CN113268809A - Design method and system for solar system marginal detection electric propulsion transfer orbit - Google Patents
Design method and system for solar system marginal detection electric propulsion transfer orbit Download PDFInfo
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Abstract
The invention provides a solar system marginal detection electric propulsion transfer orbit design method and a system, which relate to the technical field of spacecraft orbit design and optimization, and the method comprises the following steps: step S1: establishing a deep space electric propulsion-gravity transfer orbit design model and setting a flight sequence; step S2: setting a detector to start from the earth at a resonance ratio close to 1:2 according to a set flight sequence to obtain relevant parameters including emission parameters, planet gravity parameters, transfer time and fuel consumption; step S3: setting relevant constraint conditions including emission C3 and gravity-assist height, and performing preliminary optimization by using an SQP algorithm according to a Sims-Flanagan model and a gravity-assist orbit model; step S4: and according to the preliminary optimization result, carrying out re-optimization according to an indirect method to obtain a high-precision solution. The invention realizes the balance among the track precision, the calculated amount and the design efficiency, quickly and effectively completes the initial solution of the transfer track to obtain the optimized design, effectively reduces the iteration times, reduces the calculated amount and ensures the robustness of numerical iteration.
Description
Technical Field
The invention relates to the technical field of spacecraft orbit design and optimization, in particular to a solar system marginal detection electric propulsion transfer orbit design method and system.
Background
Since the last century, human deep space exploration gradually expands to remote interplanetary space, and exploration targets cover main celestial bodies ranging from the near ground to the solar system marginal range. In general, human detection of solar system margin and beyond is still limited, and currently, only a few detectors such as traveler number and new field number arrive in the space area. The detection of solar system boundary, even the detection of outer space is the inevitable choice for future aerospace technology development. Such detection tasks tend to be complicated continuously, and the urgent need of rapid transfer is provided for the design of the track, and the problems of large emission energy, high detector fuel cost and the like are brought. Sqp (sequence quadratic program) is an abbreviation for operational research definition. Sequential Quadratic Programming (SQP) algorithms are a class of effective algorithms for solving medium and small programming constraint optimization problems.
Compared with other deep space detection tasks, the solar system marginal detection task belongs to ultra-long-distance transfer, and not only faces the technical problems of large emission energy, high detector fuel cost and the like, but also needs to solve the task time requirement of rapid transfer urgently to ensure that the detector flies to a target area of 85AU or even more within reasonable time.
At present, some researches are carried out on the design of solar system marginal detection tasks, and through retrieval, the following are mainly related to the design of a transfer orbit:
the document 'solar system marginal detection research' (Chinese science: information science, 2019, No. 01: 1-16) explains 4 types of scientific targets of solar system marginal detection, provides the overall target, stage target and recent detection task assumption of solar system marginal detection in China, provides 6 types of key technologies which should be researched and broken through in a focused manner, and provides reference for the demonstration and implementation of solar system marginal detection in China. This document relates primarily to transfer track design and optimization techniques.
The invention patent with publication number CN106986049A discloses a deep space gravity orbit accurate parallel optimization design method, which aims to improve the speed of gravity sequence search and parameter optimization, reduce the time for obtaining the preliminary solution, and ensure the accuracy of the final solution. The general design method of the chemical propulsion-gravity transfer orbit concerned by the patent has the orbit model which is greatly different from that of the electric propulsion-gravity transfer orbit, and the related design method can not borrow the orbit model.
The invention discloses a small-thrust gravity-assist orbit solution space shearing method, which is disclosed by the invention with the publication number of CN103020338A, aims to solve the problems that the solution space combining the small thrust and a planet gravity-assist transfer orbit is complex and has multiple extreme points in the prior art, and provides the small-thrust gravity-assist orbit solution space shearing method which can solve and obtain the effective solution space combining the small thrust and the planet gravity-assist transfer orbit. The method adopts inverse sextic polynomial to approach the low-thrust orbit, and cannot simulate the process of thrust attenuation along with the whole satellite power, so that the method cannot be suitable for the solar system marginal detection task.
In conclusion, people have insufficient attention to the research on the solar system marginal detection electric propulsion transfer orbit, and related research results are also deficient, which is not matched with the important scientific value, so that the solar system marginal detection transfer orbit needs to be further deepened, and an effective orbit design scheme is provided for task design.
