CN110765504B - Orbit design method for rendezvous and docking of spacecraft orbits around the moon - Google Patents
Orbit design method for rendezvous and docking of spacecraft orbits around the moon Download PDFInfo
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Abstract
The invention relates to a spacecraft lunar orbit intersection butt joint orbit design method, which comprises the following steps: a. establishing a orbit maneuver mathematical model for the lunar orbit intersection butt joint of the spacecraft by adopting a descending and intersection mode of pulse orbit transfer; b. determining the height of the lunar point of the initial orbit of the rendezvous and docking of the spacecraft through phase modulation range analysis; c. estimating an initial value of the pulse orbital transfer maneuver under a two-body model, and ensuring that the initial value is quickly converged under an accurate model consisting of the two-body model and a perturbation force model; d. performing in-plane orbital transfer planning under the accurate model, and calculating orbital transfer parameters of in-plane orbital transfer; e. performing out-of-plane orbital transfer planning under the accurate model, and calculating out-of-plane orbital transfer parameters; f. and carrying out comprehensive orbital transfer planning under the accurate model, and updating orbital transfer parameters of in-plane orbital transfer after out-of-plane orbital transfer to obtain a lunar orbital rendezvous and docking orbit.
Description
Technical Field
The invention relates to the technical field of design of manned spacecraft rendezvous and docking orbits, in particular to a method for designing orbits for rendezvous and docking of a spacecraft lunar orbit, and particularly relates to a method for designing full-phase rendezvous and docking orbits for a rendezvous and docking task of the lunar orbit in the process of the spacecraft rushing to the moon.
Background
According to the experience of the Apollodenyul project, if the manned lunar task adopts a rocket with carrying capacity LEO hundred-ton grade required in a 1-time launching mode, the development difficulty of the carrier rocket is greatly increased. In order to increase the feasibility of the task, the carrying capacity of the carrier rocket can be reduced, but the manned lunar aircraft is divided into a manned spacecraft and a lunar lander 2, and the manned spacecraft and the lunar lander 2 are respectively sent into the lunar transfer orbit and are in cross joint on the lunar orbit. In order to distinguish the intersection butt joint of the lunar ascent of the lunar landing capsule and the lunar orbit of the manned spacecraft after the lunar landing task is completed in the Apollodenyum project, the lunar orbit intersection butt joint which is carried out after the lunar lander and the manned spacecraft complete the lunar transfer and capture the lunar orbit is called as lunar orbit intersection butt joint in the lunar-rushing process.
The lunar orbit rendezvous and docking is a new rendezvous and docking type in the moon-running process, and has no precedent in the world and no rendezvous and docking orbit design scheme for reference. In addition, the initial state of the lunar orbit intersection butt joint in the lunar running process is a lunar orbit captured by the lunar lander and the manned spacecraft after the lunar transfer is completed successively. The earth-moon transfer time of the lunar lander and the manned spacecraft is determined by the earth-moon transfer orbit design, and the initial phase difference of the lunar lander and the manned spacecraft entering the lunar orbit is uncertain, so that the lunar orbit intersection butt joint in the lunar process can adapt to the change of the initial phase difference in a large range.
Currently, only the American Arboloden moon project implements the lunar orbit intersection docking task with the rising moon surface, but the requirements of the type of intersection docking and the lunar orbit intersection docking task in the lunar process are completely different. The related core technology of near-earth rail intersection butt joint is mastered at present in China; in the project of the third month exploration, an unmanned lunar orbit intersection docking technology is broken through, but the lunar orbit intersection docking technology is also a lunar orbit intersection docking in the rising and returning process from the lunar surface, and a design method for lunar orbit intersection docking in the lunar process is not disclosed. Therefore, a design method for the lunar orbit rendezvous and docking orbit in the lunar running process is urgently needed to be researched, and a technical basis is provided for the design of a scheme of the lunar orbit rendezvous and docking plan rendezvous and docking orbit of people in the future in China.
Disclosure of Invention
The invention aims to solve the problems and provides a full-phase rendezvous and docking track design method for the rendezvous and docking task of the lunar orbit in the lunar process.
