CN109211225A - Obtain method, system and the equipment of highly elliptic orbit space object remaining orbital lifetime - Google Patents
Obtain method, system and the equipment of highly elliptic orbit space object remaining orbital lifetime Download PDFInfo
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Abstract
The present invention provides method, system and the equipment of a kind of acquisition highly elliptic orbit space object remaining orbital lifetime, and method includes: the preliminary orbit radical of track where obtaining initial time space object；If preliminary orbit radical meets highly elliptic orbit orbital tracking condition；The lunisolar gravitational perturbation change rate for the orbital tracking that the J2 perturbation change rate and current time lunisolar attraction for obtaining the orbital tracking of the aspherical gravitation influence of atmospheric drag perturbation change rate, the current time earth for the orbital tracking that current time atmospheric drag influences influence；And then obtain then perigee altitude that current orbit radical provides space object；When the perigee altitude is less than or equal to default perigee altitude or the time of integration is greater than the default time of integration, the space object remaining orbital lifetime is obtained.It is more acurrate that method provided by the present invention obtains highly elliptic orbit space object remaining orbital lifetime.
Description
Technical field
The present invention relates to spacecraft orbit dynamics field more particularly to a kind of acquisition highly elliptic orbit space object are remaining
Method, system and the equipment of orbital lifetime.
Background technique
Orbital lifetime refers to the time that space object retains in orbit.It is from space object injection to pass away for
Time only.For the in-orbit space object of part highly elliptic orbit mainly by the effect of lunisolar attraction, eccentricity can have long period
Variation can be gradually decreased so as to cause perigee altitude when the process that eccentricity becomes larger, finally space object be fallen into
Dense atmosphere (from the ground 100 kilometers) is burnt, wherein larger-size space object may can not be complete in dense atmosphere
Clean burn is ruined, to drop on ground, is caused damages to the personnel and property safety on ground.
For highly elliptic orbit because of the particularity of its track, many countries are all using highly elliptic orbit as in-orbit space
The designed path of object.The mechanism that passes away of highly elliptic orbit space object is complex, utilizes original consideration atmospheric drag shadow
The method of loud calculating orbital lifetime is difficult to its accurate remaining orbital lifetime of accurate forecast.
Summary of the invention
Based on this, it is necessary in view of the above technical problems, it is surplus to provide a kind of simple acquisition highly elliptic orbit space object
Method, system and the equipment of remaining orbital lifetime, the method comprise the steps that
The preliminary orbit radical of track where obtaining initial time space object；
If the preliminary orbit radical meets highly elliptic orbit orbital tracking condition；
Atmospheric drag perturbation change rate, the current time earth for obtaining the orbital tracking that current time atmospheric drag influences are non-
The life for the orbital tracking that the J2 perturbation change rate and current time lunisolar attraction for the orbital tracking that spherical gravitation influences influence is drawn
Power perturbation change rate；
The atmospheric drag perturbation change rate, J2 perturbation change rate and lunisolar gravitational perturbation change rate are overlapped, then
It is integrated by numerical integration algorithm with predetermined time step delta t, obtains the orbital tracking under the n-th Δ t time of integration；
According to the orbital tracking under the n-th Δ t time of integration, semi-major axis of orbit and orbital eccentricity are obtained, is calculated
Product between the semi-major axis of orbit and orbital eccentricity and the difference between the determining semi-major axis of orbit and the product,
And obtaining the perigee altitude of space object is the difference of the difference and earth radius；
When the perigee altitude is less than or equal to default perigee altitude or the N Δ t total mark time is greater than default integral
Between when, obtain the space object remaining orbital lifetime.
The orbital tracking atmospheric drag perturbation for obtaining current time atmospheric drag and influencing in one of the embodiments,
Change rate, comprising:
Numerical integration is carried out by variable of the very close angle of track where the space object, obtains the current time atmosphere
The orbital tracking atmospheric drag perturbation change rate of drag effects for the data integrate after value and the space object it is in-orbit
The ratio of the orbital period in road；
Wherein, further includes: according to preset Atmospheric models, obtain corresponding with the very close angle after the variable change
Solar activity parameter F10.7, geomagnetic index AP, orbit altitude and atmospheric density.
