CN117125271A - Method and system for entering large elliptic frozen orbit - Google Patents
Method and system for entering large elliptic frozen orbit Download PDFInfo
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- CN117125271A CN117125271A CN202311134429.5A CN202311134429A CN117125271A CN 117125271 A CN117125271 A CN 117125271A CN 202311134429 A CN202311134429 A CN 202311134429A CN 117125271 A CN117125271 A CN 117125271A
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- 238000000034 method Methods 0.000 title claims abstract description 33
- 238000012546 transfer Methods 0.000 claims abstract description 37
- 230000008859 change Effects 0.000 claims abstract description 30
- 230000008014 freezing Effects 0.000 claims abstract description 18
- 238000007710 freezing Methods 0.000 claims abstract description 18
- 230000001965 increasing effect Effects 0.000 claims abstract description 9
- 239000003380 propellant Substances 0.000 claims description 22
- 238000005259 measurement Methods 0.000 claims description 15
- 230000000694 effects Effects 0.000 claims description 10
- 238000009434 installation Methods 0.000 claims description 10
- 230000002411 adverse Effects 0.000 claims description 7
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 claims description 7
- 230000008569 process Effects 0.000 claims description 6
- 239000012634 fragment Substances 0.000 claims description 2
- 235000015110 jellies Nutrition 0.000 claims 1
- 239000008274 jelly Substances 0.000 claims 1
- 238000013461 design Methods 0.000 description 3
- 230000002159 abnormal effect Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 2
- 239000000446 fuel Substances 0.000 description 2
- 230000007774 longterm Effects 0.000 description 2
- 238000012423 maintenance Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000005855 radiation Effects 0.000 description 2
- 230000001133 acceleration Effects 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 230000001186 cumulative effect Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000011835 investigation Methods 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 238000012544 monitoring process Methods 0.000 description 1
- 230000003287 optical effect Effects 0.000 description 1
- 238000000926 separation method Methods 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/242—Orbits and trajectories
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Abstract
The application provides a large ellipse freezing track rail-entering method and a large ellipse freezing track rail-entering system, which comprise the steps of multiple times of low-thrust rail-changing and elevation of a near-place height at a far place; naturally increasing the perigee amplitude angle by utilizing the perturbation forward factor; the working time of each thruster is equally and reasonably distributed; the inertia pointing mode is adopted to finish each ignition and track change; forward biasing the target track near-spot height; performing low-thrust orbit transfer near the far site to reduce the height of the near site to a target value, and finally completing the orbit transfer of the satellite; the ignition height is selected to be above 28000 km. The satellite can still enter the task orbit under the fault condition and the service life of the satellite is ensured.
Description
Technical Field
The application belongs to the technical field of track entering, and particularly relates to a large oval frozen track entering method and system. In particular to an emergency track-in method and system through a 10N thruster when a 490N high-thrust engine cannot be started due to track faults.
Background
The large elliptic frozen orbit is used as a special orbit and has the characteristic of long-term residence in a target area, so that continuous observation of the target area can be realized. Along with the development of aerospace technology, the application of the space flight vehicle as a task orbit is also increasing, and the space flight vehicle is widely applied to the fields of electronic investigation, meteorological monitoring and the like. The satellite cannot directly enter a large elliptical freezing orbit due to the limitation of carrying capacity, the orbit is required to be changed by means of a 490N engine as a single-point execution mechanism for changing the orbit, and if the satellite is in failure in the orbit and cannot be used normally, a subversion problem is brought to a task and even the task fails.
Disclosure of Invention
Aiming at the defects in the prior art, the application aims to provide a large oval frozen track entering method and system.
