CN106777580B - Method for rapidly designing emission window of near-earth inclined orbit - Google Patents

Method for rapidly designing emission window of near-earth inclined orbit Download PDF

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CN106777580B
CN106777580B CN201611086251.1A CN201611086251A CN106777580B CN 106777580 B CN106777580 B CN 106777580B CN 201611086251 A CN201611086251 A CN 201611086251A CN 106777580 B CN106777580 B CN 106777580B
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许哲
沈庆丰
叶小舟
黄欣
李绿萍
李楠
陆启省
李鉴
邓武东
袁双
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Shanghai Institute of Satellite Engineering
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Abstract

The invention discloses a method for quickly designing a launching window of a near-earth inclined orbit, which comprises the following steps: step one, setting calculation conditions, including a calculation time period, networking condition constraints, illumination angle and change direction condition constraints and satellite orbit trajectory; secondly, calculating the front and back edge intervals of all windows meeting the networking condition constraint in a specified time period at one time, and giving the intervals in a julian day form for calling the subsequent links; and step three, calculating the front and back edge intervals of all windows meeting the light condition constraint in a specified time period at one time, and giving the intervals in a julian day form for calling in subsequent links. The invention can comb typical emission window constraints, has clear input and output, independently calculates each constraint and then collects the constraints, is convenient for secondary development of increasing and decreasing constraints, is convenient for global satellite designers to rapidly design emission windows, and ensures the design precision.

Description

Method for rapidly designing emission window of near-earth inclined orbit
Technical Field
The invention relates to a design method, in particular to a method for quickly designing a near-earth inclined orbit launching window.
Background
The low-earth inclined (non-solar synchronous) orbit satellite has the characteristics of traversal, freezing, good low-latitude coverage and the like, is suitable for various application scenes, and is increasingly adopted by satellite designers. The same characteristics as the daily emission windows of the sun-synchronous orbit are different, and the daily emission windows of the non-sun-synchronous orbit are different. For window calculation, currently, a commonly-used traversal calculation method sequentially determines whether the track characteristic corresponding to each time infinitesimal meets the design constraint, and arranges all the infinitesimal sets meeting the conditions to be called as emission windows. The method is simple and reliable, but the calculation amount is large, if window calculation of a whole year is carried out, the calculation time is long, and the precision is also influenced by the size of a infinitesimal element.
With the development of the aerospace technology, the development period of the satellite is shortened, and the development task is increased, and the design requirements of the satellite at the present stage cannot be met due to the problems of low design speed and insufficient design precision of the traditional design method. Therefore, a fast algorithm capable of ensuring the calculation accuracy needs to be developed, so that the overall mission planner can still design the emission window fast and accurately without being familiar with the professional knowledge of the orbit, and the development efficiency is improved.
Disclosure of Invention
The invention aims to solve the technical problem of providing a method for quickly designing a launching window of a near-earth inclined orbit, which can comb typical launching window constraints, has definite input and output, independently calculates each constraint and then collects the constraints, is convenient for increasing and decreasing the secondary development of the constraints, is convenient for a global satellite designer to quickly design the launching window, and ensures the design precision.
The invention solves the technical problems through the following technical scheme: a method for rapidly designing a launching window of a near-earth inclined orbit comprises the following steps:
step one, setting calculation conditions, including a calculation time period, networking condition constraints, illumination angle and change direction condition constraints and satellite orbit trajectory;
secondly, calculating the front and back edge intervals of all windows meeting the networking condition constraint in a specified time period at one time, and giving the intervals in a julian day form for calling the subsequent links;
thirdly, calculating the front and back edge intervals of all windows meeting the constraint of the illumination condition in a specified time period at one time, and giving the intervals in a julian day form for calling the subsequent links;
and step four, merging the front and rear edge intervals according to the date, and outputting a design scheme of a near-earth inclined track emission window.
Preferably, the networking condition constraint and the lighting condition constraint in the second step and the third step are both unified to the TrueOfDate coordinate system for recording.
Preferably, the networking condition constraint in the second step takes an interval of the right ascension difference between two stars, and gives a set of front and rear edges of the emission window corresponding to all the upper and lower limits meeting the constraint interval in a specified time period at one time.
Preferably, the networking condition constraint in the second step has a high-precision orbit recursion capability in the constraint process, and in the low-orbit constellation task, the existing orbit satellite parameters in any epoch can be used as input to directly participate in the window design of the subsequent satellites.
Preferably, the lighting condition constraint in step three directly uses the orbital plane lighting angle as input.
Preferably, the light condition constraint only calculates the front and rear edges of the window through the upper and lower limits of the constraint interval, judges whether the window belongs to the front edge or the rear edge through an algorithm, and classifies the window according to logic.
