CN115169108A - Simulation calculation method for satellite on-orbit running real-time sunned state - Google Patents
Simulation calculation method for satellite on-orbit running real-time sunned state Download PDFInfo
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Abstract
The invention provides a simulation calculation method for the on-orbit running real-time sunned state of a satellite, which comprises the following steps: establishing a time description system; establishing a spatial position description system; establishing a space star position calculation method based on ephemeris to obtain the relative positions of the sun and the earth in a universe space; establishing a satellite orbit extrapolation calculation model, and calculating the spatial position of a satellite at any moment; according to the time input by the time description system in the step S1, calculating to obtain the relative positions of the sun and the earth in the space and the space position of the satellite, then converting into the space position description system established in the step S2, and obtaining the sun-exposed state of the satellite by an earth shadow calculation method; and calculating the sun exposure state of the satellite in a specific time period. The simulation calculation method for the in-orbit running real-time sunned state of the satellite can obtain the real-time state data of the sunned state of the satellite during the space running by using the simulation method, and is low in cost, low in consumption, high in efficiency and high in reliability.
Description
Technical Field
The invention relates to the technical field of space engineering and computer simulation, in particular to a simulation calculation method for an in-orbit running real-time sunned state of a satellite.
Background
With the continuous advance of human space technology, the activity of human exploration space is continuously increased, so that the operation environment of the satellite is complex and severe, and the environmental parameters are changed violently. For example, the sun illumination can have a significant effect on the external heat flow environment of the satellite, the surface temperature of the satellite can reach above 100 ℃ when the satellite is illuminated by the sun, and the temperature can be reduced to below-100 ℃ when the satellite is not illuminated by the sun.
When the satellite moves circularly around the earth, the satellite can be alternately subjected to the heat radiation of sunlight, so that the satellite generates heat environment change and the on-orbit reliable operation of the satellite is influenced. In addition, whether the satellite is in the shadow of the earth also has an impact on the energy source of the satellite. The main energy source of the satellite in the operation period in the outer space is solar energy, and the solar energy is converted into electric energy to be stored and used by irradiating the solar cell panel through the solar optical fiber, so that the irradiation condition and the irradiation duration of the solar light have a vital effect on the energy guarantee on the satellite.
In the prior art, many discussions and researches on earth shadow calculation and satellite illumination duration have been carried out, but the study is limited to theoretical derivation and research on specific orbits, and a lot of inconveniences exist in engineering use and application of the overall designer. Therefore, a method for calculating the real-time sun exposure state (i.e. sun illumination) of the satellite during the in-orbit operation is needed.
Disclosure of Invention
The invention provides a simulation calculation method for the real-time sunned state condition of a satellite during in-orbit operation, which is convenient for engineering use and application of a total designer.
In order to solve at least one aspect of the above problems, the present invention provides a simulation calculation method for an in-orbit running real-time solarization state of a satellite, comprising the following steps:
s1, establishing a time description system, wherein coordinated world time is used as a user description standard, and a julian day is used for internal simulation calculation;
s2, establishing a spatial position description system, and using a J2000 geocentric inertial coordinate system for simulation calculation by adopting a Cartesian rectangular coordinate system;
s3, establishing a space star position calculation method based on ephemeris to obtain the relative positions of the sun and the earth in the space;
s4, establishing a satellite orbit extrapolation calculation model, defining a satellite orbit by using six Kepler orbits, and calculating the spatial position of a satellite at any moment;
s5, inputting time according to the time description system in the step S1, calculating to obtain the relative positions of the sun and the earth in the space and the space position of the satellite, then converting into the space position description system established in the step S2, and obtaining the sun exposure state of the satellite through an earth shadow calculation method;
and S6, setting task starting time, task ending time and simulation step length, sequentially accumulating the simulation step length from the task starting time as simulation calculation time, and calculating the sun exposure state of the satellite according to the method of the step S5 until the sun exposure state is accumulated to the task ending time.
Preferably, in the step S1, the user uses the coordinated universal time as a description standard, the description form is year, month, day, hour, minute, and second of a gregorian calendar, and the coordinated universal time is converted into julian day in a simulation calculation process.