Disclosure of Invention
Aiming at the defects in the prior art, the invention provides a solar system marginal detection electric propulsion transfer orbit design method and system.
According to the design method and the system for the solar system marginal detection electric propulsion transfer orbit, the scheme is as follows:
in a first aspect, a solar system marginal detection electric propulsion transfer orbit design method is provided, and the method includes:
step S1: establishing a deep space electric propulsion-gravity transfer orbit design model which comprises a planet gravity-gravity orbit model and an electric propulsion orbit model, and setting a flight sequence;
step S2: according to a set flight sequence, a detector is set to start from the earth with a resonance ratio close to 1:2, an optimization variable is guessed in a ballistic flight mode, a Sims-Flanagan model is used for orbit recursion, and relevant parameters including a launching parameter, a planet borrowing parameter, transfer time and fuel consumption are obtained;
step S3: setting relevant constraint conditions including emission C3 and gravity-assist height, and performing preliminary optimization by using an SQP algorithm according to a Sims-Flanagan model and a gravity-assist orbit model;
step S4: and according to the primary optimization result, adopting a track dynamics model, and carrying out re-optimization according to an indirect method to obtain a final solution.
Preferably, in step S1, the flight sequence is earth-wood star-sea king star.
Preferably, the step S1 specifically includes:
step S1.1: according to the characteristics that the solar propulsion thrust and the specific impulse have jumping and discontinuity, the thrust and the specific impulse are subjected to smoothing treatment, and the sum is transformed into a continuous derivable function while the precision is ensured;
step S1.2: discretizing the electric propulsion orbit, and establishing a solar-heart two-body electric propulsion Sims-Flanagan model;
step S1.3: and (4) considering the gravitational influence of the eight planets, establishing a centroidal multi-body orbit model, and deriving by adopting an indirect method to obtain a target practice equation.
Preferably, the step S2 specifically includes:
step S2.1: setting the detector to start from the earth along the earth tangential direction at a resonance ratio close to 1:2 according to the set flight sequence;
step S2.2: the thrust is set to be zero, and the earth, the wooden star and the sea king star are guessed by taking the hyperbola overspeed vector, the detector mass and the flying epoch in a ballistic flight mode.
Step S2.3: and (3) performing orbit recursion by using a Sims-Flanagan model to obtain relevant parameters including emission parameters, planet borrowing parameters, transfer time and fuel consumption, and gradually performing iterative adjustment until the continuity of the orbit is met to complete the guess of the initial value of the optimized variable.
Preferably, the step S3 specifically includes:
step S3.1: setting relevant constraints including launch C3 and height of gravity;
step S3.2: performing preliminary optimization by using an SQP algorithm by using a Sims-Flanagan model and a gravity-assisted orbit model;
step S3.3: and acquiring relevant orbit parameters including the planet borrowing force, the detector quality and the transfer time at each stage according to the initial optimization result, and obtaining an initial solution optimized by an indirect method.
Preferably, the step S4 specifically includes:
step S4.1: according to the preliminary optimization result, carrying out stage division on the transfer track, and sorting the force-borrowing parameters of each stage;
step S4.2: modeling optimization of each stage by adopting an indirect method to obtain a two-point boundary value problem of each stage;
step S4.3: and (3) performing target-shooting iteration on the two-point boundary value problem, constructing an iteration equation by adopting multivariate Newton iteration, performing cyclic calculation until the specified precision is met, and obtaining a final solution.
In a second aspect, a solar system marginal detection electric propulsion transfer orbit design system is provided, the system comprising:
module M1: establishing a deep space electric propulsion-gravity transfer orbit design model which comprises a planet gravity-gravity orbit model and an electric propulsion orbit model, and setting a flight sequence;
module M2: according to a set flight sequence, a detector is set to start from the earth with a resonance ratio close to 1:2, an optimization variable is guessed in a ballistic flight mode, a Sims-Flanagan model is used for orbit recursion, and relevant parameters including a launching parameter, a planet borrowing parameter, transfer time and fuel consumption are obtained;
module M3: setting relevant constraint conditions including emission C3 and gravity-assist height, and performing preliminary optimization by using an SQP algorithm according to a Sims-Flanagan model and a gravity-assist orbit model;
module M4: and according to the primary optimization result, adopting a track dynamics model, and carrying out re-optimization according to an indirect method to obtain a final solution.