The application provides a large oval frozen track rail-entering method, which comprises the following steps:
s1, carrying out multiple times of low-thrust orbit transfer on a far site to raise the height of the near site, so as to avoid adverse factors of atmospheric resistance, space fragments and atomic oxygen on satellites;
s2, enabling the satellite to freely run in orbit for a plurality of months, naturally increasing the perigee amplitude angle by utilizing a perturbation forward factor, and saving the propellant required for changing the perigee amplitude angle;
step S3, equally distributing working time of each thruster according to the track-changing speed increment requirement, and guaranteeing service life of each thruster and subsequent track-keeping requirement;
s4, completing ignition and track change each time in an inertial pointing mode according to the installation direction of each thruster;
s5, forward biasing the near-site height of the target track, wherein the rest parameters take the task track parameters as target values, so that the time length of the measurement and control arc segment after ignition meets the requirements and is captured in the far-site track subsequently;
s6, carrying out low-thrust orbit transfer near the far site to reduce the height of the near site to a target value, and finally completing the orbit transfer of the satellite;
and S7, in order to avoid the influence of the single event effect of the space environment on the track change process, the ignition height is selected to be more than 28000 km.
Preferably, in the step S1, the near-site height of the satellite orbit is raised to 1000 km-1100 km, so that the next orbit change is implemented after the satellite is safe.
Preferably, in said step S2, when the satellite orbit inclination is less than the critical inclination of 63.4 degrees, the positive increasing effect of the earth perturbation on the perigee amplitude is utilized:
w is the near-spot amplitude angle, n is the track average angular velocity, R e Is the equatorial radius of the earth, e is the orbital eccentricity, i is the orbital tilt angle, J 2 Is a harmonic term coefficient, t is time, a is the semi-long axis of the satellite orbit; and when the amplitude perturbation of the near-spot increases to an acceptable range, starting to change the track of the next stage.
Preferably, the acceptable range means that the consumption of the next-stage orbit transfer propellant can meet the service life of the task and the working time of the thruster is not exceeded.
Preferably, in said step S3, when the next phase of orbital transfer is performed, in the case where the orbital transfer speed increment Δv has been obtained, the total weight of the propellant to be injected for the orbital transfer is obtained, and the cumulative working time required for the orbital transfer of all the thrusters of the satellite is
Δm represents the total weight of propellant injected required to obtain a change of track;
dm represents the mass of propellant injected per unit time by the thruster.
Preferably, in the step S4, according to the track-changing speed increment direction, in combination with the installation direction and the use time requirement of each thruster, the satellite needs to establish an inertial pointing posture, so as to ensure that the thrust direction provided by the thruster is along the speed increment direction, and track-changing is completed by adopting an inertial pointing mode.
Preferably, in the step S5, the target track height is biased positively according to the requirements of measurement and control of the arc segment duration after ignition, and the remaining track parameters all target the task track parameters and enter the track to be captured.
Preferably, in the step S6, the satellite orbit is finally completed by igniting with a small thrust in the vicinity of the far spot for performing the low thrust orbit determination and lowering the near-spot altitude to the target value.
The application provides a large elliptical freezing track rail-entering system, which comprises:
the module M1 enables the satellite to orbit and raise the height of the near site for multiple times with low thrust, so as to avoid adverse factors of atmospheric resistance, space debris and atomic oxygen to the satellite;
the module M2 enables the satellite to freely run in orbit for a plurality of months, and the perturbation forward factor is utilized to naturally increase the perigee amplitude angle, so that the propellant required for changing the perigee amplitude angle is saved;
the module M3 enables the satellite to equally distribute the working time of each thruster according to the track-changing speed increment requirement, so as to ensure the service life of each thruster and the subsequent track-keeping requirement;
the module M4 enables the satellite to complete each ignition and orbit change in an inertial pointing mode according to the installation direction of each thruster;
the module M5 enables the satellite to forward bias the height of the near-site of the target orbit, and other parameters take the task orbit parameters as target values, so that the time length of the measurement and control arc section after ignition meets the requirements and is captured in the far-site orbit subsequently;
the module M6 enables the satellite to perform low-thrust orbit transfer near the far site to reduce the near site height to a target value, and finally completes the orbit transfer of the satellite;
and the module M7 is used for selecting the satellite ignition height to be more than 28000km in order to avoid the influence of the single event effect in the space environment on the orbit transfer process.