The numerical integration algorithm is Runge Kutta quadravalence or five rank integral algorithms in one of the embodiments,.
It is described when the perigee altitude is less than or equal to default perigee altitude or described in one of the embodiments,
When the time of integration is greater than the default time of integration, the space object remaining orbital lifetime is obtained, comprising:
When the perigee altitude is less than or equal to default perigee altitude, the space object remaining orbital lifetime is N
Δt；
When the time of integration being greater than the default time of integration, the space object remaining orbital lifetime is the default product
Between timesharing.
In one of the embodiments, the method also includes:
When the perigee altitude is greater than default perigee altitude and the time of integration is less than the default time of integration, return
Return the atmospheric drag perturbation change rate for obtaining the orbital tracking that current time atmospheric drag influences, current time earth aspheric
The lunisolar attraction for the orbital tracking that the J2 perturbation change rate and current time lunisolar attraction for the orbital tracking that shape gravitation influences influence
The step of perturbation change rate.
The present invention also provides the systems of acquisition highly elliptic orbit space object remaining orbital lifetime a kind of, comprising:
Orbital tracking obtains module, the preliminary orbit radical for track where obtaining initial time space object；
Judgment module, for judging that the preliminary orbit radical meets highly elliptic orbit orbital tracking condition；
Orbital tracking change rate obtains module, comprising:
Atmospheric drag influences the orbital tracking atmospheric drag perturbation change that module is used to obtain the influence of current time atmospheric drag
Rate, the aspherical gravitation of the earth influence module and are used to obtain the orbital tracking J2 perturbation that the aspherical gravitation of the current time earth influences
Change rate, lunisolar attraction influence module and are used for the orbital tracking lunisolar gravitational perturbation change rate that current time lunisolar attraction influences；
Integration module, for changing the atmospheric drag perturbation change rate, J2 perturbation change rate and lunisolar gravitational perturbation
Rate is overlapped, then is integrated by numerical integration algorithm with predetermined time step delta t, is obtained under the n-th Δ t time of integration
Orbital tracking；
Perigee altitude obtains module, for obtaining track according to the orbital tracking under the n-th Δ t time of integration
Semi-major axis and orbital eccentricity and the difference for calculating the semi-major axis of orbit Yu the semi-major axis of orbit and orbital eccentricity product,
And then the difference of the difference and earth radius is calculated, the final perigee altitude for obtaining space object；
Remaining orbital lifetime obtains module, for being less than or equal to default perigee altitude or institute when the perigee altitude
When stating the time of integration greater than the default time of integration, the space object remaining orbital lifetime is obtained.
The atmospheric drag influence module includes: in one of the embodiments,
Atmospheric drag influences change rate unit, for being counted using the very close angle of track where the space object as variable
Value integral obtains the orbital tracking atmospheric drag perturbation change rate that the current time atmospheric drag influences as data integral
The ratio of the orbital period of value and space object place track afterwards；
The atmospheric drag influences change rate unit further include: the very close corresponding parameters unit in angle of track, for according to pre-
If Atmospheric models, obtain with the corresponding solar activity parameter F10.7 in very close angle after the variable change, geomagnetic index
AP, orbit altitude and atmospheric density.
The remaining orbital lifetime obtains module in one of the embodiments, is also used to when the perigee altitude is big
In default perigee altitude and when the time of integration is less than the default time of integration, the acquisition current time atmospheric drag is returned
The J2 for the orbital tracking that the aspherical gravitation of atmospheric drag perturbation change rate, the current time earth of the orbital tracking of influence influences takes the photograph
The lunisolar gravitational perturbation change rate for the orbital tracking that dynamic change rate and current time lunisolar attraction influence.
It includes height judging unit and time judgement that the remaining orbital lifetime, which obtains module, in one of the embodiments,
Unit；
The height judging unit, for the space described when the perigee altitude is less than or equal to default perigee altitude
Object residue orbital lifetime is N Δ t；
The time judging unit, it is remaining for the space object described when the time of integration being greater than the default time of integration
Orbital lifetime is greater than the default time of integration.
The present invention also provides the equipment of acquisition highly elliptic orbit space object remaining orbital lifetime a kind of, including processor,
The computer program of memory and storage on a memory, the computer program by the processor when being handled in realization
The step of stating any one the method.
