CN106777580A - Near-earth inclined plane launch window Fast design method - Google Patents
Near-earth inclined plane launch window Fast design method Download PDFInfo
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- CN106777580A CN106777580A CN201611086251.1A CN201611086251A CN106777580A CN 106777580 A CN106777580 A CN 106777580A CN 201611086251 A CN201611086251 A CN 201611086251A CN 106777580 A CN106777580 A CN 106777580A
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Abstract
The invention discloses a kind of near-earth inclined plane launch window Fast design method, it comprises the following steps：Step one, sets design conditions, including calculate time period, networking constraint, illumination angle and its change direction constraint, satellier injection trajectory；Step 2, at the appointed time the front and rear of all windows for meeting networking constraint is given one-time calculation along interval in Julian date form in section, is called for follow-up link；Step 3, at the appointed time all the front and rear of all windows for meeting illumination condition constraint are given one-time calculation along interval in Julian date form in section, are called for follow-up link.The present invention can the constraint of combing typical emission window, input, output is clear and definite, it is independent each constraint is calculated after collect, the secondary development for being easy to increase and decrease to constrain is easy to satellite master-plan personnel quickly to design launch window, and ensure design accuracy.
Description
Technical field
The present invention relates to a kind of method for designing, the quick side of design of more particularly to a kind of near-earth inclined plane launch window
Method.
Background technology
Near-earth inclines (non-solar synchronization) orbiter has the various features such as travel through, freeze, low latitudes is covered, and is applicable
Plurality of application scenes, it is so more and more that to be used by design of satellites person.It is special with the daily launch window identical of sun-synchronous orbit
Point is different, and the daily launch window of non-solar geo-stationary orbit is differed.For the calculating of window, the traversal for generally using at present
Computational methods judge whether each corresponding orbital characteristics of time infinitesimal meets design constraint successively, arrange all eligible
Infinitesimal collection be collectively referred to as launch window.This kind of method with simple and reliable but computationally intensive, if carrying out the window meter of a year and a day
Calculate, it is calculated, and the time is long, and precision also receives the influence of infinitesimal size.
With the development of space technology, the shortening of satellite lead time, the increase of development task, due to traditional design method
The problem that desin speed is slow, design accuracy is not enough, can not meet design of satellites demand at this stage.For this reason, it is necessary to develop one
The quick algorithm for being able to ensure that computational accuracy is planted, overall tasks designer is made in the situation without be familiar with track professional knowledge
Under, remain to carry out launch window the design of fast accurate, improve development efficiency.
The content of the invention
The technical problems to be solved by the invention are to provide a kind of near-earth inclined plane launch window Fast design method, its
Can the constraint of combing typical emission window, input, output is clear and definite, it is independent each constraint is calculated after collect, be easy to increase
Subtract the secondary development of constraint, be easy to satellite master-plan personnel quickly to design launch window, and ensure design accuracy.
The present invention is to solve above-mentioned technical problem by following technical proposals：A kind of near-earth inclined plane launch window
Fast design method, it comprises the following steps：
Step one, sets design conditions, including calculate time period, networking constraint, illumination angle and its change direction bar
Part constraint, satellier injection trajectory；
Step 2, one-time calculation at the appointed time in section all windows for meeting networking constraint front and rear edge
Interval, is given in Julian date form, is called for follow-up link；
Step 3, one-time calculation at the appointed time in section all windows for meeting illumination condition constraint front and rear edges
Interval, is given in Julian date form, is called for follow-up link；
Step 4, by date along interval before and after merger, exports near-earth inclined plane launch window design.
Preferably, in the step 2 and step 3 networking constraint and illumination condition constrains unification extremely
TrueOfDate coordinate systems are recorded.
Preferably, the networking constraint in the step 2 is using the interval as designing about of two star right ascension of ascending node difference
Beam, disposably provides edge before and after all satisfactions in the specified time period constrain the corresponding launch window of interval bound and gathers.
Preferably, the networking constraint in the step 2, its constraint process has high-precision orbital recursion ability,
In low rail constellation task, the window of follow-up star can be directly participated in as input using the parameter of satellite in orbit under any epoch
Design.
Preferably, the illumination condition constraint in the step 3 directly uses orbital plane illumination angle as input.
Preferably, the illumination condition constraint only passes through to constrain edge before and after interval bound calculation window, and by algorithm
Judgement belong to forward position or after prolong, sorted out by logic.
