi-Hohmann transfer to guide the active spacecraft from the initial orbit for entering orbit to the rear and lower part of the target spacecraft; the orbit change is carried out independently in the orbital plane and outside the orbital plane, wherein the orbital maneuvering task in the orbital plane is to make the active spacecraft pass through one or more intermediate transition phase adjustment orbits from the orbital entry orbit, increase the orbital altitude and adjust the orbital phase, and the pulse control in the orbital plane is carried out at the apogee and perigee of the orbit; the orbital maneuvering task outside the orbital plane is to correct the initial orbital plane deviation, wherein the initial orbital plane deviation includes the orbital inclination deviation and the ascending node right ascension deviation, and two orbital controls are used to respectively perform two orbital inclination and ascending node right ascension out of the orbital plane control; the third pulse also plays a phase adjustment role, which reduces the number of orbital controls; considering the orbit deviation and engine execution deviation determined by each maneuver, two combined correction maneuvers are designed to eliminate the above orbital maneuver deviations.

[0007] In the above-mentioned rapid rendezvous and docking orbit design method for multi-mission spacecraft, the earliest time for the spacecraft to perform orbital maneuvers is: from the time of orbit entry T0 to the time of completing platform setting and on-orbit testing is T i , from T i The time of the nearest orbital apogee is T a The earliest time for the spacecraft to perform orbital maneuvers is from T i The time of the nearest orbital apogee.

[0008] In the above-mentioned rapid rendezvous and docking orbit design method for multi-mission spacecraft, the initial orbit deviation of the active spacecraft entering orbit includes orbital period deviation and latitude argument deviation.

[0009] In the rapid rendezvous and docking orbit design method for the above-mentioned multi-mission spacecraft, the phase deviation caused by the initial orbit deviation of the active spacecraft is at a minimum when both the orbital period deviation and the dimensional angle deviation take their maximum values, which is the earliest moment for the spacecraft to perform orbital maneuvers.

[0010] In the rapid rendezvous and docking orbit design method for the above-mentioned multi-mission spacecraft, the maximum value of the phase deviation caused by the initial orbit deviation of the active spacecraft is when both the orbital period deviation and the dimensional angle deviation take their minimum values, which is the earliest moment for the spacecraft to perform orbital maneuvers.

[0011] In the above-mentioned rapid rendezvous and docking orbit design method for multi-mission spacecraft, the orbit phase value at the time of orbit entry is obtained by the following formula:

[0012]

[0013] in, is the orbit phase at the time of orbit entry, The lower limit of phase adaptation for rapid rendezvous and docking, T a is the earliest time for the spacecraft to perform orbital maneuvers, ω t is the target spacecraft orbital angular velocity, ω c is the orbital angular velocity of the active spacecraft, and Δμ is the latitude argument deviation.

[0014] A rapid rendezvous and docking orbit design system for a multi-mission spacecraft comprises: a first module, used for obtaining a phase adaptation lower limit and a phase adaptation upper limit by adopting a rapid rendezvous and docking orbit change strategy according to a preset target spacecraft orbit altitude and a preset active spacecraft initial orbit for orbit entry; a second module, used for obtaining the earliest time for the spacecraft to perform orbital maneuvers; a third module, used for analyzing a maximum value and a minimum value of a phase deviation caused by an initial orbital deviation of the active spacecraft at the earliest time for the spacecraft to perform orbital maneuvers according to an initial orbital deviation of the active spacecraft for orbit entry; and a fourth module, used for obtaining an orbital phase value at the time of orbit entry according to the minimum value of a phase deviation caused by an initial orbital deviation of the active spacecraft at the earliest time for the spacecraft to perform orbital maneuvers being not less than a rapid rendezvous and docking phase adaptation lower limit.