The beneficial effects of the present invention are:
The aspherical gravitation of the earth influences under the conditions of highly elliptic orbit and lunisolar attraction influences to move space object track
It is critically important, therefore the method provided by the present invention for obtaining highly elliptic orbit space object remaining orbital lifetime is in addition to considering atmosphere
Dynamics of orbits under drag effects also increases and considers that the aspherical gravitation of the earth influences and the track power under the influence of lunisolar attraction
Amendment is learned, so that it is more acurrate to obtain highly elliptic orbit space object remaining orbital lifetime.
Detailed description of the invention
Fig. 1 is the flow chart of the method for obtaining highly elliptic orbit space object remaining orbital lifetime of one embodiment；
Fig. 2 is the flow chart of the method for obtaining highly elliptic orbit space object remaining orbital lifetime of another embodiment；
Fig. 3 is the structure chart of the system for obtaining highly elliptic orbit space object remaining orbital lifetime of one embodiment；
Fig. 4 is the structure chart of the system for obtaining highly elliptic orbit space object remaining orbital lifetime of another embodiment.
Specific embodiment
In order to make the objectives, technical solutions, and advantages of the present invention clearer, right with reference to the accompanying drawings and embodiments
The present invention is further elaborated.It should be appreciated that described herein, specific examples are only used to explain the present invention, not
For limiting the present invention.
Fig. 1 is the flow chart of the method for obtaining highly elliptic orbit space object remaining orbital lifetime of one embodiment, such as
The method shown in FIG. 1 for obtaining highly elliptic orbit space object remaining orbital lifetime includes:
Step S100, the preliminary orbit radical of track where obtaining initial time space object.
Specifically, the orbital tracking be semi-major axis of orbit a, orbital eccentricity e, orbit inclination angle i, right ascension of ascending node Ω,
Argument of perigee w and mean anomaly M.Orbital tracking or orbital elements or orbit parameter are for describing celestial body in its track
One group of parameter of operating status.It is often the case that describing space object with the classical law of universal gravitation is moved when institute by conic section
Required 6 parameters.The space object chosen in the present embodiment be China detect No. 1 satellite, international numbering 28140, from
The space environment Parameter File space-weather.txt of nineteen fifty-seven so far is downloaded on this website www.celestrak.com,
This environmental parameter file middle orbit radical is recorded as two row orbital trackings.It is also possible to other modes for recording orbital tracking.Two
Row orbital tracking can provide semi-major axis of orbit a, the track of track initial time where specific China detects No. 1 satellite by conversion
Eccentric ratio e, orbit inclination angle i, right ascension of ascending node Ω, argument of perigee w and mean anomaly M are as shown in table 1:
Table 1
Step S200, if the preliminary orbit radical meets highly elliptic orbit orbital tracking condition.
Specifically, for needing to study object in certain tracks, certain tracks have highly elliptic orbit orbital tracking.It is described
The preliminary orbit radical of track needs to meet highly elliptic orbit orbital tracking condition where space object, if it is next to meet continuation
Step obtains work.If be unsatisfactory for highly elliptic orbit orbital tracking condition, selected space object is not on planned orbit,
It calculates and terminates.
In one embodiment, certain tracks are highly elliptic orbit, and the predetermined trajectory semi-major axis is 15000km, described pre-
If orbital eccentricity is 0.3.As shown in table 1, the orbital tracking of track where specific China detects No. 1 satellite is specially track half
Long axis is greater than 15000km, and orbital eccentricity meets highly elliptic orbit condition also greater than 0.3.
Step S300, the atmospheric drag perturbation change rate S310 for the orbital tracking that acquisition current time atmospheric drag influences,
The J2 perturbation change rate S320 and current time lunisolar attraction for the orbital tracking that the aspherical gravitation of the current time earth influences influence
Orbital tracking lunisolar gravitational perturbation change rate S330.