Preferably, the method for controlling the satellites by adopting the large ellipse freezing orbit is adopted.
Compared with the prior art, the application has the following beneficial effects:
1. according to the application, the earth perturbation effect is positively utilized, and under the condition that the abnormal of the 490N high-thrust engine causes the specific impulse loss, the propellant is saved by utilizing the perturbation, so that the satellite can still enter a task orbit under the fault condition, and the service life of the satellite is ensured;
2. the application comprehensively considers the service life requirement of the thruster, not only can enable the satellite to enter the preset orbit, but also can ensure the long-term use requirement of the thruster after the satellite enters the preset orbit;
3. the method comprehensively considers the requirements of measuring and controlling the arc section length before and after ignition in engineering implementation, has high engineering feasibility and can guide actual tasks;
4. the application comprehensively considers the influence of the space environment on the orbit change and can ensure the safety of the satellite during the orbit change.
Drawings
Other features, objects and advantages of the present application will become more apparent upon reading of the detailed description of non-limiting embodiments, given with reference to the accompanying drawings in which:
FIG. 1 is a flow chart of the steps of the present application.
Detailed Description
The present application will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the present application, but are not intended to limit the application in any way. It should be noted that variations and modifications could be made by those skilled in the art without departing from the inventive concept. These are all within the scope of the present application.
The application provides a large ellipse freezing orbit entering method capable of saving fuel and ensuring measurement and control simultaneously, which provides an emergency large ellipse freezing orbit entering method when 490N large thrust engines cannot be started due to orbit faults by reasonably utilizing the forward effect of orbit perturbation and considering the influences of various factors such as orbit service life, thruster service life, satellite mission section service life, propellant allowance, measurement and control arc sections before and after orbit change, space environment and the like, and can provide references for satellite orbit change implementation scheme design under abnormal conditions, thereby filling the blank of the design method of the emergency orbit entering scheme of the large ellipse freezing orbit.
As shown in fig. 1, the method for entering the large oval frozen track provided by the application is a method for entering the large oval frozen track with fuel saving and measurement and control protection simultaneously, and comprises the following steps:
s1, carrying out multiple times of low-thrust orbit transfer on a distant place to raise the height of the near place, and avoiding adverse factors of atmospheric resistance, space debris, atomic oxygen and the like on satellites;
s2, enabling the satellite to freely run in orbit for a plurality of months, naturally increasing the perigee amplitude angle by utilizing a perturbation forward factor, and saving the propellant required for changing the perigee amplitude angle;
step S3, according to the track-changing speed increment requirement, the working time of each thruster is equally and reasonably distributed, and the service life of each thruster and the subsequent on-track maintenance requirement are ensured;
s4, completing ignition and track change each time in an inertial pointing mode according to the installation direction of each thruster;
s5, in order to ensure that the measurement and control arc section after ignition and the subsequent track capture at a far site are ensured, the near site height of the target track is forward biased, and the rest parameters take the task track parameters as target values;
s6, carrying out low-thrust orbit transfer near the far site to reduce the height of the near site to a target value, and finally completing the orbit transfer of the satellite;
and S7, in order to avoid the influence of the single event effect of the space environment on the track change process, the ignition height is selected to be more than 28000 km.
In step S1, because the satellite-rocket separation orbit near-site height is low (less than 500 km, the orbit life is short), in order to avoid the adverse factors of the satellite such as atmospheric resistance (causing satellite touchdown), space debris (collision risk), atomic oxygen (damaging optical components and the like), and the like, the orbit needs to be changed by a plurality of times of small thrust at the far site, the satellite orbit near-site height is raised to 1000 km-1100 km, and the next stage of orbit change is implemented after the satellite safety is ensured.