[0015] In the rapid rendezvous and docking orbit design system of the above-mentioned multi-mission spacecraft, the rapid rendezvous and docking orbit change strategy includes: the long-distance orbit change mainly considers the principle of optimal propellant, and adopts the long-distance guidance method of quasi-Hohmann transfer to guide the active spacecraft from the initial orbit for entering orbit to the rear and lower part of the target spacecraft; the orbit change is carried out independently in the orbital plane and outside the orbital plane, among which, the orbital maneuvering task in the orbital plane is to make the active spacecraft pass through one or more intermediate transition phase adjustment orbits from the orbital entry orbit, increase the orbital altitude and adjust the orbital phase, and the pulse control in the orbital plane is carried out at the apogee and perigee of the orbit; the orbital maneuvering task outside the orbital plane is to correct the initial orbital plane deviation, among which, the initial orbital plane deviation includes the orbital inclination deviation and the ascending node right ascension deviation, and two orbital controls are used to respectively perform two orbital inclination and ascending node right ascension out of the orbital plane control; the third pulse also plays a phase adjustment role, which reduces the number of orbital controls; considering the orbit deviation and engine execution deviation of each maneuver, two combined correction maneuvers are designed to eliminate the above orbital maneuver deviations.

[0016] In the rapid rendezvous and docking orbit design system of the multi-mission spacecraft, the earliest time for the spacecraft to perform orbital maneuvers is: from the time of orbit entry T0 to the time of completing platform setting and on-orbit testing is T i , from T i The time of the nearest orbital apogee is T a The earliest time for the spacecraft to perform orbital maneuvers is from T i The time of the nearest orbital apogee.

[0017] Compared with the prior art, the present invention has the following beneficial effects:

[0018] (1) The present invention combines the constraints of the flight mission duration and the distribution characteristics of the initial orbit and the deviation orbit for the launch, and jointly plans the passive spacecraft phase adjustment orbit and the active spacecraft rendezvous and docking strategy. In the design of the rapid rendezvous and docking orbit, the active spacecraft catch-up phase and the correction pulse are reserved. This not only prolongs the orbit change interval to meet the mission requirements of the on-orbit test, but also takes into account the rapid rendezvous and docking, and enhances the adaptability of the spacecraft to fault conditions, and has good engineering application value.

[0019] (2) The present invention combines the on-orbit flight plan with the characteristics of the rapid rendezvous and docking orbit to achieve the purpose of delaying orbit change to meet the mission requirements of platform setup and testing in the initial stage of orbit entry, shorten the rendezvous and docking time, and improve the fault handling capability, thus solving the rendezvous and docking problem of spacecraft with complex flight missions;

[0020] (3) The present invention analyzes the distribution characteristics of orbital deviation after launch and develops an adaptive orbit change strategy and implementation method, thereby improving the adaptability of rendezvous and docking to orbital deviation and solving the problem of active spacecraft performing rapid rendezvous and docking missions under unfavorable conditions such as a wide launch window for cryogenic launch vehicles and large orbital deviation. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present invention. Moreover, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0022] Figure 1 is a schematic diagram of a rapid rendezvous and docking pulse provided by an embodiment of the present invention;

[0023] Figure 2 It is a schematic diagram of the rapid rendezvous and docking track surface deviation correction provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0024] Exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided in order to enable a more thorough understanding of the present disclosure and to be able to fully convey the scope of the present disclosure to those skilled in the art. It should be noted that, in the absence of conflict, the embodiments of the present invention and the features described in the embodiments can be combined with each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0025] In order to realize the rendezvous and docking of active spacecraft and target spacecraft, the commonly used methods are long-term large phase difference rendezvous and docking based on ground guidance and fully autonomous small phase rapid rendezvous and docking. The advantages of large rendezvous and docking based on ground guidance are strong phase adaptability and long orbit change interval, which can be planned synchronously with other tasks on orbit after entering orbit, but the rendezvous and docking time is too long and the fault adaptability is weak. Although the rendezvous and docking time of fully autonomous small phase rapid rendezvous and docking is short, it has high requirements for orbit and poor phase adaptability, and cannot meet the implementation of other tasks synchronously. With the development of the aerospace field, spacecraft must not only meet complex flight missions, but also achieve rendezvous and docking as quickly as possible. The commonly used orbit design methods cannot meet the growing mission needs.