Specifically, the equation of the atmospheric drag perturbation change rate S310 for the orbital tracking that current time atmospheric drag influences
Are as follows:
Wherein, σ=CDA/2 is ballistic coefficient, CDFor the resistance coefficient of space object, it is the height that 2.2, ρ is taken in the present embodiment
The atmospheric density of degree, v are the rate of space object, and r is the radial distance of space object, and a is the semi-major axis of orbit of space object,
E is the orbital eccentricity of space object, and μ is Gravitational coefficient of the Earth, and A is the area quality ratio of space object, and θ is space object
The true anomaly of track.Giving above-mentioned coefficient under specific Atmospheric models can be directly obtained, and carry out by variable of the very close angle
Numerical integration, the atmospheric drag perturbation change rate S310 for the orbital tracking that current time atmospheric drag influences.
Specifically, the perturbation side of the J2 perturbation change rate S320 for the orbital tracking that the aspherical gravitation of the current time earth influences
Journey indicates are as follows:
Wherein, Re is earth mean orbit radius, and n is the mean angular velocity of space object,Draw for the earth is aspherical
The right ascension of ascending node for the space object that power influences,For the argument of perigee for the space object that the aspherical gravitation of the earth influences, i
Orbit inclination angle, half promoting menstruation of p track.According to preset earth model, JGM3 standard earth model is used in the present embodiment, acquisition is taken the photograph
Dynamic equation weight relevant parameter and J2=0.0010826269.
Specifically, the lunisolar gravitational perturbation change rate S330 for the orbital tracking that current time lunisolar attraction influences is indicated are as follows:
Wherein, AL、BL、CLRespectively indicate the direction cosines between third body (sun or the moon) and space object, L=1 table
Show that moon relevant parameter, L=2 indicate sun relevant parameter, rLFor the distance between third body and the earth, μLFor third body gravitation
Constant, n are the mean angular velocity of space object.
In the present embodiment, the orbital tracking such as table 2 of the initial time of the sun and the moon in the inertial coodinate system of the earth's core equator
It is shown:
Table 2
| a(km) | e | i(deg) | Ω(deg) | w(deg) | M(deg) | |
| The moon | 384747.981 | 0.054879905 | 28.4987 | 358.0646 | 10.0047 | 68.1390 |
| The sun | 149597885.949 | 0.01670567 | 23.4387 | 0.0 | 282.9599 | 358.1607 |
Step S400, by described atmospheric drag perturbation change rate S310, J2 perturbation change rate S320 and lunisolar gravitational perturbation
Change rate S330 is overlapped, then is integrated by numerical integration algorithm with predetermined time step delta t, and n-th Δ t product is obtained
Orbital tracking under between timesharing.
Specifically, superposition atmospheric drag perturbation change rate S310, J2 perturbation change rate S320 and lunisolar gravitational perturbation variation
It after rate S330, provides orbital tracking and changes with time rate, indicate are as follows:
Then it is integrated again by numerical integration algorithm with predetermined time step delta t, obtains the n-th Δ t time of integration
Under orbital tracking.
Step S500 obtains semi-major axis of orbit and track is inclined according to the orbital tracking under the n-th Δ t time of integration
Heart rate, calculate the product between the semi-major axis of orbit and orbital eccentricity and determine the semi-major axis of orbit and the product it
Between difference, and obtain space object perigee altitude be the difference and earth radius difference.
Specifically, according to the orbital tracking obtained in step S400, semi-major axis of orbit and orbital eccentricity are therefrom extracted, is led to
Cross formula rp=a* (1-e)-ReThe perigee altitude of the space object is obtained, wherein ReFor earth radius.
Step S600, when the perigee altitude is greater than less than or equal to default perigee altitude or N Δ t total mark time
When the default time of integration, the space object remaining orbital lifetime is obtained.
Specifically, judge to judge also by the time of integration by perigee altitude, when the perigee altitude is less than
Being greater than default time of integration satisfaction equal to default perigee altitude and the time of integration wherein can be obtained the space for one
Object residue orbital lifetime.
Fig. 2 is the flow chart of the method for obtaining highly elliptic orbit space object remaining orbital lifetime of another embodiment,
If Fig. 2 shows, the atmospheric drag perturbation change rate S310 for obtaining the orbital tracking that current time atmospheric drag influences includes:
Step S311 carries out numerical integration by variable of the very close angle of track where the space object, works as described in acquisition
The orbital tracking atmospheric drag perturbation change rate that preceding moment atmospheric drag influences is the value and the space after the data integrate
The ratio of the orbital period of track where object.