Preferably, the module M1 specifically includes:
module M1.1: according to the fact that the solar propulsion thrust and the specific impulse have jumping and discontinuous properties, the thrust and the specific impulse are subjected to smoothing treatment, and the thrust and the specific impulse function are transformed into a continuous conductible function while the precision is ensured;
module M1.2: discretizing the electric propulsion orbit, and establishing a solar-heart two-body electric propulsion Sims-Flanagan model;
module M1.3: and (3) considering the gravitational influence of eight planets, establishing a multiple-body orbit model of the centroid, and deriving by adopting an indirect method of an optimal control theory to obtain a target practice equation, wherein the iteration variable of the equation comprises 6 initial values of the cooperative state, and the position and the speed of an iteration target at the moment of gravity are also used as reference.
Preferably, the module M2 specifically includes:
module M2.1: setting the detector to start from the earth along the earth tangential direction at a resonance ratio close to 1:2 according to the set flight sequence;
module M2.2: the thrust is set to be zero, and the earth, the wooden star and the sea king star are guessed by taking the hyperbola overspeed vector, the detector mass and the flying epoch in a ballistic flight mode.
Module M2.3: and (3) performing orbit recursion by using a Sims-Flanagan model to obtain relevant parameters including emission parameters, planet borrowing parameters, transfer time and fuel consumption, and gradually performing iterative adjustment until the continuity of the orbit is met to complete the guess of the initial value of the optimized variable.
Preferably, the module M3 specifically includes:
module M3.1: setting relevant constraints including launch C3 and height of gravity;
module M3.2: performing preliminary optimization by using an SQP algorithm by using a Sims-Flanagan model and a gravity-assisted orbit model;
module M3.3: and acquiring relevant orbit parameters including the planet borrowing force, the detector quality and the transfer time at each stage according to the initial optimization result, and obtaining an initial solution optimized by an indirect method.
Compared with the prior art, the invention has the following beneficial effects:
1. the method can effectively simulate the influence of the sun ware distance on the whole satellite power, fully considers the characteristics of stepping and discontinuity of the thrust and the specific impulse of solar electric propulsion, and provides an accurate and rapid method for the design of the solar system marginal detection transfer orbit;
2. the method realizes the balance among the track precision, the calculated amount and the design efficiency, and quickly and effectively completes the optimization design of the initial solution of the transfer track by adopting a simplified track model; and the obtained initial solution is used as an initial value guess of the high-precision multi-body model, so that the iteration times can be effectively reduced, the calculated amount can be reduced, and the robustness of numerical iteration can be ensured.
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Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a flow chart of the method of the present invention;
FIG. 2 is a solar electric propulsion-gravity orbit launched in 2024 of the Earth-Jupiter-100 AU flight sequence in an embodiment;
FIG. 3 is a solar electric propulsion optimal thrust emitted in 2024 of the Earth-Jupiter-100 AU flight sequence in an embodiment;
FIG. 4 is a solar electric propulsion-gravity orbit detector power emitted 2024 of the Earth-Jupiter-100 AU flight sequence in an embodiment;
fig. 5 shows solar electric propulsion-gravity orbit emitted by the earth-muxing-starfish-100 AU flying sequence 2030 in the embodiment.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
The embodiment of the invention provides a solar system marginal detection electric propulsion transfer orbit design method, which comprises the following specific steps of:
firstly, a deep space electric propulsion-gravity transfer orbit design model is established, wherein the deep space electric propulsion-gravity transfer orbit design model comprises a planet gravity orbit model and an electric propulsion orbit model, and a flight sequence is set to be earth-wooden star-sea king star. The process is as follows:
according to the fact that the solar propulsion thrust and the specific impulse have jumping and discontinuous properties, the thrust and the specific impulse are subjected to smoothing processing, and the thrust and the specific impulse function are transformed into a continuous derivable function while accuracy is guaranteed. Discretizing the electric propulsion orbit, and establishing a solar-heart two-body electric propulsion Sims-Flanagan model; and (3) considering the gravitational influence of eight planets, establishing a multiple-body high-precision orbit model of the centroid, and deriving by adopting an indirect method of an optimal control theory to obtain a target practice equation, wherein the iteration variable of the equation comprises 6 initial values of the cooperative state, and the position and the speed of the iteration target at the moment of gravity are also used as reference.