In step S2, the available 10N engine specific impulse will lose 10% relative to the 490N engine due to 490N engine failure, and the direct use of the 10N thruster for immediate orbital transfer will consume more propellant, thereby affecting the satellite' S in-orbit lifetime or even failing to be in-orbit. Therefore, when the satellite orbit inclination is smaller than the critical inclination of 63.4 degrees, the positive increasing effect of the earth perturbation on the near-place amplitude angle can be utilized:
w is the near-spot amplitude angle, n is the track average angular velocity, R e Is the equatorial radius of the earth, e is the orbital eccentricity, i is the orbital tilt angle, J 2 Is the harmonic term coefficient, t is time, and a is the semi-long axis of the satellite orbit. When the amplitude perturbation of the near-place increases to an acceptable range (the consumption of the next-stage orbital transfer propellant can meet the service life of the task, the working time of the thruster is not out of limit, and the like), starting to carry out next-stage orbital transfer:
in the step S3, the factors that the consumption of the next-stage orbital transfer propellant can meet the service life of the task, the working time of the thruster is not out of limit and the like are comprehensively considered, and when the amplitude perturbation of the near place is increased to an acceptable range, next-stage orbital transfer is started. Because the continuous accumulated ignition time of the thrusters is limited, and the thrusters for performing attitude and orbit control tasks such as orbit maintenance and attitude control on orbit are determined in the satellite scheme design, the service life requirements of the thrusters must be comprehensively considered, and the accumulated working time of the thrusters in the orbit transfer stage is determined. In the next phase of the change of track, in the case where the change speed increase Δv has been obtained, the change speed increase may be determined according to the Ji Erao kofsky theorem (Δm=m (1-exp (- Δv/(I) s G)) to obtain the total weight of propellant (Δm) required for orbital transfer, M being the total weight of the satellite, exp being an exponential function based on natural logarithms, I s Is the engine specific impulse and g is the gravitational acceleration. For a 10N thruster, the propellant mass dm=f/(I) per unit time of the thruster is injected s G), F is the thrust of the thruster. The accumulated working time required by the orbit change of all thrusters of the satellite isWhen the working time of each thruster of the track-changing task is distributed, the time of the thrusters of the attitude track control task is required to be born after track entering is considered, and the working time of the thrusters is correspondingly less distributed, so that the service life requirement of the thrusters is met.
In step S4, according to the track change speed increment direction, combining the installation direction of each thruster and the use time requirement, and adopting an inertial pointing mode to finish track change. Because the speed increment entering the large elliptical freezing orbit is not along the direction of the orbit surface and is not perpendicular to the direction of the orbit surface, and the installation orientations of the thrusters on the satellites are different, the thrusters are oriented differently, and when each speed pulse is completed, the inertial pointing gesture of the satellites needs to be established to ensure that the thrust direction provided by the thrusters is along the direction of the speed increment.
In step S5, according to the requirements of measuring and controlling the duration of the arc section after ignition, the target track height forward bias is carried out, and the rest track parameters all take the task track parameters as targets and enter the track to be captured. Because the second speed pulse of the large elliptical freezing track is positioned at the track descending section, the measurement and control arc section with lower height after ignition is tense, the near-place height needs to be manually forward biased to ensure the measurement and control arc section time length after ignition; and because the near-site of the large elliptical freezing track is positioned near the south pole, the ground station of China cannot perform measurement and control, and the long-time ignition and track transfer arc section loss is large near the near-site, which is not beneficial to propellant saving, the final track capturing operation is required to be performed near the far-site, so that the measurement and control conditions before and after ignition are ensured. Both factors require a positive offset near-ground height, the magnitude of the offset being determined by the arc length requirement after ignition. In addition, in order to reduce the ignition time of the captured orbit and thus ensure the accuracy, the orbit parameters of the satellite except the near-site altitude before capturing are required to reach the task orbit requirement.