[0026] This embodiment combines the constraints of flight mission duration and the distribution characteristics of the initial orbit and deviation orbit for launch into orbit, and through joint planning of the passive aircraft phase adjustment orbit and the active spacecraft rendezvous and docking strategy, reserves the active spacecraft catch-up phase and increases the correction pulse in the rapid rendezvous and docking orbit design. This not only extends the orbit change interval to meet the mission requirements of on-orbit testing, but also takes into account rapid rendezvous and docking, enhances the spacecraft's adaptability to fault conditions, and has good engineering application value.

[0027] This embodiment provides a method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft, the method comprising:

[0028] According to the preset target spacecraft orbital altitude and the preset active spacecraft initial orbit, the phase adaptation lower limit is obtained by adopting the rapid rendezvous and docking orbit change strategy. and phase adaptation upper limit

[0029] Get the earliest time when the spacecraft can perform orbital maneuvers;

[0030] According to the initial orbital deviation of the active spacecraft entering orbit, the maximum phase deviation caused by the initial orbital deviation of the active spacecraft is analyzed at the earliest time when the spacecraft performs orbital maneuvers. and minimum

[0031] The minimum phase deviation caused by the initial orbit deviation of the active spacecraft at the earliest time when the spacecraft performs orbital maneuvers Not less than the lower limit of rapid rendezvous and docking phase adaptation Get the orbit phase value at the time of orbit entry.

[0032] According to the orbital altitude of the target spacecraft and the initial orbit of the active spacecraft, without considering the duration constraints of the active spacecraft on-orbit platform setting and on-orbit testing, that is, when the spacecraft meets the orbit change conditions in real time, the phase adaptation lower limit is known according to the rapid rendezvous and docking orbit change strategy. and upper limit

[0033] According to the on-orbit flight plan of the spacecraft, the time from the orbit entry time T0 to the completion of the platform setup and on-orbit test is T i , from T i The time of the nearest orbital apogee is T a .

[0034] According to the initial orbit deviation of the active spacecraft into orbit, T a The maximum phase deviation caused by the initial orbit deviation of the active spacecraft within the time range and minimum The influencing factors include orbital period deviation and latitude argument deviation. According to the orbital characteristics, the phase deviation takes a minimum value when the orbital period and latitude argument take a maximum value, and correspondingly, the phase deviation takes a maximum value when the orbital period and latitude argument take a minimum value.

[0035] The target spacecraft obtains the orbit phase adjustment target, T a The minimum phase deviation of the active aircraft at any time is not less than the lower limit of the phase adaptation for rapid rendezvous and docking Right now

[0036] The rapid rendezvous and docking orbit change strategy includes: the long-distance orbit change mainly considers the principle of optimal propellant, and adopts the long-distance guidance method of quasi-Hohmann transfer to guide the active spacecraft from the initial orbit for orbit entry to the rear and lower part of the target spacecraft; the orbit change is carried out independently in the orbital plane and outside the orbital plane, among which, the orbital maneuvering task in the orbital plane is to make the active spacecraft pass through one or more intermediate transition phase adjustment orbits from the orbital entry orbit, increase the orbital altitude and adjust the orbital phase, and the pulse control in the orbital plane is carried out at the apogee and perigee of the orbit; the orbital maneuvering task outside the orbital plane is to correct the initial orbital plane deviation, among which, the initial orbital plane deviation includes the orbital inclination deviation and the ascending node right ascension deviation, and two orbital controls are used to perform two orbital inclination and ascending node right ascension out of the orbital plane respectively; the third pulse also plays a phase adjustment role, which reduces the number of orbital controls; considering the orbit deviation and engine execution deviation determined by each maneuver, two combined correction maneuvers are designed to eliminate the above orbital maneuver deviations.