Secondly, according to a set flight sequence, a detector is set to start from the earth with a resonance ratio close to 1:2, optimization variables are guessed in a ballistic flight mode, orbit recursion is carried out by using a Sims-Flanagan model, and relevant parameters including emission parameters, planet borrowing parameters, transfer time and fuel consumption are obtained. Specifically, the method comprises the following steps:
the probe is set to start at approximately 1:2 resonance ratio from earth tangentially along earth according to the set flight sequence, ensuring that the probe can return to earth with less fuel penalty. And then, the thrust is set to be zero, and the earth, the wooden star and the sea king star are guessed by taking the hyperbola overspeed vector, the detector mass and the flying epoch in a ballistic flight mode. And (3) performing orbit recursion by using the Sims-Flanagan model to obtain parameters such as emission parameters, planet borrowing parameters, transfer time and fuel consumption, gradually performing iterative adjustment until the continuity of the orbit is met, and completing the guessing of the initial value of the optimized variable.
And then, setting constraint conditions such as launch C3, gravity height and the like, carrying out preliminary optimization by using an SQP algorithm according to the Sims-Flanagan model and the gravity orbit model, and then carrying out re-optimization according to an indirect method to finally obtain a high-precision solution. The specific process of the step is as follows:
and setting constraint conditions such as emission C3, gravity height and the like, a Sims-Flanagan model and a gravity orbit model, and performing preliminary optimization by using an SQP algorithm.
Acquiring orbit parameters such as planet borrowing force, detector quality, transfer time and the like in each stage according to the initial optimization result, and taking the initial solution as an indirect optimization;
and (4) carrying out re-optimization according to an indirect method, carrying out target practice iteration on the obtained two-point boundary value problem, and finally obtaining a high-precision solution.
And finally, according to the primary optimization result, adopting a high-precision orbit dynamics model, and carrying out re-optimization according to an indirect method to finally obtain a high-precision solution. The specific process of the step is as follows:
according to the preliminary optimization result, carrying out stage division on the transfer track, sorting the force-borrowing parameters of each stage, modeling the optimization of each stage by adopting an indirect method to obtain the two-point boundary value problem of each stage, carrying out target-shooting iteration on the two-point boundary value problem, constructing an iterative equation by adopting multivariate Newton iteration, carrying out cyclic calculation until the specified precision is met, and finally obtaining a high-precision solution.
The following is a numerical simulation verification of a design method of solar system marginal detection electric propulsion transfer orbit.
The mass of the detector is considered according to 1000kg, and at a distance of 1AU from the sun to the ground, the power generation capacity of the solar sailboard is 10kW, and the maximum available power of the electric propulsion system is 9.5 kW. The thrust and specific impulse of the electric propulsion system are considered in the step-by-step mode, and the step is changed along with the step-by-step mode of the available power. The earth gravity height is not less than 1000km, the wooden star gravity height is not less than 5 wooden star radiuses, and the sea king star gravity height is not less than 1000 km. Also, during the process of borrowing force, the application of pulse orbit change is not allowed. The thrust arc section is only acted on the track arc section before the momentum of the wooden star, namely the detector always keeps ballistic flight after the momentum of the wooden star. For carrying, the inclination angle of the launching orbit is required to be not less than 28.5deg, and the argument of the perigee is required to be not more than 245 deg. And during orbit optimization, selecting a gravity sequence earth-wooden star-100 AU and earth-wooden star-starfish-100 for optimization design.
The simulation calculation results refer to the following pictures, and the solar electric propulsion-gravity orbit and the magnitude of the thrust emitted in 2024 and 2030 are respectively given. Specifically, as shown in fig. 2 and 3, the solar electric propulsion-gravity orbit emitted by the earth-muxing-100 AU flight sequence in 2024 year and the solar electric propulsion optimal thrust emitted by the earth-muxing-100 AU flight sequence in 2024 year are respectively; as shown in fig. 4 and 5, the solar electric propulsion-gravity orbit detector power emitted in 2024 for the earth-mars-100 AU flight sequence and the solar electric propulsion-gravity orbit emitted in 2030 for the earth-mars-asterias-100 AU flight sequence are respectively.