In order to achieve the aim, the invention provides an orbit design method for rendezvous and docking of a spacecraft lunar orbit, which comprises the following steps:
a. establishing a orbit maneuver mathematical model for the lunar orbit intersection butt joint of the spacecraft by adopting a descending and intersection mode of pulse orbit transfer;
b. determining the height of the lunar point of the initial orbit of the rendezvous and docking of the spacecraft through phase modulation range analysis;
c. estimating an initial value of the pulse orbital transfer maneuver under a two-body model, and ensuring that the initial value is quickly converged under an accurate model consisting of the two-body model and a perturbation force model;
d. performing in-plane orbital transfer planning under the accurate model, and calculating orbital transfer parameters of in-plane orbital transfer;
e. performing out-of-plane orbital transfer planning under the accurate model, and calculating out-of-plane orbital transfer parameters;
f. and carrying out comprehensive orbital transfer planning under the accurate model, and updating orbital transfer parameters of in-plane orbital transfer after out-of-plane orbital transfer to obtain a lunar orbital rendezvous and docking orbit.
According to one aspect of the invention, a 5-pulse orbital transfer drop intersection is used, wherein the 1 st, 3 rd and 4 th orbital transfers are in-plane orbital transfers, the 2 nd orbital transfer is out-of-plane orbital transfers, and the 5 th orbital transfer is comprehensive orbital transfer.
According to an aspect of the invention, in the a step:
the 1 st orbital transfer is carried out at a near moon point, and the height of a far moon point is reduced; the 2 nd orbital transfer is a track surface correction maneuver;
the 3 rd orbital transfer is carried out at a moon point, and the height of a moon point is reduced;
the 4 th orbital transfer is carried out near the moonpoint, and the orbit enters a rounded orbit to meet the requirement of a remote guidance terminal;
the 5 th orbital transfer is a combined correcting maneuver and is not implemented in a nominal state;
and setting orbital transfer times according to the total flight time at orbital transfer intervals of 3 circles or more every time, establishing a motor mathematical model of the rendezvous and docking orbit according to a 5-pulse orbital-descending orbital transfer mode, and inputting 5 orbital transfer parameters to calculate the terminal position and speed error of the active spacecraft relative to the passive spacecraft.
According to an aspect of the invention, in the step b, the mathematical model is used to adjust the initial orbit lunar point height so that the phase modulation range covers a phase angle of 0-360 degrees.
According to one aspect of the invention, in the step c, the determined initial orbit lunar point height is used as the intersection butt joint initial orbit lunar point height, intersection butt joint orbit simulation is performed under a two-body model, an intersection butt joint initial phase angle is calculated according to an intersection butt joint initial orbit and a target orbit, a phase angle of a terminal time is calculated according to a terminal condition, an intersection butt joint phase adjusting angle is obtained, phase adjusting angle constraint is met by adjusting the lunar point height after 1 st orbital transfer, and orbital transfer time and orbital transfer speed increment of 1 st, 3 rd and 4 th orbital transfer are calculated, and the 2 nd orbital transfer speed increment is 0 because the simulation model is a two-body model and has no out-of-plane perturbation.
According to one aspect of the invention, in the step d, the 1 st orbital transfer speed increment, the 3 rd orbital transfer speed increment, the 4 th orbital transfer time and the speed increment calculated in the step c are used as initial values of control variables, target constraints are set as radial and transverse position errors and radial and transverse speed errors of the terminal time relative to the remote guidance terminal condition, the orbital transfer times of the 1 st orbital transfer and the 3 rd orbital transfer are respectively determined according to features of a near moon point and a far moon point, the 2 nd orbital transfer speed increment is set as 0, orbit simulation is carried out under an accurate model, and the 1 st orbital transfer speed increment, the 3 rd orbital transfer speed increment, the 4 th orbital transfer time and the speed increment are continuously corrected through a differential correction method to adjust the in-plane terminal condition until the in-plane terminal condition constraint is met.
According to one aspect of the invention, in the step e, the 1 st orbital transfer time and speed increment, the 3 rd orbital transfer time and speed increment, and the 4 th orbital transfer time and speed increment calculated in the step d are used as known quantities, the 2 nd orbital transfer time and speed increment are used as control variables, target constraints are set as a normal position error of the terminal time relative to the remote guidance terminal condition and a normal speed error, and the 2 nd orbital transfer time and speed increment are continuously corrected through a differential correction method to adjust the out-of-plane terminal condition until the out-of-plane terminal condition constraint is met.