[0037] The earliest time for the spacecraft to perform orbital maneuvers is: from the time of orbit entry T0 to the time of completing platform setup and on-orbit testing is T i , from T i The time of the nearest orbital apogee is T a The earliest time for the spacecraft to perform orbital maneuvers is from T i The time of the nearest orbital apogee.

[0038] The initial orbital deviation of the active spacecraft entering orbit includes the orbital period deviation and the latitude argument deviation. When both the orbital period deviation and the latitude argument deviation are at their maximum, it is the earliest time for the spacecraft to perform orbital maneuvers, and the phase deviation caused by the initial orbital deviation of the active spacecraft is at its minimum. When both the orbital period deviation and the latitude argument deviation are at their minimum, it is the earliest time for the spacecraft to perform orbital maneuvers, and the phase deviation caused by the initial orbital deviation of the active spacecraft is at its maximum.

[0039] The orbit phase value at the time of orbit entry is obtained by the following formula:

[0040]

[0041] in, is the orbit phase at the time of orbit entry, The lower limit of phase adaptation for rapid rendezvous and docking, T a is the earliest time for the spacecraft to perform orbital maneuvers, ω t is the target spacecraft orbital angular velocity, ω c is the orbital angular velocity of the active spacecraft, and Δμ is the latitude argument deviation.

[0042] Rapid long-distance track change strategy and capability analysis:

[0043] Long-distance trajectory change mainly considers the principle of optimal propellant and adopts the long-distance guidance method of quasi-Hohmann transfer to guide the active spacecraft from the initial orbit of orbit entry to the vicinity of the target spacecraft, usually behind and below the target spacecraft.

[0044] Orbit change is carried out independently in the orbital plane and outside the orbital plane. The orbital maneuvering task in the orbital plane is to make the active spacecraft pass through one or more intermediate transition phase adjustment orbits from the orbital entry orbit, increase the orbital altitude and adjust the orbital phase. The pulse control in the orbital plane is mainly carried out at the orbital apogee and perigee. Figure 1 As shown in the figure. The main correction of the initial orbital plane deviation outside the orbital plane includes the orbital inclination deviation and the ascending node right ascension deviation. Two orbital controls are used to perform the orbital inclination and ascending node right ascension respectively. Figure 2 As shown, the problem of simultaneously correcting the inclination and the right ascension deviation of the ascending node in one orbit control resulting in an uncertain orbit control implementation point that needs to be determined in real time according to the out-of-plane deviation of the entry orbit is avoided. The third pulse also plays a role in phase modulation, reducing the number of orbit control times.

[0045] After performing in-orbit maneuvers and orbital correction maneuvers respectively, two combined correction maneuvers are designed taking into account the orbital deviation and engine execution deviation of each maneuver to eliminate the above orbital maneuver deviations and meet the terminal orbital condition requirements.

[0046] Table 1: Rapid rendezvous and docking long-distance orbit control strategy

[0047]

[0048] During the design, except for the orbital plane correction pulse, all other phase adjustment and comprehensive correction pulses are positive pulses, that is, the velocity increment is positive. Without considering the spacecraft mission arrangement, only considering the rapid rendezvous and docking orbit change requirements, the optimal phase difference requirement with the target spacecraft under the condition of zero deviation in the launch orbit was obtained. And the lower limit of phase adaptation considering the launch altitude and orbit period deviation and upper limit

[0049] Rapid orbit change opportunity assessment:

[0050] After the spacecraft enters orbit, the platform is set up and key equipment is tested first. Its ability to change orbit quickly is limited, and it is impossible to implement orbit phase control at the first apogee after entering orbit. Considering the initial attitude deviation elimination, propulsion pipeline filling, relay antenna deployment, solar wing deployment, and unlocking of large extravehicular equipment, the basic platform initial state establishment and setting and rendezvous and docking function testing, including translation engine testing, rendezvous and docking sensor testing, docking mechanism, rendezvous and docking equipment inspection, etc., the total time is T i According to the analysis in the previous section, the long-distance orbit control phase adjustment phase of the rendezvous and docking is carried out at the orbital apogee and perigee, and the first orbit control is carried out at the apogee. Therefore, it is concluded that the earliest time for the spacecraft to perform orbital maneuvers is the time of the nearest orbital apogee T a .