Specifically, the very close angle is divided into several equal parts, the very close angle is divided into 360 equal parts in one embodiment.
After being overlapped with each equal part, divided by the orbital period, to obtain the change rate and orbital eccentricity of semi-major axis of orbit
Atmospheric drag perturbation change rate.According to NRLMSISE2000 Standard Earth (SE) Atmospheric models, the current time atmospheric drag that provides
The orbital tracking atmospheric drag perturbation change rate of influence.
During calculating step S311, further includes: step S312 is obtained and according to preset Atmospheric models with the change
The corresponding solar activity parameter F10.7 in very close angle, geomagnetic index AP, orbit altitude and atmospheric density after amount variation.
Specifically, preset Atmospheric models are NRLMSISE2000 Standard Earth (SE) Atmospheric models in one embodiment.According to
NRLMSISE2000 Standard Earth (SE) Atmospheric models obtain angle phase very close with space object place track under different very close angles
Corresponding sun extreme ultraviolet effect parameter, that is, solar activity parameter F10.7, geomagnetic index AP, orbit altitude and then basis
NRLMSISE2000 Standard Earth (SE) Atmospheric models provide atmospheric density.
In one embodiment, the numerical integration algorithm is Runge Kutta quadravalence or five rank integral algorithms.
Described to be integrated by numerical integration algorithm with predetermined time step delta t, the numerical integration algorithm is dragon
Ge Kuta quadravalence or five order algorithms.Other suitable predetermined times are chosen not grow, should the integral calculation time it is shorter, so as to shorten
The entire acquisition time for obtaining remaining orbital lifetime, does not again influence the accuracy calculated.In one embodiment, described pre-
Step-length of fixing time is Δ t=86400 seconds.
In one embodiment, step S600, it is described to work as the perigee altitude less than or equal to default perigee altitude, or
When the time of integration is greater than the default time of integration, the space object remaining orbital lifetime is obtained, comprising: step S610, when
When the perigee altitude is less than or equal to default perigee altitude, the space object remaining orbital lifetime is N Δ t；Step
S620, when the time of integration being greater than the default time of integration, the space object remaining orbital lifetime is the default integral
Time.
Specifically, perigee altitude and the time of integration are judged respectively.When default perigee altitude is 100km, when
rpWhen≤100km, the corresponding time is passing away the time for space object at this time, and remaining life is N Δ t.The default time of integration
It is 100 years, if the time of integration is greater than 100 years, calculates termination, the remaining orbital lifetime for exporting space object is greater than 100 years.
This default perigee altitude and the default time of integration set different numerical value for different planned orbits.
In one embodiment, when the perigee altitude is greater than default perigee altitude and the time of integration is less than in advance
If when the time of integration, N+1 is assigned to N, the orbital tracking that acquisition current time atmospheric drag influences described in return step S300
The J2 perturbation change rate for the orbital tracking that the aspherical gravitation of atmospheric drag perturbation change rate, the current time earth influences and it is current when
The lunisolar gravitational perturbation change rate for carving the orbital tracking that lunisolar attraction influences, until the perigee altitude is less than or equal to preset closely
When location higher or the time of integration are greater than the default time of integration, the space object remaining orbital lifetime is obtained.
Specifically, in step S600, in most cases predetermined time step-length once had better not just make the perigee
Height is less than or equal to default perigee altitude or the N Δ t time of integration is greater than the default time of integration.When the perigee is high
N+1 is assigned to N, return step 300 greater than default perigee altitude and when the time of integration is less than the default time of integration by degree
Obtain orbital tracking atmospheric drag perturbation change rate, the aspherical gravitation of the current time earth that the atmospheric drag of N time Δt influences
The lunisolar gravitational perturbation for the orbital tracking that the J2 perturbation change rate and current time lunisolar attraction of the orbital tracking of influence influence becomes
Rate, until the perigee altitude is less than or equal to default perigee altitude or the time of integration greater than the default time of integration
When, obtain the space object remaining orbital lifetime.