According to one aspect of the invention, in the step f, the 1 st orbital transfer time and speed increment, the 2 nd orbital transfer time and speed increment calculated in the step e are used as known quantities, the 3 rd orbital transfer speed increment, the 4 th orbital transfer time and speed increment are used as initial values of control variables, target constraints are set as radial and transverse position errors of the terminal time relative to the remote guidance terminal condition and radial and transverse speed errors, the orbital transfer time of the 3 rd orbital transfer is determined according to far moon point characteristics, and the 3 rd orbital transfer speed increment, the 4 th orbital transfer time and speed increment are continuously corrected by a differential correction method to adjust the in-plane terminal condition until the in-plane terminal condition constraint is met.
According to the orbit design method for the rendezvous and docking of the lunar orbits of the spacecraft, the height of the lunar point of the active spacecraft, which enters the rendezvous and docking initial orbit during the lunar braking, is obtained through phase modulation capability analysis, the initial value design of the orbital transfer maneuver is carried out on a two-body model according to a multi-pulse orbital transfer strategy, and the accurate orbital transfer maneuver value is gradually obtained in a high-precision shooting model through differential correction and multilayer iteration, so that the rendezvous and docking orbit of the lunar orbits meeting the full phase is obtained.
According to the orbit design method for the lunar orbit intersection butt joint of the spacecraft, the solving speed is high, the convergence is good, the robustness is strong, and the phase modulation requirement of 0-360 degrees is met.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings needed in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art to obtain other drawings without creative efforts.
FIG. 1 schematically represents a flow chart of a method of orbit design for spacecraft lunar orbit rendezvous and docking in accordance with the present invention;
FIG. 2 is a schematic diagram of a lunar process lunar orbit intersection docking 5 pulse orbit-lowering and transferring scheme;
FIG. 3 is a graph of the relationship between the lunar orbit intersection phase modulation range and the initial orbit lunar point height;
FIG. 4 is a flight trajectory diagram of the active and passive spacecrafts under the lunar center inertial system in the embodiment;
FIG. 5 is a flight path diagram of the active and passive spacecrafts under the inertial system of the geocentric J2000 in the embodiment;
FIG. 6 is a graph of the change of the lunar center height of the active spacecraft in the embodiment.
Detailed Description
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the embodiments will be briefly described below. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
In describing embodiments of the present invention, the terms "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like are used in an orientation or positional relationship that is based on the orientation or positional relationship shown in the associated drawings, which is for convenience and simplicity of description only, and does not indicate or imply that the referenced device or element must have a particular orientation, be constructed and operated in a particular orientation, and thus, the above-described terms should not be construed as limiting the present invention.
The present invention is described in detail below with reference to the drawings and the specific embodiments, which are not repeated herein, but the embodiments of the present invention are not limited to the following embodiments.
Fig. 1 schematically shows a flow chart of an orbit design method for spacecraft lunar orbit rendezvous and docking according to the invention. As shown in fig. 1, the orbit design method for the rendezvous and docking of the lunar orbits of the spacecraft according to the invention comprises the following steps:
a. establishing a orbit maneuver mathematical model for the lunar orbit intersection butt joint of the spacecraft by adopting a descending and intersection mode of pulse orbit transfer;
b. determining the height of the lunar point of the initial orbit of the rendezvous and docking of the spacecraft through phase modulation range analysis;
c. estimating an initial value of the pulse orbital transfer maneuver under a two-body model, and ensuring that the initial value is quickly converged under an accurate model consisting of the two-body model and a perturbation force model;
d. performing in-plane orbital transfer planning under the accurate model, and calculating orbital transfer parameters of in-plane orbital transfer;
e. performing out-of-plane orbital transfer planning under the accurate model, and calculating out-of-plane orbital transfer parameters;
f. and carrying out comprehensive orbital transfer planning under the accurate model, and updating orbital transfer parameters of in-plane orbital transfer after out-of-plane orbital transfer to obtain a lunar orbital rendezvous and docking orbit.