[0051] Initial orbit phase requirements after orbit deviation analysis and strategy adjustment:

[0052] The rocket orbital deviation items mainly include orbital time, orbital period, perigee height, orbital inclination, ascending node longitude and perigee argument, a total of 6 items. Among them, orbital time, orbital period, perigee height and perigee argument are orbital plane deviations, while orbital inclination and ascending node longitude are orbital plane deviations.

[0053] The in-plane deviation mainly affects the phase difference between the two spacecraft. The orbital entry time and perigee argument deviation can be combined into the latitude argument deviation. The orbital period deviation affects the semi-major axis of the orbital entry, thereby affecting the orbital angular velocity after orbital entry. When the orbital period deviation is negative, the orbital semi-major axis becomes smaller, the tracker orbit is low, and the angular velocity is faster, resulting in T a When the orbital period deviation is positive, the semi-major axis of the orbit becomes larger, the tracker orbit is high, and the angular velocity is slow, resulting in T a The phase difference between the two targets increases at time T. The perigee height deviation also affects the orbital semi-major axis and thus changes the orbital angular velocity. a At time , the orbital phase difference between the target spacecraft and the active spacecraft is

[0054]

[0055] where ω t is the target spacecraft orbital angular velocity, is the phase difference between the two spacecraft at the theoretical orbital entry time, ω c is the orbital angular velocity of the active spacecraft, and Δμ is the latitude argument deviation.

[0056] Table 2: Phase deviation value determination method

[0057] <![CDATA[ΔT0(s)]]> Δμ(°) <![CDATA[ΔA p (km)]]> + + + + + - + - + + - - - + + - + - - - + - - -

[0058] The phase difference extreme value is obtained according to the method in Table 2 above. Where "+" indicates that the deviation is at the maximum value, and "-" indicates that the deviation is at the minimum value. Substitute into formula (1) to calculate T a Take the extreme values ​​of phase deviation under different working conditions. Taking the orbital parameters of a certain type of launch vehicle as an example, the working condition of the first row of values ​​in Table 2 indicates an orbital period deviation of 10s, a latitude angle deviation of 2°, and a perigee height deviation of 2km. The working condition of the last row of values ​​in Table 2 indicates an orbital period deviation of -10s, a latitude angle deviation of -2°, and a perigee height deviation of -2km. Finally, it is required that the phase difference under each working condition in Table 2 meets T a The time is not less than the lower limit Right now

[0059]

[0060] Substituting into formula (1), we can get the orbit phase value at the time of orbit entry:

[0061]

[0062] The orbital inclination and ascending node longitude deviation caused by the launch are constant values, and are independently corrected by the third and fourth orbital controls after the launch. In addition, these two deviations are usually corrected as small as possible on-orbit by the rocket's orbital inclination bias and launch time bias before the launch.

[0063] Semi-physical target shooting simulation verification: Through semi-physical simulation, considering factors such as orbit disturbance and orbit control deviation caused by engine test, target shooting simulation is carried out for various orbit deviation parameters according to the sampling method of equal probability distribution, focusing on verifying 1: the adaptability of the initial orbit to the flight plan, that is, the first start-up time of the orbit control is not earlier than the time when the active spacecraft platform is set up and the on-orbit test is completed. 2: After the long-distance orbit change is completed, the orbit meets the entry conditions of the close-range guidance control.