Atmospheric drag shadow is considered using JGM3 standard earth model according to NRLMSISE2000 Standard Earth (SE) Atmospheric models
It rings, the aspherical gravitation influence of the earth, lunisolar attraction influence, choosing 17:00 on January 1 in 2007 is initial time, is detected to China
The motion dynamics equations of track carry out Runge Kutta quadravalence integral algorithm where No. 1 satellite, be integrated to operation provide it is current
The semi-major axis of orbit and orbital eccentricity at moment obtain perigee altitude rp≤ 100km is calculated and is terminated, and exports the remaining track longevity
Life is 284 days.And the real surplus orbital lifetime observed is 286 days, the acquisition methods of the present embodiment are more accurate.
The present invention also provides the systems of acquisition highly elliptic orbit space object remaining orbital lifetime a kind of.
Fig. 3 is the structure chart of the system for obtaining highly elliptic orbit space object remaining orbital lifetime of one embodiment, Fig. 3
Shown in the system of acquisition highly elliptic orbit space object remaining orbital lifetime include:
Preliminary orbit radical obtains module 100, the preliminary orbit root for track where obtaining initial time space object
Number；
Judgment module 200, for judging that the preliminary orbit radical meets highly elliptic orbit orbital tracking condition；
Orbital tracking change rate obtains module 300, comprising: it is big for obtaining current time that atmospheric drag influences module 310
The orbital tracking atmospheric drag perturbation change rate that atmidometer influences, it is current for obtaining that the aspherical gravitation of the earth influences module 320
The orbital tracking J2 perturbation change rate that the aspherical gravitation of the moment earth influences, lunisolar attraction influence module 330 and are used for current time
The orbital tracking lunisolar gravitational perturbation change rate that lunisolar attraction influences；
Integration module 400, for becoming the atmospheric drag perturbation change rate, J2 perturbation change rate and lunisolar gravitational perturbation
Rate is overlapped, then is integrated by numerical integration algorithm with predetermined time step delta t, and the n-th Δ t time of integration is obtained
The orbital tracking of track where down space object；
Perigee altitude obtains module 500, for obtaining rail according to the orbital tracking under the n-th Δ t time of integration
Road semi-major axis and orbital eccentricity and the difference for calculating the semi-major axis of orbit Yu the semi-major axis of orbit and orbital eccentricity product
Value, and then the difference of the difference and earth radius is calculated, the final perigee altitude for obtaining space object；
Remaining orbital lifetime obtains module 600, is less than or equal to default perigee altitude for working as the perigee altitude, or
When the time of integration is greater than the default time of integration, the space object remaining orbital lifetime is obtained.
Fig. 4 is the structure chart of the system for obtaining highly elliptic orbit space object remaining orbital lifetime of another embodiment.
Shown in Fig. 4, one of embodiment, the atmospheric drag influences module 310 and includes:
Atmospheric drag influences change rate unit 311, for using the very close angle of track where the space object as variable into
Row numerical integration, obtaining the orbital tracking atmospheric drag perturbation change rate that the current time atmospheric drag influences is the data
The ratio of the orbital period of track where value and the space object after integral；
The atmospheric drag influences change rate unit 311 further include: the corresponding parameters unit 312 in the very close angle of track is used for
According to preset Atmospheric models, obtain with the corresponding solar activity parameter F10.7 in very close angle after the variable change,
Magnetic Index A P, orbit altitude and atmospheric density.
In one embodiment, the numerical integration algorithm used in the integration module 400 is Runge Kutta four
Rank.The numerical integration algorithm can also use five rank integral algorithm of Runge Kutta.
One of embodiment, remaining orbital lifetime obtain module 600, are also used to preset when the perigee altitude is greater than
When perigee altitude and the time of integration are less than the default time of integration, N+1 is assigned to N, returns to the acquisition current time
The track root that the aspherical gravitation of atmospheric drag perturbation change rate, the current time earth for the orbital tracking that atmospheric drag influences influences
The lunisolar gravitational perturbation change rate for the orbital tracking that several J2 perturbation change rates and current time lunisolar attraction influence, until described
When perigee altitude is less than or equal to default perigee altitude or the time of integration greater than the default time of integration, the sky is obtained
Between object remaining orbital lifetime.