According to one embodiment of the present invention, a 5-pulse orbital transfer drop intersection is used, wherein the 1 st, 3 rd and 4 th orbital transfers are in-plane orbital transfers, the 2 nd orbital transfer is out-of-plane orbital transfers, and the 5 th orbital transfer is comprehensive orbital transfer.
Based on the above embodiment, in the step a: as shown in fig. 2, in the lunar-driving process, a lunar orbit intersection docking task adopts a rail descending intersection scheme to save propellant consumption, and the active spacecraft adopts a 5-pulse maneuvering strategy to complete remote intersection with the passive spacecraft so as to meet the remote guidance terminal conditions. The 1 st orbital transfer is carried out at a near moon point, and the height of a far moon point is reduced; the 2 nd orbital transfer is a track surface correction maneuver; the 3 rd orbital transfer is carried out at a moon point, and the height of a moon point is reduced; the 4 th orbital transfer is carried out near the moonpoint, and the orbit enters a rounded orbit to meet the requirement of a remote guidance terminal; the 5 th orbital transfer is a combined correcting maneuver, not implemented in the nominal state. The track changing interval is 3 circles and more each time. And setting orbital transfer circles according to the total flight time, establishing a motor mathematical model of the rendezvous and docking orbit according to a 5-pulse orbital-falling orbital transfer strategy, and inputting 5 orbital transfer parameters to calculate the terminal position and speed error of the active spacecraft relative to the passive spacecraft.
In the step b, the initial cross-joint is determined by phase modulation analysisThe orbit is high at the moon point, and the full phase modulation capability is realized. For the lunar orbit intersection docking task in the lunar running process, the target orbit is a certain height H2The circular moon orbit; the initial orbit is a moon-shaped elliptical orbit, and the height H of the moon-shaped pointp10Generally slightly greater than the target track height, the moon point height Ha10The adjustment can be carried out according to the determination of the near-moon braking of the active spacecraft. And a moon point height Ha10Directly determines the phase modulation range of the whole rendezvous and docking process. Adjusting the height H of the lunar point of the initial orbit by adopting the dynamic mathematical model of the lunar orbit intersection butt joint orbit established in the step aa10The phase modulation range covers 0-360 DEG phase angle.
In the step c, the H determined in the step b is addeda10As the initial rail moon height of the rendezvous dock. In order to quickly calculate the initial value of orbital transfer, the simulation of the rendezvous and butt joint orbit is carried out under a two-body model. And calculating a rendezvous initial phase angle delta according to the rendezvous initial orbit and the target orbit. And calculating a phase angle alpha of the terminal time according to the terminal condition, thereby obtaining a rendezvous and docking phase adjusting angle delta theta which is delta-alpha. By adjusting the height H of the moon point after the 1 st orbital transfera11Satisfying the constraint of phase modulation angle delta theta and calculating the track change time t of the 1 st, 3 rd and 4 th track change1,t3,t4And orbital transfer velocity increment Δ V1,ΔV3,ΔV4. Since the simulation model is a two-body model at this time and has no out-of-plane perturbation, the 2 nd orbital transfer velocity increment is delta V2＝0。t1,t3Respectively according to the characteristics of the near moon point and the far moon point. Therefore, it mainly calculates Δ V1,ΔV3,t4,ΔV4。
In the step d, the 1 st orbital transfer speed increment delta V calculated in the step c is obtained13 rd orbital transfer speed increment delta V3The 4 th track transfer time and the speed increment t4,ΔV4As initial values of the control variables, the target constraints are set to radial and lateral position errors δ R of the terminal time versus the remote pilot terminal conditionsr,δRtAnd radial and lateral velocity error δ Vr,δVt. Orbital transfer for 1 st orbital transfer and 3 rd orbital transferTime t1,t3Respectively determining according to the characteristics of the near moon point and the far moon point, and setting the 2 nd orbital transfer speed increment as delta V 20. Performing orbit simulation under an accurate model, and continuously correcting delta V by a differential correction method1、ΔV3、t4And Δ V4And adjusting the in-plane terminal condition until the in-plane terminal condition constraint is met.