The embodiment of the invention provides a design method of a solar system marginal detection electric propulsion transfer orbit, which can effectively simulate the influence of sun ware distance on the whole satellite power, fully considers the characteristics of stepping and discontinuity of the thrust and the specific impulse of solar electric propulsion, and provides an accurate and rapid method for the design of the solar system marginal detection transfer orbit. The method realizes the balance among the track precision, the calculated amount and the design efficiency, can quickly and effectively complete the initial solution of the transfer track to obtain the optimized design by adopting the simplified track model, and then can effectively reduce the iteration times, reduce the calculated amount and ensure the robustness of numerical iteration by taking the solution as the initial value guess of the high-precision multi-body model.
Those skilled in the art will appreciate that, in addition to implementing the system and its various devices, modules, units provided by the present invention as pure computer readable program code, the system and its various devices, modules, units provided by the present invention can be fully implemented by logically programming method steps in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system and various devices, modules and units thereof provided by the invention can be regarded as a hardware component, and the devices, modules and units included in the system for realizing various functions can also be regarded as structures in the hardware component; means, modules, units for performing the various functions may also be regarded as structures within both software modules and hardware components for performing the method.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (10)
1. A design method for solar system marginal detection electric propulsion transfer orbit is characterized by comprising the following steps:
step S1: establishing a deep space electric propulsion-gravity transfer orbit design model which comprises a planet gravity-gravity orbit model and an electric propulsion orbit model, and setting a flight sequence;
step S2: according to a set flight sequence, a detector is set to start from the earth with a resonance ratio close to 1:2, an optimization variable is guessed in a ballistic flight mode, a Sims-Flanagan model is used for orbit recursion, and relevant parameters including a launching parameter, a planet borrowing parameter, transfer time and fuel consumption are obtained;
step S3: setting relevant constraint conditions including emission C3 and gravity-assist height, and performing preliminary optimization by using an SQP algorithm according to a Sims-Flanagan model and a gravity-assist orbit model;
step S4: and according to the primary optimization result, adopting a track dynamics model, and carrying out re-optimization according to an indirect method to obtain a final solution.
2. The method for designing a solar system marginal survey electric propulsion transfer orbit according to claim 1, wherein the flight sequence set in step S1 is earth-wood star-sea king star.
3. The solar system marginal detection electric propulsion transfer orbit design method of claim 1, wherein the step S1 specifically comprises:
step S1.1: according to the fact that the solar propulsion thrust and the specific impulse have jumping and discontinuous properties, the thrust and the specific impulse are subjected to smoothing treatment, and the thrust and the specific impulse function are transformed into a continuous conductible function while the precision is ensured;
step S1.2: discretizing the electric propulsion orbit, and establishing a solar-heart two-body electric propulsion Sims-Flanagan model;
step S1.3: and (3) considering the gravitational influence of eight planets, establishing a multiple-body orbit model of the centroid, and deriving by adopting an indirect method of an optimal control theory to obtain a target practice equation, wherein the iteration variable of the equation comprises 6 initial values of the cooperative state, and the position and the speed of an iteration target at the moment of gravity are also used as reference.
4. The solar system marginal detection electric propulsion transfer orbit design method of claim 1, wherein the step S2 specifically comprises:
step S2.1: setting the detector to start from the earth along the earth tangential direction at a resonance ratio close to 1:2 according to the set flight sequence;
step S2.2: the thrust is set to be zero, and the earth, the wooden star and the sea king star are guessed by taking the hyperbola overspeed vector, the detector mass and the flying epoch in a ballistic flight mode.
Step S2.3: and (3) performing orbit recursion by using a Sims-Flanagan model to obtain relevant parameters including emission parameters, planet borrowing parameters, transfer time and fuel consumption, and gradually performing iterative adjustment until the continuity of the orbit is met to complete the guess of the initial value of the optimized variable.