The positive progress effects of the invention are as follows: the invention can comb typical emission window constraints, has clear input and output, independently calculates each constraint and then collects the constraints, is convenient for secondary development of increasing and decreasing constraints, is convenient for global satellite designers to rapidly design emission windows, and ensures the design precision.
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FIG. 1 is a block flow diagram of the present invention.
FIG. 2 is a block diagram of the illumination-constrained window calculation of the present invention.
Fig. 3 is a diagram illustrating the conversion relationship between the track root and the position velocity.
Detailed Description
The following provides a detailed description of the preferred embodiments of the present invention with reference to the accompanying drawings.
As shown in fig. 1 to 3, the method for rapidly designing the emission window of the near-earth inclined orbit comprises the following steps:
step one, setting calculation conditions, including information such as a calculation time period, networking condition constraints, illumination angle and change direction condition constraints, satellite orbit-entering trajectory and the like;
secondly, calculating the front and back edge intervals of all windows meeting the networking condition constraint in a specified time period at one time, and giving the intervals in a julian day form for calling the subsequent links;
thirdly, calculating the front and back edge intervals of all windows meeting the constraint of the illumination condition in a specified time period at one time, and giving the intervals in a julian day form for calling the subsequent links;
and step four, merging the front and rear edge intervals according to the date, and outputting a design scheme of a near-earth inclined track emission window.
And step two and step three, the networking condition constraint and the illumination condition constraint are unified to a TrueOfDate coordinate system for recording, and the design error caused by the coordinate system is avoided.
In the networking condition constraint in the step two, an interval of the right ascension difference of two star ascending intersection points is used as a design constraint, and all emission window front and rear edge sets corresponding to upper and lower limits of the constraint interval in a specified time period are given at one time;
for down-track launch, the ascent point of two stars is the right ascension omega2(t) can be calculated as in formula (1) as follows:
Figure GDA0002299013100000031
in the formula, t is the track entering time;
SG(t) is the emission daily greenwich mean sidereal hour angle; λ is the track entry point longitude;
Figure GDA0002299013100000032
the geocentric latitude of the orbit entering point is obtained; i is the track inclination angle;
searching the time t of track entering when the omega is satisfied2(t)-Ω1And stopping searching when the (t) is equal to delta omega. Time of transmission tLCan be expressed as formula (2):
tL＝t0–tactive section(2)
And step two, the networking condition constraint is carried out, the constraint process has high-precision orbit recursion capability, and existing orbit satellite parameters in any epoch can be used as input to directly participate in the window design of subsequent satellites in the low orbit constellation task.
The illumination condition constraint in the third step directly uses the orbit plane illumination angle as input, directly calculates the satellite orbit entering time through a small amount of iteration, has less time consumption and high precision, and avoids the defects of very long time consumption and low precision of a traversal algorithm taking the orbit entering time as input, and the link outputs 4 corresponding rising-intersection red meridians according to β angle constraint upper and lower limits for subsequent link processing (see a flow chart in figure 2).
The calculation process calculates the ascent point right ascent points satisfying a given β angle for each daily sun position throughout the whole year, and deduces the emission time accordingly.
The calculation method of the emission window is complicated, and for the convenience of understanding, the emission window estimation method taking β angle as input is given in the following formula form, and the steps are as follows:
(6) taking a certain moment of a specified date as an iteration initial value of the sun vector S;
(7) the crossover right ascension Ω can be obtained from the following relationship, as shown in formula (3):
Figure GDA0002299013100000051
where h is orbital angular momentum, β is orbital plane β angle, N is orbital elevation line, and S is sun vector (read from DE405 ephemeris);
(8) the orbit entering time of the iteration can be obtained according to the following relation:
the track-in time t can be obtained by the following formula0As in formula (4):
Figure GDA0002299013100000052
wherein, ω iseIs the angular velocity of the earth's rotation, omega is the ascension, the right ascension, lambda,
Figure GDA0002299013100000053
Respectively longitude and latitude, S of the point of entryGIn greenwich mean time of day 0 of launch, can be represented by the following formula (5):
SG＝18h.697374558+879000h.051336907T+0s.093104T2(5)
in the formula, T is the Ru-Shi century;
(9) by the found t0As the next iteration input of the sun vector S; repeating the steps (1) to (4) until convergence;
(10) obtaining emission time t by deducting emission timeLAs in formula (6):
tL＝t0–tactive section(6)
The design constraint of networking condition constraint is an evolution direction of β angles, four intersection points right ascension output by a previous link are judged according to the evolution direction, two of the four intersection points right ascension are divided into a group of feasible intervals corresponding to the front edge and the rear edge of an emission window, and the judgment principle is as follows:
trajectories with an inclination below 90 ° are the case for west drifts, for which case the equation (7) is given:
Δ＝Ω-λs+90°-180°≤Δ≤+180° (7)
in the formula, λsThe right ascension and the omega the current ascending crossing right ascension.