In the step e, the 1 st track change time and the speed increment t calculated in the step d are obtained1,ΔV1The 3 rd track transfer time and the speed increment t3,ΔV3The 4 th track transfer time and the speed increment t4,ΔV4As known quantities, the 2 nd track change time and the speed increment t2,ΔV2As a control variable, the target constraint is set to the normal position error δ R of the terminal time versus the remote pilot terminal conditionsnAnd normal velocity error δ Vn. Constantly correcting t by differential correction2,ΔV2And adjusting out-of-plane terminal conditions until out-of-plane terminal condition constraints are met.
In the step f, the 1 st track change time and the speed increment t calculated in the step 5 are obtained1,ΔV 12 nd orbital transfer time and speed increment t2,ΔV2As a known quantity, the 3 rd track change speed is increased by Δ V3The 4 th track transfer time and the speed increment t4,ΔV4As initial values of the control variables, the target constraints are set to radial and lateral position errors δ R of the terminal time versus the remote pilot terminal conditionsr,δRtAnd radial and lateral velocity error δ Vr,δVt. Track change time t of 3 rd track change3Respectively determined according to the characteristics of the moon points. Continuously correcting delta V by differential correction method3、t4And Δ V4And adjusting the in-plane terminal condition until the in-plane terminal condition constraint is met. Finally, the lunar orbit intersection butt joint orbit and orbital transfer strategy in the lunar process under the accurate model can be obtained.
According to the above embodiment of the present invention, the present invention further provides the following specific example:
according to the step a of the present invention, the total number of turns of the remote guidance section is set to be about 21 turns; the 1 st to 5 th orbital transfer times are respectively set as 4, 10, 15, 19 and 20; the terminal condition is that the radial transverse normal relative position is [ 10500 ] km. And establishing a track maneuvering mathematical model according to a track transfer strategy.
According to the step b of the present invention, the rendezvous and docking target track is set to be H2100km circular orbit. According to the remote guidance terminal condition, the height of the near moon point of the initial rail of the rendezvous and docking can be set to be H p10120 km. Calculate Ha10When the phase modulation range of the rendezvous and docking task is changed to 120-281 km, the phase modulation range is shown in figure 3. As can be seen from FIG. 3, when H isa10When the phase modulation capacity is 281km, the phase modulation capacity is 181.8-542.2 degrees, which corresponds to the phase modulation ranges of 181.8-360 degrees and 0-182.2 degrees, so that any initial phase angle between 0-360 degrees can be modulated.
According to the step c of the present invention, the height of the lunar point of the initial rail of the rendezvous and docking is set as H a10281 km. The orbit parameters of the active spacecraft and the passive spacecraft at the initial rendezvous and docking moment are shown in the following table 1, and the initial phase angle delta can be calculated to be 250 degrees; and calculating the terminal phase angle alpha to be 1.55 degrees according to the remote guidance terminal condition, and further obtaining the intersection docking task phase angle delta theta to be 248.45 degrees. Calculating the height H of the moon point after the 1 st orbital transfer under a two-body modela11When 149.988km, the phase modulation angle constraint is satisfied, and further calculation results in delta V1＝-26.909m/s，ΔV3＝-2.185m/s，t4＝131918.447s，ΔV4＝-8.698m/s。
The orbit parameters at the initial moment of the lunar orbit intersection docking are as the following table 1:
| parameters of the track | Initial time | Initial time |
| Aircraft with a flight control device | Passive spacecraft | Active spacecraft |
| Time of day | 1Jan202012:00:00.000 | 1Jan202012:00:00.000 |
| Height/km at |
100 | 120 |
| Height/km of |
100 | 281 |
| Track inclination deg | 160 | 160 |
| Elevation crossing right ascension deg | 60 | 60 |
| Lunar point argument deg | 0 | 300 |
| True proximal angle deg | 50 | 0 |
TABLE 1
According to the invention, step d, step c is divided into Δ V1＝-26.909m/s，ΔV3＝-2.185m/s，t4＝131918.447s，ΔV4-8.698m/s as initial value of control variable; target constraint set to δ Rr≤10m，δRt≤100m，δVr≤0.02m/s，δVtLess than or equal to 0.02 m/s. Through differential correction, the control variable delta V is obtained by converging only through iteration of 4 steps1＝-27.038m/s，ΔV3＝-2.347m/s，t4＝131893.497s，ΔV4-8.535 m/s. Obtaining t from the characteristics of the near moon point and the far moon point1＝22962.050s，t3＝106535.198s。