In step S6, on the basis of step S5, a plurality of times of low-thrust ignition is performed near the far site to perform low-thrust orbit transfer to reduce the near-site height to a target value, and finally the satellite orbit transfer is completed
In step S7, the large ellipse freezing orbit crosses the earth radiation band 2 times a day, if the satellite is affected by the single event effect of the space environment during ignition, the satellite register is knocked over, which will cause the calculation of the satellite software to generate an error calculation value, so that an error control command is sent to cause the error action of the attitude orbit control thruster to cause the uncontrolled overturn of the satellite attitude, and the satellite ignition height is required to be above 28000km according to the distribution characteristics of the earth radiation band, thereby avoiding the orbit transfer risk caused by the single event effect of the space environment.
The application also provides a large elliptical freezing track entering system which can be realized by executing the flow steps of the large elliptical freezing track entering method, namely, the large elliptical freezing track entering method can be understood as a preferred implementation mode of the large elliptical freezing track entering system by a person skilled in the art. Those skilled in the art will appreciate that the application provides a system and its individual devices, modules, units, etc. that can be implemented entirely by logic programming of method steps, in addition to being implemented as pure computer readable program code, in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers, etc. Therefore, the system and various devices, modules and units thereof provided by the application can be regarded as a hardware component, and the devices, modules and units for realizing various functions included in the system can also be regarded as structures in the hardware component; means, modules, and units for implementing the various functions may also be considered as either software modules for implementing the methods or structures within hardware components.
The application provides a large elliptical freezing track rail-entering system, which comprises:
the module M1 enables the satellite to orbit and raise the height of the near site for multiple times with low thrust, so as to avoid adverse factors of atmospheric resistance, space debris and atomic oxygen to the satellite;
the module M2 enables the satellite to freely run in orbit for a plurality of months, and the perturbation forward factor is utilized to naturally increase the perigee amplitude angle, so that the propellant required for changing the perigee amplitude angle is saved;
the module M3 enables the satellite to equally distribute the working time of each thruster according to the track-changing speed increment requirement, so as to ensure the service life of each thruster and the subsequent track-keeping requirement;
the module M4 enables the satellite to complete each ignition and orbit change in an inertial pointing mode according to the installation direction of each thruster;
the module M5 enables the satellite to forward bias the height of the near-site of the target orbit, and other parameters take the task orbit parameters as target values, so that the time length of the measurement and control arc section after ignition meets the requirements and is captured in the far-site orbit subsequently;
the module M6 enables the satellite to perform low-thrust orbit transfer near the far site to reduce the near site height to a target value, and finally completes the orbit transfer of the satellite;
and the module M7 is used for selecting the satellite ignition height to be more than 28000km in order to avoid the influence of the single event effect in the space environment on the orbit transfer process.
The foregoing describes specific embodiments of the present application. It is to be understood that the application is not limited to the particular embodiments described above, and that various changes or modifications may be made by those skilled in the art within the scope of the appended claims without affecting the spirit of the application. The embodiments of the application and the features of the embodiments may be combined with each other arbitrarily without conflict.
Claims (10)
1. A method for tracking a large elliptical frozen track, comprising:
s1, carrying out multiple times of low-thrust orbit transfer on a far site to raise the height of the near site, so as to avoid adverse factors of atmospheric resistance, space fragments and atomic oxygen on satellites;
s2, enabling the satellite to freely run in orbit for a plurality of months, naturally increasing the perigee amplitude angle by utilizing a perturbation forward factor, and saving the propellant required for changing the perigee amplitude angle;
step S3, equally distributing working time of each thruster according to the track-changing speed increment requirement, and guaranteeing service life of each thruster and subsequent track-keeping requirement;
s4, completing ignition and track change each time in an inertial pointing mode according to the installation direction of each thruster;
s5, forward biasing the near-site height of the target track, wherein the rest parameters take the task track parameters as target values, so that the time length of the measurement and control arc segment after ignition meets the requirements and is captured in the far-site track subsequently;
s6, carrying out low-thrust orbit transfer near the far site to reduce the height of the near site to a target value, and finally completing the orbit transfer of the satellite;
and S7, in order to avoid the influence of the single event effect of the space environment on the track change process, the ignition height is selected to be more than 28000 km.