When the angle is less than or equal to minus 90 degrees and less than or equal to plus 90 degrees, the angle is changed downwards; otherwise it is changed upwards.
The light condition constraint only calculates the front edge and the rear edge of the window through the upper limit and the lower limit of the constraint interval, judges the front edge or the rear edge through an algorithm, and classifies the front edge and the rear edge according to logic, and the inner part and the outer part of the constraint interval do not need to be calculated, so that the calculation time is greatly saved.
The working principle of the invention is as follows:
step one, setting calculation conditions, including information such as a calculation time period (date), networking condition constraints, illumination angles and change direction constraints thereof, satellite orbit-entering trajectory and the like.
Step two, unifying all track parameters to a TrueOfDate coordinate system, wherein the conversion mode is as follows: the number is firstly converted to the position speed coordinate, then to the new coordinate system, and then to the number.
The orbital plane is located in a near-focus coordinate system pqw with the p, q, and w axes pointing in the perigee, semi-path, and orbital plane normal directions, respectively, of the elliptical orbit. The position and velocity of the satellite in the near-focus coordinate system can be expressed as in equations (8) and (9):
Figure GDA0002299013100000061
Figure GDA0002299013100000062
the X axis of the geocentric equator coordinate system points to the spring equinox, the Z axis points to the celestial north pole, and the Y axis is obtained by the cross multiplication of the Z axis and the X axis. And sequentially rotating the geocentric equator coordinate system by an angle omega around the Z axis, rotating the geocentric equator coordinate system by an angle i around the X axis, and rotating the geocentric equator coordinate system by an angle omega around the Z axis to obtain the pqw coordinate system. The coordinate transformation matrix obtained by three consecutive rotations can be expressed as formula (10):
RXYZ←pqw＝Rz(ω)*Rx(i)*Rz(Ω) (10)
wherein, the formula is (11), (12) and (13):
Figure GDA0002299013100000063
Figure GDA0002299013100000071
Figure GDA0002299013100000072
thus, the transformation from the near-focus coordinate system to the geocentric equatorial coordinate system can be expressed as equation (14):
rXYZ＝RXYZ←pqw·rpqw(14)
the transformation of the coordinate system is expressed by the following formula (15):
Figure GDA0002299013100000073
HG＝(NR)(PR) (15)
wherein the content of the first and second substances,
Figure GDA0002299013100000074
is the coordinate under the TrueOfDate system,
Figure GDA0002299013100000075
is coordinates under the J2000 inertial system. HG is the transformation matrix output by the function, including the age (PR), the Nutation (NR) successively. They are respectively expressed by the following respective formulae (16).
(PR)＝Rz(-zA)Ry(θA)Rz(-ζA)
(NR)＝Rx(-Δ)Ry(Δθ)Rz(-Δμ) (16)
Wherein Z isA,θA,ζAIs the equatorial slip angle, calculated by the following formula
Δ μ, Δ θ, Δ are the right ascension nutation, declination nutation and crossing angle nutation, respectively.
Figure GDA0002299013100000076
Is Greenwich mean time, as shown in formula (17):
SG＝18h.697374558+879000h.051336907T+0s.093104T2(17)
wherein the content of the first and second substances,
Figure GDA0002299013100000077
t is UT1 time.
In the low-orbit constellation task, the existing orbit satellite parameters (namely, the input for networking constraint) under any epoch can be used as input to directly participate in the window design of the subsequent satellite; inputting a satellite orbit state and recursion to epoch time, and recursion to a parallel root by adopting a J4 model, wherein the formula is as follows (18):
(a0,e0,i0,Ω0,ω0,M0,t)＝J4(a0,e0,i0,Ω0,ω0,M0,t0) (18)
where t is the recurrence duration, a0,e0,i0,Ω0,ω0,M0Is the number of tracks at the initial moment; mapping J4The method has the advantages of high precision and high speed by adopting a quasi-flat root recurrence method.
And step three, directly using the orbit plane illumination angle (namely the angle β) as input, directly calculating the satellite orbit entering time through a small amount of iteration, consuming less time, having high precision, and avoiding the defects of long time consumption and low precision of a traversal algorithm taking the orbit entering time as input, outputting 4 corresponding ascent point right ascension points according to β angle constraint upper and lower limits for subsequent link processing, considering that the illumination constraint form is β angle instead of descending intersection point, and different sun positions per day can influence β angle, so the calculation result of the emission window can present different conditions per day, respectively calculating the ascent point right ascension points meeting the given β angle for the sun positions per day in the whole anniversary range in the calculation process, and deducing the emission time.