[0064] The present embodiment also provides a rapid rendezvous and docking orbit design system for a multi-mission spacecraft, the system comprising: a first module, used to obtain a phase adaptation lower limit and a phase adaptation upper limit by adopting a rapid rendezvous and docking orbit change strategy according to a preset target spacecraft orbit altitude and a preset active spacecraft initial orbit for orbit entry; a second module, used to obtain the earliest time for the spacecraft to perform an orbital maneuver; a third module, used to analyze a maximum and minimum value of the phase deviation caused by the initial orbital deviation of the active spacecraft at the earliest time for the spacecraft to perform an orbital maneuver according to the initial orbital deviation of the active spacecraft for orbit entry; a fourth module, used to obtain an orbital phase value at the time of orbit entry according to the minimum value of the phase deviation caused by the initial orbital deviation of the active spacecraft at the earliest time for the spacecraft to perform an orbital maneuver being not less than the rapid rendezvous and docking phase adaptation lower limit.

[0065] This embodiment, by combining the on-orbit flight plan and the rapid rendezvous and docking orbit characteristics, can achieve the purpose of delaying orbit change to meet the mission requirements of platform setup and testing in the initial stage of orbit entry and shortening the rendezvous and docking time to improve the fault handling capability, thereby solving the rendezvous and docking problem of spacecraft in complex flight missions; This embodiment analyzes the distribution characteristics of orbital deviation after launch, and formulates an adaptive orbit change strategy and implementation method, thereby improving the adaptability of rendezvous and docking to orbital deviation, and solving the problem of active aircraft performing rapid rendezvous and docking missions under unfavorable conditions such as a wide launch window for cryogenic carrier rockets and large orbital deviations.

[0066] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft, characterized in that include: According to the preset target spacecraft orbit altitude and the preset active spacecraft initial orbit for orbit entry, the phase adaptation lower limit and phase adaptation upper limit are obtained by adopting the rapid rendezvous and docking orbit change strategy; Get the earliest time when the spacecraft can perform orbital maneuvers; According to the initial orbital deviation of the active spacecraft entering orbit, the maximum and minimum phase deviations caused by the initial orbital deviation of the active spacecraft at the earliest moment of orbital maneuvering of the spacecraft are analyzed; The orbital phase value at the time of entering orbit is obtained based on the minimum value of the phase deviation caused by the initial orbital deviation of the active vehicle when the spacecraft performs orbital maneuvers, which is not less than the lower limit of the phase adaptation for rapid rendezvous and docking.

2. The method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft according to claim 1, characterized in that: Rapid rendezvous and docking trajectory change strategies include: Long-distance trajectory change mainly considers the principle of optimal propellant, and adopts the long-distance guidance method of quasi-Hohmann transfer to guide the active spacecraft from the initial orbit of orbit entry to the rear and lower part of the target spacecraft; The orbit change is carried out independently within the orbital plane and outside the orbital plane. The orbital maneuvering task within the orbital plane is to make the active spacecraft pass through one or more intermediate transition phase adjustment orbits from the entry orbit, increase the orbital altitude and adjust the orbital phase. The pulse control within the orbital plane is carried out at the orbital apogee and perigee; the orbital maneuvering task outside the orbital plane is to correct the initial orbital plane deviation, wherein the initial orbital plane deviation includes the orbital inclination deviation and the ascending node right ascension deviation. Two orbital controls are used to perform two orbital inclination and ascending node right ascension out of the orbital plane respectively; the third pulse also plays a role in phase adjustment, reducing the number of orbital controls; considering the orbit deviation determined by each maneuver and the engine execution deviation, two combined correction maneuvers are designed to eliminate the above orbital maneuver deviations.

3. The method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft according to claim 1, characterized in that: The earliest time for the spacecraft to perform orbital maneuvers is: The time from the time of entering orbit T0 to the time of completing platform setup and on-orbit testing is T i , from T i The time of the nearest orbital apogee is T a The earliest time for the spacecraft to perform orbital maneuvers is from T i The time of the nearest orbital apogee.