Preferably, in the step S2, the origin of coordinates of the J2000 centroid inertial coordinate system coincides with the centroid of the earth, the X axis of the J2000 centroid inertial coordinate system points to the J2000 time division point, the Z axis of the J2000 centroid inertial coordinate system points to the north pole, and the Y axis, the X axis, and the Z axis of the J2000 centroid inertial coordinate system form a right-handed rectangular coordinate system.
Preferably, in step S3, a space-celestial-sphere position calculation method based on ephemeris is established by using JPL ephemeris data, so as to obtain the relative positions of the sun and the earth in the space.
Preferably, a space celestial sphere position calculation method based on ephemeris is established by adopting ephemeris data in a DE405 ephemeris, and the relative positions of the sun and the earth in the space are obtained; the DE405 ephemeris adopts a coordinate system which is a rectangular coordinate system with the sun system centroid as the origin, the J2000 earth equatorial plane as the xy plane, and the J2000 vernal equinox direction as the x direction.
Preferably, in step S4, the six kepler orbit components include a semi-major axis a, an eccentricity e, an inclination angle i, a right ascension angle Ω of a rising intersection, an argument ω of a near point, and a true near point angle f.
Preferably, the semi-major axis a, the eccentricity e, the inclination angle i, the ascension angle Ω of the ascending intersection point, the argument ω of the perigee, and the true perigee angle f are calculated by equations (1) to (6):
wherein μ is an earth gravity constant, μ = GM =3.986005 × 10 14 m 3 /s 2 G is a gravitational constant, and M is the earth mass; let t 0 Time of day, in the inertial system, the velocity vector of the satellite
And position vector
Are respectively as
And
the magnitude of velocity v 0 And the position size r 0 Are respectively as
And
momentum moment of the orbit of
And
preferably, in step S4, satellite trajectory data at any time is calculated by using an SGP4 model and using the initialized six kepler orbits and time, so as to obtain the spatial position of the satellite at any time.
Preferably, in step S5, the earth shadow calculating method includes:
step T1, calculating sun parallel ray prevention vector v 1 ；
Step T2, calculating a direction vector v from the geocentric to the satellite 2 ；
Step T3, calculating v 1 And v 2 Angle theta therebetween 1 ；
Step T4, calculating a critical angle theta 2 Wherein the critical angle refers to the minimum value of the earth shadow angle in a circular orbit with the distance from the satellite to the geocenter as a radius;
step T5, if theta 1 ＞θ 2 The satellite is in the sunned state if theta 1 ＜θ 2 Then the satellite is in a non-sunned state.
Preferably, in step S6, the simulation step size is 20S.
The method is characterized by establishing a proper time description system and a proper space position description system by combining a computer simulation technology, calculating the relative positions of the sun and the earth in the space by a space planet position calculation method of ephemeris, then obtaining the space position of a satellite by an established satellite orbit extrapolation model, then unifying the relative positions of the sun and the earth in the space and the space position of the satellite in the established time description system and space position description system, and obtaining the sun-exposed state of the satellite at any moment by an earth shadow algorithm, wherein the sun-exposed state can be conveniently used by a user by adopting the coordinated world as a user description standard, and the time calculation can be carried out by using continuous real numbers in the simulation calculation process by adopting the julian sun for internal simulation calculation; in addition, the sun exposure state of the satellite in the target time period can be obtained by setting the task starting time, the task ending time and the simulation step length, and sun exposure state data with different accuracies can be obtained by adjusting the simulation step length; the simulation calculation method for the in-orbit running real-time sunned state of the satellite can obtain the real-time state data of the sunned state of the satellite during the space running by using the simulation method, and is low in cost, low in consumption, high in efficiency and high in reliability.
Drawings
FIG. 1 is a flowchart of a simulation calculation method for an in-orbit operation real-time solarization state of a satellite according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of a J2000 geocentric inertial coordinate system in an embodiment of the invention;
FIG. 3 is a schematic diagram of eccentricity and satellite orbit shape in an embodiment of the present invention;
FIG. 4 is a diagram illustrating six Kepler orbits of a satellite according to an embodiment of the present invention;
FIG. 5 is a diagram illustrating a geocentric shadow calculation method according to an embodiment of the invention;
FIG. 6 is a schematic diagram illustrating a process of calculating a sun exposure status of a satellite during a specific time period according to an embodiment of the present invention;
FIG. 7 is a diagram illustrating a periodic variation of the exposure status of a satellite in accordance with an embodiment of the present invention;
FIG. 8 is a diagram of the spatial variation of a satellite around the earth in an embodiment of the present invention.