According to the above-mentioned e step of the present invention, t in the step d is1＝22962.050s，ΔV1＝-27.038m/s，t3＝106535.198s，ΔV3＝-2.347m/s，t4131893.497s and Δ V4-8.535m/s as known; setting the 2 nd orbital transfer parameter initial value as t2＝64800s，ΔV2-0.2m/s and as control variable; target constraint set to δ Rn≤10m，δVnLess than or equal to 0.01 m/s. Through differential correction, the convergence is realized only by iterating 4 steps to obtain a control variable t2＝65740.311s，ΔV2＝-1.640m/s。
According to the above f step of the present invention, t in e step1＝22962.050s，ΔV1＝-27.038m/s，t2＝65740.311s，ΔV2-1.640m/s as known; setting the initial value of the 3 rd and 4 th orbital transfer parameters as delta V3＝-2.347m/s，t4＝131893.497s，ΔV4-8.535m/s as a control variable; target constraint set to δ Rr≤10m，δRt≤100m，δVr≤0.02m/s，δVtLess than or equal to 0.02 m/s. Through differential correction, the control variable delta V is obtained by converging only through iteration of 1 step3＝-2.346m/s，t4＝131893.471s，ΔV4-8.539 m/s. Obtaining t from the moon point features3106535.167 s. Finally, the orbit parameters of the active spacecraft and the passive spacecraft during the lunar orbit intersection docking can be obtained, and the active spacecraftAnd 4 times of orbital transfer parameters of the orbital transfer.
Fig. 4 shows the flight trajectory of the active and passive spacecraft in the lunar inertial system, which corresponds to the flight scenario in fig. 2. Fig. 5 shows the flight trajectory of the active and passive spacecrafts in the earth center J2000 inertial system, and it can be known from fig. 5 that the trajectory of the active and passive spacecrafts is a spiral line moving along with the moon in the earth center inertial system, and conforms to the earth-moon operation rule. Fig. 6 shows the height change of the active spacecraft during the rendezvous and docking, and the influence of 1 st, 3 th and 4 th orbital transfer on the orbit height can be seen, so that the 5-time pulse orbital-lowering maneuvering strategy is met.
According to the orbit design method for the rendezvous and docking of the lunar orbits of the spacecraft, the height of the lunar point of the active spacecraft, which enters the rendezvous and docking initial orbit during the lunar braking, is obtained through phase modulation capability analysis, the initial value design of the orbital transfer maneuver is carried out on a two-body model according to a multi-pulse orbital transfer strategy, and the accurate orbital transfer maneuver value is gradually obtained in a high-precision shooting model through differential correction and multilayer iteration, so that the rendezvous and docking orbit of the lunar orbits meeting the full phase is obtained.
According to the orbit design method for the lunar orbit intersection butt joint of the spacecraft, the solving speed is high, the convergence is good, the robustness is strong, and the phase modulation requirement of 0-360 degrees is met.
The above description is only one embodiment of the present invention, and is not intended to limit the present invention, and it is apparent to those skilled in the art that various modifications and variations can be made in the present invention. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (1)
1. A spacecraft lunar orbit intersection butt joint orbit design method comprises the following steps:
a. establishing an orbit maneuver mathematical model for the rendezvous and docking of the lunar orbits of the spacecraft by adopting a 5-pulse orbital transfer orbit descending rendezvous mode;
the 1 st, 3 rd and 4 th orbital transfer is in-plane orbital transfer, the 2 nd orbital transfer is out-of-plane orbital transfer, and the 5 th orbital transfer is comprehensive orbital transfer;
the 1 st orbital transfer is carried out at a near moon point, and the height of a far moon point is reduced; the 2 nd orbital transfer is a track surface correction maneuver;
the 3 rd orbital transfer is carried out at a moon point, and the height of a moon point is reduced;
the 4 th orbital transfer is carried out near the moonpoint, and the orbit enters a rounded orbit to meet the requirement of a remote guidance terminal;
the 5 th orbital transfer is a combined correcting maneuver and is not implemented in a nominal state;
setting orbital transfer times according to total flight time at orbital transfer intervals of 3 circles or more every time, establishing a motor mathematical model of the rendezvous and docking orbit according to a 5-pulse orbital-lowering orbital transfer mode, inputting 5 orbital transfer parameters, and calculating to obtain a terminal position and a speed error of the active spacecraft relative to the passive spacecraft;
b. determining the height of the lunar point of the initial orbit of the rendezvous and docking of the spacecraft through phase modulation range analysis;