One of embodiment, it includes sentencing height judging unit 610 and time that the residue orbital lifetime, which obtains module 600,
Disconnected unit 620；
The height judging unit 610, described in when the perigee altitude is less than or equal to default perigee altitude
Space object residue orbital lifetime is N Δ t；
The time judging unit 620, for the space object described when the time of integration being greater than the default time of integration
Remaining orbital lifetime is greater than the default time of integration.
The present invention also provides the equipment of acquisition highly elliptic orbit space object remaining orbital lifetime a kind of.Including processor,
The computer program of memory and storage on a memory, computer program realize above-mentioned when being handled by the processor
It anticipates an embodiment the method.
The aspherical gravitation of the earth influences under the conditions of certain planned orbit and lunisolar attraction is influenced to space object track power
Learn that amendment is critically important, the method provided by the present invention for obtaining highly elliptic orbit space object remaining orbital lifetime, system and
Equipment also increases in addition to considering that atmospheric drag influences and considers that the aspherical gravitation of the earth influences and lunisolar attraction influences, so that obtaining
Highly elliptic orbit space object remaining orbital lifetime is more acurrate.
Only several embodiments of the present invention are expressed for above embodiments, and the description thereof is more specific and detailed, but can not
Therefore it is construed as limiting the scope of the patent.It should be pointed out that for those of ordinary skill in the art,
Under the premise of not departing from present inventive concept, various modifications and improvements can be made, and these are all within the scope of protection of the present invention.
Therefore, the scope of protection of the patent of the invention shall be subject to the appended claims.
Claims (10)
1. a kind of method for obtaining highly elliptic orbit space object remaining orbital lifetime, which is characterized in that the described method includes:
The preliminary orbit radical of track where obtaining initial time space object；
If the preliminary orbit radical meets highly elliptic orbit orbital tracking condition；
Atmospheric drag perturbation change rate, the current time earth for obtaining the orbital tracking that current time atmospheric drag influences are aspherical
The lunisolar attraction for the orbital tracking that the J2 perturbation change rate and current time lunisolar attraction for the orbital tracking that gravitation influences influence is taken the photograph
Dynamic change rate；
The atmospheric drag perturbation change rate, J2 perturbation change rate and lunisolar gravitational perturbation change rate are overlapped, then passed through
Numerical integration algorithm is integrated with predetermined time step delta t, obtains the orbital tracking under the n-th Δ t time of integration；
According to the orbital tracking under the n-th Δ t time of integration, semi-major axis of orbit and orbital eccentricity are obtained, described in calculating
Product between semi-major axis of orbit and orbital eccentricity and the difference between the determining semi-major axis of orbit and the product, and
The perigee altitude for obtaining space object is the difference of the difference and earth radius；
When the perigee altitude is less than or equal to default perigee altitude or N Δ t total mark time greater than the default time of integration
When, obtain the space object remaining orbital lifetime.
2. obtaining the method for highly elliptic orbit space object remaining orbital lifetime as described in claim 1, which is characterized in that institute
It states and obtains the orbital tracking atmospheric drag perturbation change rate that current time atmospheric drag influences, comprising:
Numerical integration is carried out by variable of the very close angle of track where the space object, obtains the current time atmospheric drag
The orbital tracking average rate of change of influence；
Wherein, further includes: according to preset Atmospheric models, obtain with the corresponding sun in very close angle after the variable change
Movement parameter F10.7, geomagnetic index AP, orbit altitude and atmospheric density.
3. obtaining the method for highly elliptic orbit space object remaining orbital lifetime as described in claim 1, which is characterized in that institute
Stating numerical integration algorithm is Runge Kutta quadravalence or five rank integral algorithms.
4. obtaining the method for highly elliptic orbit space object remaining orbital lifetime as described in claim 1, which is characterized in that institute
It states when the perigee altitude is less than or equal to default perigee altitude or the time of integration is greater than the default time of integration, obtains
Obtain the space object remaining orbital lifetime, comprising:
When the perigee altitude is less than or equal to default perigee altitude, the space object remaining orbital lifetime is N Δ t；
When the time of integration being greater than the default time of integration, when the space object remaining orbital lifetime is the default integral
Between.