Detailed Description
In order to make the aforementioned objects, features and advantages of the present invention more comprehensible, embodiments of the present invention are described in detail below.
It should be noted that the features in the embodiments of the present invention may be combined with each other without conflict. The terms "comprising," "including," "containing," and "having" are intended to be inclusive, i.e., that additional steps and other ingredients may be added without affecting the result. The above terms encompass the terms "consisting of … …" and "consisting essentially of … …". Materials, equipment and reagents are commercially available unless otherwise specified. Also, it is noted that the terms "first," "second," and the like in the description and claims of the present invention and in the above-described drawings are used for distinguishing between similar elements and not necessarily for describing a particular sequential or chronological order.
In general, the application of engineering design can design and calculate the design parameters of the satellite orbit, such as the height of the orbit, the inclination angle of the orbit, the eccentricity, the right ascension of the ascending intersection point, the range of the near place, and the like, and obtain the results of different sunned states. For detailed research of a specific space position in a specific time period, real-time solarization state change data of satellite operation is needed, and the conventional research is difficult to consider the scenes of the engineering application.
The embodiment of the invention provides a simulation calculation method for an on-orbit running real-time sunned state of a satellite, which comprises the following steps as shown in figure 1:
s1, establishing a time description system, wherein the coordinated world time is used as a user description standard, and a julian day is used for internal simulation calculation;
s2, establishing a spatial position description system, and using a J2000 geocentric inertial coordinate system for simulation calculation by adopting a Cartesian rectangular coordinate system;
s3, establishing a space star position calculation method based on ephemeris to obtain the relative positions of the sun and the earth in a space;
s4, establishing a satellite orbit extrapolation calculation model, defining a satellite orbit by using six Kepler orbit roots, and calculating the space position of the satellite at any moment;
s5, inputting time according to the time description system in the step S1, calculating to obtain the relative positions of the sun and the earth in the space and the space position of the satellite, then converting into the space position description system established in the step S2, and obtaining the sun exposure state of the satellite through an earth shadow calculation method;
and S6, setting task starting time, task ending time and simulation step length, sequentially accumulating the simulation step length from the task starting time as simulation calculation time, and calculating the sun exposure state of the satellite according to the method of the step S5 until the sun exposure state is accumulated to the task ending time.
In the step S1, the user adopts the coordinated universal time as a description standard, the description form is year, month, day, hour, minute and second of a gregorian calendar, and the coordinated universal time is converted into julian days in the simulation calculation process. I.e., the user takes the form of universal time coordination input, and the coordinated time input needs to be converted to julian days before being used in the simulation calculations.
Coordinated Universal Time (UTC), also known as universal time, is based on atomic seconds. To ensure that the universal time coordinated does not differ from the universal time by more than 0.9 seconds, the international terrestrial rotation office in paris will decide to add leap seconds to the universal time coordinated if necessary. In the simulation calculation method for the in-orbit running real-time sunned state of the satellite, the transmitting time and the current position of the satellite are input by adopting the UTC, and the year, month, day, hour, minute and second of the gregorian calendar are used as a time description form, so that the simulation calculation method can be conveniently used and set by engineering designers.
In order to ensure that the simulation calculation method of the satellite in-orbit running real-time sunned state can use a continuous real number to calculate time, the simulation calculation adopts julian days. The julian day is the number of days with the initial point of 4573 years 1 month 1day 12h, and by converting UTC into julian day, the description of the difference between two common calendar moments can be obtained under unified specification, so that the simulation step length can be conveniently formulated, and the julian day is also suitable for the calculation of the self-rotation angle of the earth.
Specifically, the gregorian calendar is converted into julian days by formula (7):
in the formula, INT () is rounded, and Year, month, day, hour, minute, and Second are calendar Year, month, day, hour, minute, and Second, respectively.
In step S2, in order to describe the position of the sun, the earth, and the ray direction of the sun in space, and also to describe the speed of the satellite in space, a suitable reference coordinate system needs to be established, in the embodiment of the present invention, a cartesian rectangular coordinate system is adopted, and a J2000 geocentric inertial coordinate system is used for calculation in simulation calculation.