adjusting the height of the initial orbit lunar point by adopting the mathematical model to enable the phase modulation range to cover a phase angle of 0-360 degrees;
c. estimating an initial value of the pulse orbital transfer maneuver under a two-body model, and ensuring that the initial value is quickly converged under an accurate model consisting of the two-body model and a perturbation force model;
taking the determined initial orbit lunar point height as the intersection butt joint initial orbit lunar point height, performing intersection butt joint orbit simulation under a two-body model, calculating an intersection butt joint initial phase angle according to the intersection butt joint initial orbit and a target orbit, calculating a phase angle of a terminal moment according to terminal conditions to obtain an intersection butt joint phase adjusting angle, satisfying phase adjusting angle constraint by adjusting the lunar point height after 1 st orbital transfer, and calculating orbital transfer moments and orbital transfer speed increment of 1 st, 3 rd and 4 th orbital transfer, wherein the 2 nd orbital transfer speed increment is 0 because the simulation model is a two-body model and has no out-of-plane perturbation at the moment;
d. performing in-plane orbital transfer planning under the accurate model, and calculating orbital transfer parameters of in-plane orbital transfer;
taking the 1 st orbital transfer speed increment, the 3 rd orbital transfer speed increment, the 4 th orbital transfer time and the speed increment obtained by calculation in the step c as initial values of control variables, setting target constraints as radial and transverse position errors and radial and transverse speed errors of the terminal time relative to a remote guide terminal condition, respectively determining the orbital transfer time of the 1 st orbital transfer and the 3 rd orbital transfer according to characteristics of a moonpoint and a moonpoint, setting the 2 nd orbital transfer speed increment as 0, performing track simulation under an accurate model, and continuously correcting the 1 st orbital transfer speed increment, the 3 rd orbital transfer speed increment, the 4 th orbital transfer time and the speed increment through a differential correction method to adjust the in-plane terminal condition until the in-plane terminal condition constraint is met;
e. performing out-of-plane orbital transfer planning under the accurate model, and calculating out-of-plane orbital transfer parameters;
d, taking the 1 st orbital transfer time and speed increment, the 3 rd orbital transfer time and speed increment and the 4 th orbital transfer time and speed increment obtained by calculation in the step d as known quantities, taking the 2 nd orbital transfer time and speed increment as control variables, setting target constraints as a normal position error and a normal speed error of the terminal time relative to a remote guidance terminal condition, and continuously correcting the 2 nd orbital transfer time and speed increment through a differential correction method to adjust the out-of-plane terminal condition until the out-of-plane terminal condition constraint is met;
f. carrying out comprehensive orbital transfer planning under the accurate model, and updating orbital transfer parameters of in-plane orbital transfer after out-of-plane orbital transfer to obtain a lunar orbital rendezvous and docking orbit;
and e, taking the 1 st orbital transfer time and speed increment and the 2 nd orbital transfer time and speed increment obtained by calculation in the step e as known quantities, taking the 3 rd orbital transfer speed increment, the 4 th orbital transfer time and speed increment as initial values of control variables, setting target constraints as radial and transverse position errors and radial and transverse speed errors of the terminal time relative to the remote guidance terminal condition, respectively determining the orbital transfer time of the 3 rd orbital transfer according to the characteristics of the moon point, and continuously correcting the 3 rd orbital transfer time increment, the 4 th orbital transfer time and the speed increment through a differential correction method to adjust the in-plane terminal condition until the in-plane terminal condition constraint is met.
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| CN113128032B (en) * | 2021-04-01 | 2022-09-16 | 北京航空航天大学 | Intersection point time and position solving algorithm based on orbit analysis perturbation solution |
| CN113311854A (en) * | 2021-05-19 | 2021-08-27 | 北京空间飞行器总体设计部 | Fixed-point landing orbit design method in lunar sampling return task |
| CN113656939B (en) * | 2021-07-08 | 2023-12-26 | 中国人民解放军63919部队 | Manned month-entering track design method based on month-surrounding track |
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