CN111232246B - Overall optimization design method based on illumination conditions of inclined orbit satellite - Google Patents

Overall optimization design method based on illumination conditions of inclined orbit satellite Download PDF

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CN111232246B
CN111232246B CN202010037946.0A CN202010037946A CN111232246B CN 111232246 B CN111232246 B CN 111232246B CN 202010037946 A CN202010037946 A CN 202010037946A CN 111232246 B CN111232246 B CN 111232246B
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CN111232246A (en
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许海玉
汪自军
***
缪鹏飞
王震
朱海江
王君磊
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Shanghai Institute of Satellite Engineering
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Abstract

The invention provides a total optimization design method based on illumination conditions of an inclined orbit satellite, which comprises the following steps: a calculation step: obtaining a sun vector by satellite orbit parameter recursion or sun sensor measurement, and obtaining a sun altitude angle by calculation, wherein the whole satellite energy supply change and the radiating surface change are represented by taking the sun altitude angle as a parameter; determining a flight scheme: the satellite U-turn flight scheme which takes the whole satellite energy and the fixed radiating surface into consideration; determining the flight polarity of the satellite: determining the flight polarity of the satellite according to the change direction of the sun vector; and determining applicability: adaptability of shadow area judging method and shadow area turning flight scheme; according to the selection step: and selecting basis for turning around flight time. The invention starts from the overall optimization design of the whole satellite, and is beneficial to the thermal control design of the whole satellite; enough illumination time is ensured, and the whole satellite energy is balanced; the fixed lighting surface also solves the constraint of sensor and load layout design.