The J2000 geocentric inertial coordinate system defines a celestial sphere reference coordinate system by using the equator and the dichotomy point at the time of J2000 (1/12 in 2000), as shown in fig. 2, the origin of coordinates of the J2000 geocentric inertial coordinate system coincides with the centroid of the earth, the X-axis of the J2000 geocentric inertial coordinate system points to the equatorial patch at the time of J2000, the Z-axis of the J2000 geocentric inertial coordinate system points to the north pole, and the Y-axis of the J2000 geocentric inertial coordinate system, the X-axis and the Z-axis form a right-hand rectangular coordinate system.
In step S3, in order to obtain the relative positions of the sun and the earth in the space through simulation calculation, an ephemeris-based space celestial sphere position calculation method is established.
JPL ephemeris gives past and future location information for the sun, moon and the nine major planets and is open for licensed use. The JPL ephemeris is established by a jet propulsion laboratory in the 60 th 20 th century, is initially used for the purpose of planet exploration navigation, and is continuously corrected and improved along with the continuous improvement of observation technology and continuous acquisition of new observation data. JPL ephemeris is the result of numerical integration of a system of differential equations describing the solar system dynamics, and is therefore established based on two assumptions: (1) The system of differential equations accurately represents the known laws of dynamics, at least with the present accuracy of observation; and (2) the precision of the numerical integration program is high enough. Under these two assumptions, a dynamical system can be constructed, which is consistent with the solar system dynamical system, and it is also necessary to see if the initial conditions and parameters are consistent, and the initial conditions and parameters are subjected to least squares fitting by observing data. The observed data includes: ranging data of a planetary exploration spacecraft, radar planetary ranging data, lunar laser ranging, angle measurement data of optical observation and some latest measurement means. In order to accurately represent the celestial body position in a long time range, the JPL ephemeris divides the long time range (hundreds of years) into short time intervals (days), provides a group of Chebyshev interpolation coefficients for each short time interval, calculates the celestial body position at a certain moment by firstly finding the short time interval to obtain the Chebyshev interpolation coefficients, and then calculates the celestial body position according to a Chebyshev interpolation formula.
Commonly used JPL ephemeris is DE200, DE403, DE405 and DE406, and DE405 ephemeris data is selected for use in the embodiment of the present invention, and the ephemeris covers anywhere from 12 months 9 days 1599 to 2 months 20 days 2201.
Converting a known instant time (Greenwich calendar date + UTC time) into a julian day (julian epoch) and inputting the julian day (julian epoch) into a JPL-DE405 ephemeris, obtaining the positions and the speeds of the sun, the moon and each planet under ICRS (BCRS), and then switching to ITRS or J2000, wherein ICRS is an international celestial sphere reference system, BCRS is a centroid celestial sphere reference system, ITRS is an earth reference system, and J2000 is a J2000 geocentric inertial coordinate system. The coordinate system adopted by the DE405 ephemeris is a rectangular coordinate system with the centroid of the solar system as the origin, the equatorial plane of the J2000 earth as the xy plane, and the vernal point direction of the J2000 as the x direction, and the position coordinates obtained by interpolation are values in this coordinate system (except for the moon, the moon coordinates use the centroid as the origin).
If the acquired celestial body position in other coordinate systems needs to be further changed, translation and rotation change of coordinates need to be carried out, an inertial coordinate system used in orbital dynamics generally takes the earth centroid as an origin, the J2000 earth equator as an xy plane and the vernal point direction as the x direction, the coordinates of DE405 are converted into a geocentric inertial coordinate system, and only coordinate translation needs to be carried out. The time unit of DE405 is day, 1day =86400s, (SI) 1day =86400s (SI), the distance unit is km, and the speed unit is km/day.
In step S4, a satellite orbit extrapolation calculation model is established by adopting an SGP4 model, and the space position of the satellite at any moment is calculated by defining the satellite orbit by using six numbers of Kepler orbits.