5. obtaining the method for highly elliptic orbit space object remaining orbital lifetime as described in claim 1, which is characterized in that institute
State method further include:
When the perigee altitude is greater than default perigee altitude and the time of integration is less than the default time of integration, by N+1
Be assigned to N, return the atmospheric drag perturbation change rate for obtaining the orbital tracking that current time atmospheric drag influences, it is current when
The orbital tracking that the J2 perturbation change rate and current time lunisolar attraction for carving the orbital tracking that the aspherical gravitation of the earth influences influence
Lunisolar gravitational perturbation change rate the step of.
6. a kind of system for obtaining highly elliptic orbit space object remaining orbital lifetime characterized by comprising
Orbital tracking obtains module, the preliminary orbit radical for track where obtaining initial time space object；
Judgment module, for judging that the preliminary orbit radical meets highly elliptic orbit orbital tracking condition；
Orbital tracking change rate obtains module, comprising:
Atmospheric drag influences the orbital tracking atmospheric drag perturbation change rate that module is used to obtain the influence of current time atmospheric drag,
The aspherical gravitation of the earth influences module and is used to obtain the orbital tracking J2 perturbation variation that the aspherical gravitation of the current time earth influences
Rate, lunisolar attraction influence module and are used for the orbital tracking lunisolar gravitational perturbation change rate that current time lunisolar attraction influences；
Integration module, for by the atmospheric drag perturbation change rate, J2 perturbation change rate and lunisolar gravitational perturbation change rate into
Row superposition, then integrated by numerical integration algorithm with predetermined time step delta t, obtain the rail under the n-th Δ t time of integration
Road radical；
Perigee altitude obtains module, for it is long to obtain track half according to the orbital tracking under the n-th Δ t time of integration
Axis and orbital eccentricity and the difference for calculating the semi-major axis of orbit Yu the semi-major axis of orbit and orbital eccentricity product, in turn
The difference of the difference and earth radius is calculated, the final perigee altitude for obtaining space object；
Remaining orbital lifetime obtains module, for being less than or equal to default perigee altitude or the product when the perigee altitude
When being greater than the default time of integration between timesharing, the space object remaining orbital lifetime is obtained.
7. obtaining the system of highly elliptic orbit space object remaining orbital lifetime as claimed in claim 6, which is characterized in that
The atmospheric drag influences module
Atmospheric drag influences change rate unit, for carrying out numerical value product by variable of the very close angle of track where the space object
Point, obtaining the orbital tracking atmospheric drag perturbation change rate that the current time atmospheric drag influences is after the data integrate
The ratio of value and the orbital period of track where the space object；
The atmospheric drag influences change rate unit further include: the very close corresponding parameters unit in angle of track, for according to preset
Atmospheric models, obtain with the corresponding solar activity parameter F10.7 in very close angle after the variable change, geomagnetic index AP,
Orbit altitude and atmospheric density.
8. obtaining the system of highly elliptic orbit space object remaining orbital lifetime as claimed in claim 6, which is characterized in that institute
Remaining orbital lifetime acquisition module is stated, is also used to be greater than default perigee altitude and the time of integration when the perigee altitude
When less than the default time of integration, N+1 is assigned to N, returns to the orbital tracking for obtaining the influence of current time atmospheric drag
The J2 perturbation change rate for the orbital tracking that the aspherical gravitation of atmospheric drag perturbation change rate, the current time earth influences and it is current when
Carve the lunisolar gravitational perturbation change rate for the orbital tracking that lunisolar attraction influences.
9. obtaining the system of highly elliptic orbit space object remaining orbital lifetime as claimed in claim 6, which is characterized in that institute
Stating remaining orbital lifetime acquisition module includes height judging unit and time judging unit；
The height judging unit, for the space object described when the perigee altitude is less than or equal to default perigee altitude
Remaining orbital lifetime is N Δ t；
The time judging unit, for the space object residue track described when the time of integration being greater than the default time of integration
Service life is greater than the default time of integration.
10. a kind of equipment for obtaining highly elliptic orbit space object remaining orbital lifetime, including processor, memory and storage
Computer program on a memory, which is characterized in that the computer program realizes right when being handled by the processor
It is required that the step of 1-5 any one the method.
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