And step four, classifying the window back edge obtained by calculation to be the window with two different dates under the condition that the window front edge is examined.
In conclusion, the invention can comb typical emission window constraints, has clear input and output, independently calculates each constraint and then collects the constraints, is convenient for secondary development of increase and decrease constraints, is convenient for global satellite designers to rapidly design emission windows, and ensures the design precision.
The above embodiments are described in further detail to solve the technical problems, technical solutions and advantages of the present invention, and it should be understood that the above embodiments are only examples of the present invention and are not intended to limit the present invention, and any modifications, equivalent substitutions, improvements and the like made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims (7)

1. A method for rapidly designing a launching window of a near-earth inclined orbit is characterized by comprising the following steps:
step one, setting calculation conditions, including a calculation time period, networking condition constraints, illumination angle and change direction condition constraints and satellite orbit trajectory;
secondly, calculating the front and back edge intervals of all windows meeting the networking condition constraint in a specified time period at one time, and giving the intervals in a julian day form for calling the subsequent links;
the networking condition constraint takes the interval of the right ascension difference of the two-star ascending intersection point as a design constraint, and gives a front edge set and a rear edge set of all emission windows corresponding to the upper limit and the lower limit of the constraint interval in a specified time period at one time;
for down-track launch, the ascent point of two stars is the right ascension omega2(t) is calculated according to the formula (1):
Figure FDA0002483038040000011
in the formula, t is the track entering time;
SG(t) is the emission daily greenwich mean sidereal hour angle; λ is the track entry point longitude;
Figure FDA0002483038040000012
the geocentric latitude of the orbit entering point is obtained; i is the track inclination angle;
searching the time t of track entering when the omega is satisfied2(t)-Ω1Stopping searching when (t) is delta omega, and transmitting for time tLRepresented by formula (2):
tL＝t0–tactive section(2)；
The emission window estimation method taking the track surface illumination angle β as input comprises the following steps:
step S1, taking a certain time of the appointed date as an iteration initial value of the sun vector S;
step S2, the rising-point right ascension Ω is obtained from equation (3):
Figure FDA0002483038040000021
wherein h is orbital angular momentum, β is orbital plane β angle, N is orbital elevation line, and S is sun vector;
step S3, the tracking time of the current iteration is obtained from the following relationship:
the track-in time t is obtained from the following equation (4)0：
Figure FDA0002483038040000022
Wherein, ω iseIs the angular velocity of the earth's rotation, omega is the ascension, the right ascension, lambda,
Figure FDA0002483038040000023
Respectively longitude and latitude, S of the point of entryGIs greenwich mean time at transmission day 0, represented by formula (5):
SG＝18h.697374558+879000h.051336907T+0s.093104T2(5)
in the formula, T is the Ru-Shi century;
step S4, using the obtained t0As the next iteration input of the sun vector S; repeating the steps (1) to (4) until convergence;
step S5, deducting the transmitting time to obtain the transmitting time tL：
tL＝t0–tActive section(6)；
Thirdly, calculating the front and back edge intervals of all windows meeting the constraint of the illumination condition in a specified time period at one time, and giving the intervals in a julian day form for calling the subsequent links;
and step four, merging the front and rear edge intervals according to the date, and outputting a design scheme of a near-earth inclined track emission window.
2. The method as claimed in claim 1, wherein the networking constraint and the lighting constraint in step two and step three are unified to TrueOfDate coordinate system for recording.
3. The method according to claim 1, wherein the networking condition constraint in the second step takes an interval of a right ascension difference between two stars higher than the ground as a design constraint, and gives a set of front and rear edges of the emission window corresponding to all upper and lower limits of the constraint interval within a specified time period at one time.
4. The method according to claim 1, wherein the constraint process of the networking condition constraint in the second step has a high-precision orbit recursion capability, and in the low-orbit constellation task, the existing orbit satellite parameters in any epoch can be used as input to directly participate in the window design of the subsequent satellites.
5. The method of claim 1, wherein the illumination condition constraints in step three directly use the orbital plane illumination angle as input.
6. The method as claimed in claim 4, wherein the design constraint of the networking condition constraint is an evolution direction of β degrees.
7. The method for rapidly designing the emission window of the near-earth inclined orbit as claimed in claim 1, wherein the light condition constraint only calculates the front and rear edges of the window through the upper and lower limits of the constraint interval, judges whether the window belongs to the front edge or the rear edge through an algorithm, and classifies the window according to logic.
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