The six Kepler orbits comprise a semi-long axis a, an eccentricity e, an inclination angle i, a rising intersection right ascension omega, an perigee amplitude angle omega and a true perigee angle f. (ii) a The semimajor axis a is used for describing the size of the orbit, the eccentricity e is used for describing the shape of the orbit, the inclination angle i is used for describing the inclination angle of the orbit plane relative to the equatorial plane, the ascension point right ascension omega is used for describing the position of the ascension point relative to the vernal point, the perigee amplitude omega is used for describing the position of the perigee relative to the ascension point, and the true perigee angle is used for describing the position of the satellite relative to the perigee.
Specifically, the semimajor axis a, the eccentricity e, the inclination angle i, the ascension angle Ω of the intersection point, the argument ω of the perigee, and the true perigee angle f are calculated by the formulas (1) to (6):
wherein μ is an earth gravity constant, μ = GM =3.986005 × 10 14 m 3 /s 2 G is a gravitational constant, and M is the earth mass; let t0 be the time, in the inertial system, the velocity vector of the satellite
And position vector
Are respectively as
And
the magnitude of velocity v 0 And the position size r 0 Are respectively as
And
momentum moment of the orbit of
And
as shown in fig. 3, the eccentricity e determines the type of orbit shape, including circular orbit, elliptical orbit, parabola and hyperbola, and the semimajor axis a is an important parameter of the elliptical orbit and has a value equal to half the distance from the apogee to the perigee of the orbit of the aircraft.
As shown in fig. 4, the x-axis of the coordinate system points to the vernal equinox from the center of the earth, the z-axis points to the north pole from the center of the earth, the arc BPD (i.e. the arc of the dotted line part in the figure) is the track of the satellite subtotal point, point a is the vernal equinox point, point B is the point where the aircraft passes through the equatorial plane of the earth from south to north during on-orbit operation, called the elevation intersection point, point D is the point where the aircraft passes through the equatorial plane of the earth from south to north during on-orbit operation, called the descent intersection point, point P is the projection of the near point of the orbit on the track of the subtotal point, point S represents the position where the aircraft is located at the current moment of the orbit, the inclination angle of the orbit i describes the inclination angle of the orbital plane relative to the equatorial plane, the angle range is 0-180 °, angle AOB is the ascent intersection point Ω in the figure, angle BOP is the near point depression angle ω, and angle POS is the true near point angle f in the figure.
In the step S5, inputting time according to the time description system in the step S1, calculating to obtain the relative positions of the sun and the earth in the space and the space position of the satellite, then converting into the space position description system established in the step S2, and obtaining the sun exposure state of the satellite through an earth shadow calculation method; specifically, the relative positions of the sun and the earth in the space and the spatial positions of the satellites are obtained by inputting time according to the time description system of step S1 and calculating, then the relative positions of the sun and the earth in the space and the spatial positions of the satellites are converted into the spatial position description system (J2000 geocentric inertial coordinate system) of step S2, and the sun exposure state of the satellites is obtained by the earth shadow calculation method.
Specifically, the earth shadow calculation method includes:
step T1, calculating sun parallel light ray prevention vector v 1 ；
Step T2, calculating a direction vector v from the geocentric to the satellite 2 ；
Step T3, calculating v 1 And v 2 Angle theta therebetween 1 ；
Step T4, calculating a critical angle theta 2 Wherein the critical angle refers to the minimum value of the earth shadow angle in a circular orbit with the distance from the satellite to the geocenter as a radius;
step T5, if theta 1 ＞θ 2 The satellite is in the sunned state if theta 1 ＜θ 2 If the satellite is in a non-sunned state;
fig. 5 is a schematic diagram of a method for calculating earth shadow.
In step S6, as shown in fig. 6, the sun exposure state of the satellite in a specific time period can be obtained by setting the task start time and the task end time, and calculation results with different accuracies can be obtained by using different simulation step sizes, where the smaller the simulation step size, the higher the accuracy, the larger the calculation amount, the larger the simulation step size, the lower the accuracy, but the calculation amount is also reduced, and the simulation step size is optimally set to 20S.
For convenience of use of a user, UTC time is adopted for task starting time and task ending time, a description form of gregorian year, month, day, time, minute and second is adopted, the UTC time input during simulation calculation is converted into julian day, the use time in the simulation calculation process can be guaranteed to be continuous real number, calculation and solving are facilitated, from the task starting time, simulation step length is sequentially accumulated to serve as simulation calculation time, the sun exposure state of the satellite is calculated by the method obtained in the step S5, and calculation is stopped until the time accumulated value task ending time.