5. The solar system marginal detection electric propulsion transfer orbit design method of claim 1, wherein the step S3 specifically comprises:
step S3.1: setting relevant constraints including launch C3 and height of gravity;
step S3.2: performing preliminary optimization by using an SQP algorithm by using a Sims-Flanagan model and a gravity-assisted orbit model;
step S3.3: and acquiring relevant orbit parameters including the planet borrowing force, the detector quality and the transfer time at each stage according to the initial optimization result, and obtaining an initial solution optimized by an indirect method.
6. The solar system marginal detection electric propulsion transfer orbit design method of claim 1, wherein the step S4 specifically comprises:
step S4.1: according to the preliminary optimization result, carrying out stage division on the transfer track, and sorting the force-borrowing parameters of each stage;
step S4.2: modeling optimization of each stage by adopting an indirect method to obtain a two-point boundary value problem of each stage;
step S4.3: and (3) performing target-shooting iteration on the two-point boundary value problem, constructing an iteration equation by adopting multivariate Newton iteration, performing cyclic calculation until the specified precision is met, and obtaining a final solution.
7. A solar system marginal detection electric propulsion transfer orbit design system is characterized by comprising:
module M1: establishing a deep space electric propulsion-gravity transfer orbit design model which comprises a planet gravity-gravity orbit model and an electric propulsion orbit model, and setting a flight sequence;
module M2: according to a set flight sequence, a detector is set to start from the earth with a resonance ratio close to 1:2, an optimization variable is guessed in a ballistic flight mode, a Sims-Flanagan model is used for orbit recursion, and relevant parameters including a launching parameter, a planet borrowing parameter, transfer time and fuel consumption are obtained;
module M3: setting relevant constraint conditions including emission C3 and gravity-assist height, and performing preliminary optimization by using an SQP algorithm according to a Sims-Flanagan model and a gravity-assist orbit model;
module M4: and according to the primary optimization result, adopting a track dynamics model, and carrying out re-optimization according to an indirect method to obtain a final solution.
8. The solar system marginal detection electric propulsion transfer track design system according to claim 7, wherein the module M1 specifically comprises:
module M1.1: according to the fact that the solar propulsion thrust and the specific impulse have jumping and discontinuous properties, the thrust and the specific impulse are subjected to smoothing treatment, and the thrust and the specific impulse function are transformed into a continuous conductible function while the precision is ensured;
module M1.2: discretizing the electric propulsion orbit, and establishing a solar-heart two-body electric propulsion Sims-Flanagan model;
module M1.3: and (3) considering the gravitational influence of eight planets, establishing a multiple-body orbit model of the centroid, and deriving by adopting an indirect method of an optimal control theory to obtain a target practice equation, wherein the iteration variable of the equation comprises 6 initial values of the cooperative state, and the position and the speed of an iteration target at the moment of gravity are also used as reference.
9. The solar system marginal detection electric propulsion transfer track design system according to claim 7, wherein the module M2 specifically comprises:
module M2.1: setting the detector to start from the earth along the earth tangential direction at a resonance ratio close to 1:2 according to the set flight sequence;
module M2.2: the thrust is set to be zero, and the earth, the wooden star and the sea king star are guessed by taking the hyperbola overspeed vector, the detector mass and the flying epoch in a ballistic flight mode.
Module M2.3: and (3) performing orbit recursion by using a Sims-Flanagan model to obtain relevant parameters including emission parameters, planet borrowing parameters, transfer time and fuel consumption, and gradually performing iterative adjustment until the continuity of the orbit is met to complete the guess of the initial value of the optimized variable.
10. The solar system marginal detection electric propulsion transfer track design system according to claim 7, wherein the module M3 specifically comprises:
module M3.1: setting relevant constraints including launch C3 and height of gravity;
module M3.2: performing preliminary optimization by using an SQP algorithm by using a Sims-Flanagan model and a gravity-assisted orbit model;
module M3.3: and acquiring relevant orbit parameters including the planet borrowing force, the detector quality and the transfer time at each stage according to the initial optimization result, and obtaining an initial solution optimized by an indirect method.
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| CN113636106A (en) * | 2021-09-15 | 2021-11-12 | 上海卫星工程研究所 | Continuous small-thrust high-orbit target track-changing approaching method and system |
| CN113741197A (en) * | 2021-09-15 | 2021-12-03 | 上海卫星工程研究所 | Single-time approaching three-pulse control method and system for high-orbit target |
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