4. The method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft according to claim 1, characterized in that: The initial orbital deviation of the active spacecraft entering orbit includes an orbital period deviation and a latitude argument deviation.

5. The method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft according to claim 4, characterized in that: When both the orbital period deviation and the dimensional argument deviation take their maximum values, it is the earliest moment for the spacecraft to perform orbital maneuvers, and the phase deviation caused by the initial orbital deviation of the active spacecraft is at its minimum value.

6. The method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft according to claim 4, characterized in that: When both the orbital period deviation and the dimensional argument deviation take their minimum values, it is the earliest moment when the spacecraft performs orbital maneuvers, and the phase deviation caused by the initial orbital deviation of the active spacecraft is maximum.

7. The method for designing a rapid rendezvous and docking trajectory for a multi-mission spacecraft according to claim 1, characterized in that: The orbit phase value at the time of orbit entry is obtained by the following formula: in, is the orbit phase at the time of orbit entry, The lower limit of phase adaptation for rapid rendezvous and docking, T a is the earliest time for the spacecraft to perform orbital maneuvers, ω t is the target spacecraft orbital angular velocity, ω c is the orbital angular velocity of the active spacecraft, and Δμ is the latitude argument deviation.

8. A rapid rendezvous and docking trajectory design system for multi-mission spacecraft, characterized in that include: The first module is used to obtain the phase adaptation lower limit and the phase adaptation upper limit by adopting the rapid rendezvous and docking trajectory change strategy according to the preset target spacecraft orbit altitude and the preset active spacecraft initial orbit for orbit entry; The second module is used to obtain the earliest time when the spacecraft can perform orbital maneuvers; The third module is used to analyze the maximum and minimum phase deviations caused by the initial orbital deviation of the active vehicle at the earliest moment of the spacecraft's orbital maneuver according to the initial orbital deviation of the active vehicle entering orbit; The fourth module is used to obtain the orbital phase value at the time of entering orbit based on the fact that the minimum value of the phase deviation caused by the initial orbital deviation of the active aircraft at the earliest time when the spacecraft performs orbital maneuvers is not less than the lower limit of the phase adaptation for rapid rendezvous and docking.

9. The rapid rendezvous and docking trajectory design system for multi-mission spacecraft according to claim 8, characterized in that: Rapid rendezvous and docking trajectory change strategies include: Long-distance trajectory change mainly considers the principle of optimal propellant, and adopts the long-distance guidance method of quasi-Hohmann transfer to guide the active spacecraft from the initial orbit of orbit entry to the rear and lower part of the target spacecraft; The orbit change is carried out independently within the orbital plane and outside the orbital plane. The orbital maneuvering task within the orbital plane is to make the active spacecraft pass through one or more intermediate transition phase adjustment orbits from the entry orbit, increase the orbital altitude and adjust the orbital phase. The pulse control within the orbital plane is carried out at the orbital apogee and perigee; the orbital maneuvering task outside the orbital plane is to correct the initial orbital plane deviation, wherein the initial orbital plane deviation includes the orbital inclination deviation and the ascending node right ascension deviation. Two orbital controls are used to perform two orbital inclination and ascending node right ascension out of the orbital plane respectively; the third pulse also plays a role in phase adjustment, reducing the number of orbital controls; considering the orbit deviation determined by each maneuver and the engine execution deviation, two combined correction maneuvers are designed to eliminate the above orbital maneuver deviations.

10. The rapid rendezvous and docking trajectory design system for multi-mission spacecraft according to claim 8, characterized in that: The earliest time for the spacecraft to perform orbital maneuvers is: The time from the time of entering orbit T0 to the time of completing platform setup and on-orbit testing is T i , from T i The time of the nearest orbital apogee is T a The earliest time for the spacecraft to perform orbital maneuvers is from T i The time of the nearest orbital apogee.

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