The invention will be further illustrated with reference to the following specific examples. It should be understood that these examples are for illustrative purposes only and are not intended to limit the scope of the present invention. The following examples are examples of experimental procedures not specified under specific conditions, generally according to the conditions recommended by the manufacturer.
1.1, establishing a time description system, wherein UTC is adopted as a user description standard, and a julian day is adopted for internal simulation calculation;
1.3, establishing a spatial position description system, and using a J2000 geocentric inertial coordinate system for simulation calculation by adopting a Cartesian rectangular coordinate system;
1.3, establishing a space planet position calculation method based on JPL-DE405 to obtain the relative positions of the sun and the earth in the universe space;
1.4, establishing a satellite orbit extrapolation calculation model, defining a satellite orbit by using six Kepler orbits, and calculating the spatial position of the satellite at any moment, wherein the height of the satellite flight orbit is 500 kilometers of a near earth orbit, the orbit inclination angle is 0 degree, the eccentricity is 0 degree, the ascent point right ascent is 0, the true near point angle is 0, and the near point angular distance is 0;
1.5, the task starting time is 2017, month 1, day 12, 00, the task ending time is 2017, month 1, day 13, 51; the earth shadow calculation method comprises the following steps:
step T1, calculating sun parallel ray prevention vector v 1 ；
Step T2, calculating a direction vector v from the geocentric to the satellite 2 ；
Step T3, calculating v 1 And v 2 Angle theta therebetween 1 ；
Step T4, calculating a critical angle theta 2 Wherein the critical angle refers to the minimum value of the earth shadow angle in a circular orbit with the distance from the satellite to the geocenter as a radius;
step T5, if theta 1 ＞θ 2 The satellite is in the sunned state if theta 1 ＜θ 2 Then the satellite is in a non-sunned state.
Fig. 7 is a diagram of the exposure period change of a satellite, in fig. 7, the ordinate shows that the satellite is in different exposure states, and the abscissa shows that the exposure body of the satellite changes continuously with the change of time in the operation process at different times.
In the embodiment of the invention, the orbital inclination angle of the satellite is 0, and the satellite subsatellite point has no latitude change, so that the space change condition of the satellite around the earth can be known by observing the geographical longitude change of the satellite subsatellite point, as shown in fig. 8, the ordinate represents the longitude of the satellite subsatellite point, and the abscissa corresponds to different moments.
Although the present disclosure has been described above, the scope of the present disclosure is not limited thereto. Various changes and modifications may be effected therein by one of ordinary skill in the pertinent art without departing from the spirit and scope of the present disclosure, and these changes and modifications are intended to be within the scope of the present disclosure.
Claims (10)
1. A simulation calculation method for the on-orbit running real-time solarization state of a satellite is characterized by comprising the following steps:
s1, establishing a time description system, wherein the coordinated world time is used as a user description standard, and a julian day is used for internal simulation calculation;
s2, establishing a spatial position description system, and using a J2000 geocentric inertial coordinate system for simulation calculation by adopting a Cartesian rectangular coordinate system;
s3, establishing a space star position calculation method based on ephemeris to obtain the relative positions of the sun and the earth in a space;
s4, establishing a satellite orbit extrapolation calculation model, defining a satellite orbit by using six Kepler orbit roots, and calculating the space position of the satellite at any moment;
s5, inputting time according to the time description system in the step S1, calculating to obtain the relative positions of the sun and the earth in the space and the space position of the satellite, then converting into the space position description system established in the step S2, and obtaining the sun exposure state of the satellite through an earth shadow calculation method;
and S6, setting task starting time, task ending time and simulation step length, sequentially accumulating the simulation step length from the task starting time as simulation calculation time, and calculating the sun exposure state of the satellite according to the method of the step S5 until the sun exposure state is accumulated to the task ending time.
2. The method for simulation calculation of the in-orbit running real-time sunned state of the satellite according to claim 1, wherein in the step S1, the user adopts the coordinated universal time as a description standard in the form of year, month, day, hour, minute and second of a gregorian calendar, and the coordinated universal time is converted into julian day in the simulation calculation process.