Positive effect of the invention is：The present invention can combing typical emission window constraint, be input into, export bright
Really, collect after independently calculating each constraint, satellite master-plan personnel are easy in the secondary development for being easy to increase and decrease to constrain
Launch window is quickly designed, and is ensured design accuracy.
Brief description of the drawings
Fig. 1 is FB(flow block) of the invention.
Fig. 2 is the window calculation block diagram of illumination constraint of the invention.
Fig. 3 is schematic diagram of the orbital tracking to the transformational relation of position and speed.
Specific embodiment
Present pre-ferred embodiments are given below in conjunction with the accompanying drawings, to describe technical scheme in detail.
As shown in Figure 1 to Figure 3, near-earth inclined plane launch window Fast design method comprises the following steps：
Step one, sets design conditions, including calculate time period, networking constraint, illumination angle and its change direction bar
The information such as part constraint, satellier injection trajectory；
Step 2, one-time calculation at the appointed time in section all windows for meeting networking constraint front and rear edge
Interval, is given in Julian date form, is called for follow-up link；
Step 3, one-time calculation at the appointed time in section all windows for meeting illumination condition constraint front and rear edges
Interval, is given in Julian date form, is called for follow-up link；
Step 4, by date along interval before and after merger, exports near-earth inclined plane launch window design.
Networking constraint and illumination condition constraint in step 2 and step 3 is unified to TrueOfDate coordinate systems
Recorded, prevented the design error of coordinate system reason generation.
Networking constraint in step 2 is disposably given using the interval of two star right ascension of ascending node difference as design constraint
Go out edge before and after all satisfactions in the specified time period constrain the corresponding launch window of interval bound to gather；
For drop rail transmitting, the right ascension of ascending node Ω of two stars2T () can be calculated as follows such as formula (1)：
In formula, t is to enter the orbit the time；
SGT () is transmitting Greenwich sidereal time day angle；λ is injection point longitude；It is injection point reduced latitude；I is track
Inclination angle；
To entering the orbit, time t is scanned for, when meeting Ω2(t)-Ω1Stopped search during (t)=Δ Ω.Launch time tLCan table
It is shown as such as formula (2)：
tL=t0–tPowered phase (2)
Networking constraint in step 2, its constraint process has high-precision orbital recursion ability, appoints in low rail constellation
In business, the window design of follow-up star can be directly participated in as input using the parameter of satellite in orbit under any epoch.
Illumination condition constraint in step 3 directly uses orbital plane illumination angle as input, is directly counted by a small amount of iteration
The satellier injection time is calculated, few, high precision is taken, it is to avoid with the traversal type algorithm time-consuming very long and precision that the time of entering the orbit is input
Low deficiency；This link exports 4 corresponding right ascension of ascending node according to β angles constraint bound, and for the treatment of follow-up link, (flow is joined
See Fig. 2).
During in view of illumination constraint form for β angles rather than southbound node place, and daily different position of sun influences whether β
Angle, therefore the result of calculation of launch window can be presented daily different situation.Calculating process is directed to respectively in the range of the whole anniversary
Daily position of sun calculates the right ascension of ascending node for meeting given β angles, and thus pushes away to obtain x time.
The computational methods of launch window are complex, and for ease of understanding, it is input to be given with β angles with formula form below
Launch window evaluation method, step is as follows：
(1) using a certain moment of scheduled date as the iteration initial value of solar vector S；
(2) right ascension of ascending node Ω can be tried to achieve by following relation, such as formula (3)：
Wherein, h is orbital angular momentum, and β is orbital plane β angles, and N is that track rises nodel line, and S is the sun
Vector (reads from DE405 ephemeris)；
(3) entering the orbit the time for current iteration can be tried to achieve by following relation：
The time t that enters the orbit can be tried to achieve by following formula0, such as formula (4)：
Wherein, ωeIt is rotational-angular velocity of the earth, Ω is right ascension of ascending node, λ,Respectively injection point longitude and latitude, SGIt is hair
Greenwich mean sidereal time during day 0 is penetrated, can be expressed from the next such as formula (5)：
SG=18h.697374558+879000h.051336907T+0s.093104T2 (5)
In formula, T is the Julian century；
(4) in the hope of t0It is input into as the next step iteration of solar vector S；(1)~(4) are repeated, until convergence；
(5) deduct transmitting the used time obtain launch time tL, such as formula (6)：
tL=t0–tPowered phase (6)
The design constraint of networking constraint is the evolution direction at β angles, on this basis to four of the output of front and continued link
Right ascension of ascending node differentiated, marks off two of which as edge, decision principle before and after one group of feasible interval correspondence launch window
It is as follows：
Track below 90 ° of inclination angle is western drift situation, for such situation, is made such as formula (7)：
Δ=Ω-λs+90°-180°≤Δ≤+180° (7)
In formula, λsIt is Sun Dec, Ω is current right ascension of ascending node.