2. The method according to claim 1, wherein in the step S1, the near-spot height of the satellite orbit is raised to 1000 km-1100 km, so that the next orbit change is performed after the satellite is safe.
3. The method according to claim 1, wherein in the step S2, when the satellite orbit incidence angle is smaller than the critical incidence angle 63.4 degrees, the positive increasing effect of earth perturbation on the perigee amplitude angle is utilized:
w is the near-spot amplitude angle, n is the track average angular velocity, R e Is the equatorial radius of the earth, e is the orbital eccentricity, i is the orbital tilt angle, J 2 Is a harmonic term coefficient, t is time, a is the semi-long axis of the satellite orbit; and when the amplitude perturbation of the near-spot increases to an acceptable range, starting to change the track of the next stage.
4. A method of orbital approach to large elliptical freezing according to claim 3 wherein the acceptable range is that the next stage of orbital transfer propellant consumption can meet the mission life and the thruster operating length is not exceeded.
5. A macrooval jelly according to claim 1The orbit joining method is characterized in that in the step S3, when the orbit is changed in the next stage, the total weight of the propellant required to be injected for orbit change is obtained under the condition that the orbit change speed increment Deltav is obtained, and the accumulated working time required for orbit change of all the thrusters of the satellite is as follows
Δm represents the total weight of propellant injected required to obtain a change of track;
dm represents the mass of propellant injected per unit time by the thruster.
6. The method according to claim 1, wherein in step S4, in combination with the installation direction and the usage time requirement of each thruster, the satellite needs to establish an inertial pointing attitude to ensure that the thrust direction provided by the thruster is along the speed increment direction, and the orbiting is completed by adopting an inertial pointing manner.
7. The method for entering a large elliptical frozen orbit according to claim 1, wherein in the step S5, the target orbit is highly positively biased according to the requirements of the measurement and control arc segment duration after ignition, and the rest orbit parameters all target the task orbit parameters and enter the orbit to be captured.
8. The method according to claim 1, wherein in step S6, the satellite orbit is finally completed by igniting with a small thrust in the vicinity of the remote site for a small thrust orbit reduction of the near-site altitude to a target value.
9. A large elliptical freeze rail in-track system, comprising:
the module M1 enables the satellite to orbit and raise the height of the near site for multiple times with low thrust, so as to avoid adverse factors of atmospheric resistance, space debris and atomic oxygen to the satellite;
the module M2 enables the satellite to freely run in orbit for a plurality of months, and the perturbation forward factor is utilized to naturally increase the perigee amplitude angle, so that the propellant required for changing the perigee amplitude angle is saved;
the module M3 enables the satellite to equally distribute the working time of each thruster according to the track-changing speed increment requirement, so as to ensure the service life of each thruster and the subsequent track-keeping requirement;
the module M4 enables the satellite to complete each ignition and orbit change in an inertial pointing mode according to the installation direction of each thruster;
the module M5 enables the satellite to forward bias the height of the near-site of the target orbit, and other parameters take the task orbit parameters as target values, so that the time length of the measurement and control arc section after ignition meets the requirements and is captured in the far-site orbit subsequently;
the module M6 enables the satellite to perform low-thrust orbit transfer near the far site to reduce the near site height to a target value, and finally completes the orbit transfer of the satellite;
and the module M7 is used for selecting the satellite ignition height to be more than 28000km in order to avoid the influence of the single event effect in the space environment on the orbit transfer process.
10. The large elliptical freeze orbit in-orbit system of claim 9, wherein the satellites are controlled using the large elliptical freeze orbit in-orbit method of any one of claims 1 to 8.
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| CN202311134429.5A CN117125271A (en) | 2023-09-04 | 2023-09-04 | Method and system for entering large elliptic frozen orbit |
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| CN202311134429.5A CN117125271A (en) | 2023-09-04 | 2023-09-04 | Method and system for entering large elliptic frozen orbit |
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