3. The method for simulation calculation of the in-orbit running real-time sun exposure of the satellite according to claim 1, wherein in the step S2, the origin of coordinates of the J2000 centroid inertial coordinate system coincides with the earth centroid, the X axis of the J2000 centroid inertial coordinate system points to the J2000 time division flat spring point, the Z axis of the J2000 centroid inertial coordinate system points to the north pole, and the Y axis, the X axis and the Z axis of the J2000 centroid inertial coordinate system form a right-hand rectangular coordinate system.
4. The method for simulating and calculating the in-orbit real-time sun exposure state of the satellite according to claim 1, wherein in the step S3, JPL ephemeris data is adopted to establish a space and planet position calculation method based on ephemeris, and the relative positions of the sun and the earth in the space are obtained.
5. The simulation calculation method for the in-orbit running real-time sun-exposed state of the satellite according to claim 4, wherein the ephemeris data in DE405 ephemeris is adopted to establish a space celestial position calculation method based on the ephemeris, so as to obtain the relative positions of the sun and the earth in the space; the DE405 ephemeris adopts a coordinate system which is a rectangular coordinate system with the sun system centroid as the origin, the J2000 earth equatorial plane as the xy plane, and the J2000 vernal equinox direction as the x direction.
6. The method for simulation calculation of the in-orbit running real-time sun exposure state of the satellite according to claim 1, wherein in the step S4, the six kepler orbit components include a semi-major axis a, an eccentricity e, an inclination angle i, a ascension threshold angle Ω, an argument ω of near place, and a true argument f of near point.
7. The method for simulating and calculating the in-orbit real-time sun exposure state of the satellite according to claim 6, wherein the semi-major axis a, the eccentricity e, the inclination angle i, the ascent intersection akathisia Ω, the perigee argument ω and the true perigee angle f are calculated by the following formulas (1) to (6):
wherein μ is an earth gravity constant, μ = GM =3.986005 × 10 14 m 3 /s 2 G is a gravitational constant, and M is the earth mass; let t 0 Time of day, in the inertial system, the velocity vector of the satellite
And position vector
Are respectively as
And
the magnitude of velocity v 0 And the position size r 0 Are respectively as
And
momentum moment of the orbit of
And
8. the method according to claim 1, wherein in step S4, the SGP4 model is used to calculate satellite orbit data at any time by using six initialized kepler orbit numbers and time, so as to obtain the spatial position of the satellite at any time.
9. The method for simulating and calculating the in-orbit real-time sun exposure of the satellite according to claim 1, wherein in the step S5, the method for calculating the earth shadow comprises the following steps:
step T1, calculating sun parallel ray prevention vector v 1 ；
Step T2, calculating a direction vector v from the geocentric to the satellite 2 ；
Step T3, calculating v 1 And v 2 Angle theta therebetween 1 ；
Step T4, calculating a critical angle theta 2 Wherein the critical angle refers to the minimum value of the earth shadow angle in a circular orbit with the distance from the satellite to the geocenter as a radius;
step T5, if theta 1 ＞θ 2 The satellite is in the sunned state if theta 1 ＜θ 2 Then the satellite is in a non-sunned state.
10. The method for calculating the simulation of the in-orbit running real-time sun exposure of the satellite according to claim 1, wherein in the step S6, the simulation step size is 20S.
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| CN116151052A (en) * | 2023-04-24 | 2023-05-23 | 中国人民解放军32035部队 | Method and device for heavily orbit-changing losing recapture of certain satellite and electronic equipment |
| CN116796522A (en) * | 2023-06-05 | 2023-09-22 | 中国人民解放军战略支援部队航天工程大学 | Satellite data processing method, device, equipment and storage medium |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116151052A (en) * | 2023-04-24 | 2023-05-23 | 中国人民解放军32035部队 | Method and device for heavily orbit-changing losing recapture of certain satellite and electronic equipment |
| CN116151052B (en) * | 2023-04-24 | 2023-07-04 | 中国人民解放军32035部队 | Method and device for heavily orbit-changing losing recapture of certain satellite and electronic equipment |
| CN116796522A (en) * | 2023-06-05 | 2023-09-22 | 中国人民解放军战略支援部队航天工程大学 | Satellite data processing method, device, equipment and storage medium |
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