It is change downwards when -90 °≤Δ≤+ 90 °；Otherwise it is change upwards.
Illumination condition constraint judges to belong to forward position only by edge before and after the interval bound calculation window of constraint by algorithm
Or after prolong, sorted out by logic, the interval internal and external parts of constraint greatlys save the calculating time without calculating.
Operation principle of the invention is as follows：
Step one, sets design conditions, including calculate time period (date), networking constraint, illumination angle and its change
The information such as direction constraint, satellier injection trajectory.
Step 2, by each orbit parameter unification to TrueOfDate coordinate systems, conversion regime is：First radical goes to position again
Speed coordinate, after go to new coordinate system, then go back to radical.
Orbital plane is located in perifocus coordinate system pqw, and p, q, w axle are respectively directed to the perigee direction of elliptic orbit, partly lead to
Footpath direction and orbital plane positive normal.Position and speed of the satellite in perifocus coordinate system are represented by such as formula (8) and (9)：
The X-axis of geocentric equatorial polar coordinate points to the first point of Aries, and Z axis point to the celestial sphere arctic, and Z axis obtain Y-axis with X-axis multiplication cross.Will
Geocentric equatorial polar coordinate rotates Ω angles about the z axis successively, and i angles are rotated around X-axis, is that can obtain pqw further around Z axis rotation ω angles
Coordinate system.The transformation matrix of coordinates for obtaining is rotated continuous three times to be represented by such as formula (10)：
Wherein, such as formula (11), (12), (13)：
So, the conversion for being tied to geocentric equatorial polar coordinate from perifocus coordinate is represented by such as formula (14)：
rXYZ=RXYZ←pqw·rpqw (14)
The conversion of coordinate system is expressed such as formula (15) by following formula：
HG=(NR) (PR) (15)
Wherein,It is the lower coordinate of TrueOfDate systems,It is coordinate under J2000 inertial systems.HG is the conversion exported by function
Matrix, successively including the precession of the equinoxes (PR), nutating (NR).They are expressed by following various (16) respectively.
(PR)=Rz(-zA)Ry(θA)Rz(-ζA)
(NR)=Rx(-Δε)Ry(Δθ)Rz(-Δμ) (16)
Wherein, ZA,θA,ζAIt is mean equator precession angle, is calculated by following equation
Δ μ, Δ θ, Δ ε are then respectively equation of the equinoxes, declination nutating and nutation in obliquity.
It it is the Greenwich mean sidereal time, such as formula (17)：
SG=18h.697374558+879000h.051336907T+0s.093104T2 (17)
Wherein,T is the UT1 times.
In low rail constellation task, using the parameter of satellite in orbit (input for being constrained for networking) under any epoch
The window design of follow-up star can be directly participated in as input；It is input into a star track condition and recursion is to moment epoch, using J4
Model recursion intends flat root, such as formula (18)：
(a0,e0,i0,Ω0,ω0,M0, t)=J4(a0,e0,i0,Ω0,ω0,M0,t0) (18)
In formula, t is recursion duration, a0,e0,i0,Ω0,ω0,M0It is initial time orbital tracking；Mapping J4Using the flat root of plan
Recurrence method, with high precision, fireballing advantage.
Step 3, directly uses orbital plane illumination angle (i.e. " β " angle) as input, and satellite is directly calculated by a small amount of iteration
Enter the orbit the time, take few, high precision, it is to avoid be time-consuming very long with traversal type algorithm that the time of entering the orbit is input and precision is low not
Foot；This link exports 4 corresponding right ascension of ascending node according to β angles constraint bound, for the treatment of follow-up link.In view of illumination about
When beam form is β angles rather than southbound node place, and daily different position of sun influences whether β angles, therefore the calculating of launch window
Result can be presented daily different situation.Calculating process is calculated for daily position of sun respectively in the range of the whole anniversary
Meet the right ascension of ascending node at given β angles, and thus push away to obtain x time.
Step 4, for after the window that calculates along than window forward position before examination in the case of, processed by classification and make it belong to
In two windows of not same date.
In sum, the present invention can the constraint of combing typical emission window, input, output is clear and definite, it is independent to it is each constrain into
Row is collected after calculating, the secondary development for being easy to increase and decrease to constrain, and is easy to satellite master-plan personnel to carry out launch window soon
Speed design, and ensure design accuracy.
Particular embodiments described above, technical problem, technical scheme and beneficial effect to solution of the invention are carried out
Further describe, should be understood that and the foregoing is only specific embodiment of the invention, be not limited to
The present invention, all any modification, equivalent substitution and improvements within the spirit and principles in the present invention, done etc., should be included in this
Within the protection domain of invention.
Claims (7)
1. a kind of near-earth inclined plane launch window Fast design method, it is characterised in that it comprises the following steps：
Step one, sets design conditions, including calculate time period, networking constraint, illumination angle and its change direction condition about
Beam, satellier injection trajectory；
Step 2, one-time calculation at the appointed time in section all windows for meeting networking constraint it is front and rear along area
Between, be given in Julian date form, called for follow-up link；
Step 3, one-time calculation at the appointed time in section all windows for meeting illumination condition constraint it is front and rear along area
Between, be given in Julian date form, called for follow-up link；
Step 4, by date along interval before and after merger, exports near-earth inclined plane launch window design.
2. near-earth inclined plane launch window Fast design method as claimed in claim 1, it is characterised in that the step 2
Unify to TrueOfDate coordinate systems to be recorded with the networking constraint in step 3 and illumination condition constraint.
3. near-earth inclined plane launch window Fast design method as claimed in claim 1, it is characterised in that the step 2
In networking constraint using two star right ascension of ascending node difference interval as design constraint, be disposably given in the specified time period
It is all to meet edge set before and after the corresponding launch window of the interval bound of constraint.
4. near-earth inclined plane launch window Fast design method as claimed in claim 1, it is characterised in that the step 2
In the constraint process of networking constraint there is high-precision orbital recursion ability, in low rail constellation task, gone through using any
The parameter of satellite in orbit under unit can directly participate in the window design of follow-up star as input.
5. near-earth inclined plane launch window Fast design method as claimed in claim 1, it is characterised in that the step 3
In illumination condition constraint directly use orbital plane illumination angle as input.
6. near-earth inclined plane launch window Fast design method as claimed in claim 4, it is characterised in that the networking bar
The design constraint of part constraint is the evolution direction at β angles.
7. near-earth inclined plane launch window Fast design method as claimed in claim 1, it is characterised in that the illumination bar
Part constraint only by constraining edge before and after interval bound calculation window, and by algorithm judge to belong to forward position or after prolong, by patrolling
Collect and sorted out.
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN108491647A (en) * | 2018-03-28 | 2018-09-04 | 中国科学院国家空间科学中心 | A kind of launch mission design system and its method for emergent application |
| CN110489779A (en) * | 2019-07-03 | 2019-11-22 | 上海卫星工程研究所 | A kind of jupiter's exploration swing-by trajectory optimum design method |
| CN112361887A (en) * | 2020-11-09 | 2021-02-12 | 北京理工大学 | Emission window planning method for near-earth target interception |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN108491647A (en) * | 2018-03-28 | 2018-09-04 | 中国科学院国家空间科学中心 | A kind of launch mission design system and its method for emergent application |
| CN108491647B (en) * | 2018-03-28 | 2021-11-09 | 中国科学院国家空间科学中心 | Launching task design system and method for emergency application |
| CN110489779A (en) * | 2019-07-03 | 2019-11-22 | 上海卫星工程研究所 | A kind of jupiter's exploration swing-by trajectory optimum design method |
| CN110489779B (en) * | 2019-07-03 | 2022-11-29 | 上海卫星工程研究所 | Optimization design method for Mars exploration assisted flight orbit |
| CN112361887A (en) * | 2020-11-09 | 2021-02-12 | 北京理工大学 | Emission window planning method for near-earth target interception |
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