Description

Overall optimization design method based on illumination conditions of inclined orbit satellite
Technical Field
The invention relates to the field of satellite overall optimization design methods, in particular to an overall optimization design method based on illumination conditions of an inclined orbit satellite.
Background
With the continuous development of the aerospace technology, higher requirements are put forward on satellite detection areas, the south-north low-latitude areas gradually become hot spot areas for detection, and the development requirements of inclined orbit satellites are more and more for meeting the requirement of high-time revisit under the condition of less satellite quantity.
The inclined orbit satellite is characterized in that the included angle between a sun vector and an orbit plane is constantly changed, the illumination condition of the whole satellite is complex, the satellite body does not have a fixed radiating surface, the thermal control design of the whole satellite is difficult, the illumination time obtained by the traditional fixed wing single-shaft one-dimensional driven solar sailboard is short due to the change of the illumination condition of the whole satellite, the requirement of energy supply cannot be met, the energy of the whole satellite cannot be balanced, the driving mechanism of the non-fixed wing two-dimensional driven solar sailboard is complex, the reliability is low, and the development and development stage is still in progress. In addition, the whole star layout is also limited, the attitude sensor is often influenced by solar illumination to fail, and the remote sensing load cannot be imaged. In summary, the inclined orbit satellite has challenges in thermal control subsystems, power subsystems, mechanical layout, and the like.
The invention provides a U-turn flight scheme taking the solar altitude as a parameter basis from the overall optimization design angle of the whole satellite, so that the satellite has a fixed illumination surface and a radiating surface, and the thermal control design of the whole satellite is facilitated; enough illumination time is ensured, and the whole satellite energy is balanced; the fixed illumination surface also solves the constraint of the attitude sensor and the load layout design.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide an overall optimization design method based on the illumination conditions of an inclined orbit satellite.
The invention provides a total optimization design method based on illumination conditions of an inclined orbit satellite, which is characterized by comprising the following steps:
a calculation step: obtaining a sun vector by satellite orbit parameter recursion or sun sensor measurement, and obtaining a sun altitude angle by calculation, wherein the whole satellite energy supply change and the radiating surface change are represented by taking the sun altitude angle as a parameter;
determining a flight scheme: the satellite U-turn flight scheme which takes the whole satellite energy and the fixed radiating surface into consideration;
determining the flight polarity of the satellite: determining the flight polarity of the satellite according to the change direction of the sun vector;
and determining applicability: adaptability of shadow area judging method and shadow area turning flight scheme;
according to the selection step: and selecting basis for turning around flight time.
Preferably, the calculating step comprises:
orbit parameters recursion sun vector on satellite: calculating the number of the julian century, wherein the relative epoch J2000 unit time is T, and obtaining the vector of the sun under an inertial system through 6 solar orbits
Figure BDA0002366698240000021
as＝149598022.589827
es＝0.01670862-0.00004204*T-0.00000124*T2
is＝0.409146133727251-0.000226965524811429*T-2.86040071854626e-009*T2
ws＝4.93818828588621+0.0300115182644598*T+0.00045972*T2
Ms＝6.24005996669206+628.301955132127*T-0.00980875039620813*T2
Wherein as is the solar orbit height, es is the eccentricity, is the inclination angle, ws is the perigee angular distance, and Ms is the mean perigee angle;
Es＝Ms+(es-es3/8)*sin(Ms)+1/2*es2*sin(2*Ms)+3/8*es3*sin(3*Ms)
wherein Es is a near point angle;
Figure BDA0002366698240000022
wherein,
Figure BDA0002366698240000023
is the sun vector;
wherein, the expressions of Ps and Qs are as follows:
Figure BDA0002366698240000024
Figure BDA0002366698240000025
preferably, the calculating step comprises:
the sun sensor calculates to obtain a sun vector: the coordinate value of the solar facula on the APS detector is used for obtaining the high and low angles of the sun sensor relative to the incident ray, namely the expression of the sun ray vector under the coordinates of the sun sensor, and the sun vector under the inertial system is obtained through coordinate conversion
Figure BDA0002366698240000026
Is represented by (a);
the solar incident direction angle θ can be obtained according to the following expression:
Figure BDA0002366698240000031
wherein: x is the number ofi，yiAs coordinates of the spot on the detector, x0，y0As coordinates of the center origin of the detector, fiThe distance from the sun sensor diaphragm to the APS image detector imaging surface is obtained;
sun vector in space-sensitive coordinate system
Figure BDA0002366698240000032
Is expressed as:
Figure BDA0002366698240000033
wherein r is the radius of the light spot relative to the calibration origin of the detector;
it should be noted that the sun vector
Figure BDA0002366698240000034
For the representation of the sun sensor in a single-machine coordinate system, the sun vector in an inertial system is obtained through a coordinate conversion matrix according to a single-machine installation mode
Figure BDA0002366698240000035
Is shown.
Preferably, the calculating step comprises:
the sun vector can be obtained by recursion of the orbit parameters on the star in the illumination area and calculation of the sun vector obtained by the sun sensor
Figure BDA0002366698240000036
Obtaining sun vector in shadow region by recursion of sun vector only from on-satellite orbit parameter
Figure BDA0002366698240000037
Preferably, the calculating step comprises:
the sun altitude angle beta is defined as the earth to sun vector
Figure BDA0002366698240000038
The included angle between the track surface and the track surface can be obtained according to the following formula:
Figure BDA0002366698240000039
wherein,
Figure BDA00023666982400000310
is an orbital momentum vector.
Preferably, the calculating step comprises:
the solar altitude represents the whole satellite supply change and the radiating surface change as parameters: defining a satellite body coordinate system ObXbYbZbThe coordinate system is fixedly connected with the star body and is a rectangular coordinate system; xbAxis, YbAxis, ZbThe axes are three geometric axes of the satellite; at nominal attitude, XbThe axis is a rolling axis and is consistent with the flight direction of the satellite, namely positive flight time + XbThe axis is in accordance with the flight direction, inverted flight time-XbThe axis is in line with the flight direction; zbThe axis being the yaw axis, pointing towards the centre of the earth, i.e. + ZbThe axis points to the earth center; y isbThe axis being the pitch axis, the star being formed by the right-hand rule, i.e. perpendicular to the orbital plane and in the direction of the negative normalthe-Y and + Y planes respectively refer to the body axis-Yb，+YbA corresponding plane;
when the solar altitude angle approaches 90 degrees, the solar panel has good illumination condition and the whole satellite has sufficient energy; when the solar altitude approaches 0 degree, the illumination condition of the solar sailboard is deteriorated, and the whole satellite energy source is insufficient; when the solar altitude is close to minus 90 degrees, the solar sailboard has no illumination, and the whole satellite energy source is insufficient; when the solar altitude changes from 90 degrees to → 0 degrees, the star-Y surface is a radiating surface; when the solar altitude changes from 0 DEG → 90 DEG, the star body + Y surface is a heat dissipation surface.
Preferably, the step of determining a flight plan comprises:
the design of the satellite U-turn flight scheme considering the whole satellite energy and the fixed radiating surface is as follows: when the solar altitude is near 0 degree, the satellite performs turning flight control through the attitude and orbit control subsystem attitude actuating mechanism, and after the satellite turns and flies, the rolling axis-X of the satellite bodybAxial in the direction of flight, star yaw + ZbPointing at the earth; after the satellite turns around and flies, when the solar altitude approaches 90 degrees, the solar sailboard has good illumination condition and the whole satellite has sufficient energy; when the solar altitude is close to 0 degree, the illumination condition of the solar sailboard gradually becomes better from poor along with the change of the solar altitude; when the solar altitude angle approaches minus 90 degrees, the solar sailboard is changed from no illumination to illumination, and the whole satellite has sufficient energy; before the satellite turns around and flies, when the solar altitude angle changes from 90 degrees to → 0 degrees, the star body and the Y surface are used as heat dissipation surfaces, and after the satellite turns around and flies, when the solar altitude angle changes from 0 degrees to 90 degrees, the star body and the Y surface can still be used as heat dissipation surfaces.
Preferably, the step of determining the satellite flight polarity comprises:
determining the flight polarity of the satellite according to the change direction of the sun vector: the constraint condition of the turning flight of the satellite in the orbit flight period is the solar altitude, the solar altitude is a slow variable relative to the orbital motion of the satellite, the solar altitude can be turned to fly after changing within a period of time, and the judgment of the change direction of the solar vector is increased. When the sun vector is above the orbital plane, namely when the sun vector is consistent with the normal direction of the orbital plane, and the sun altitude changes from 90 degrees to → 0 degrees, the satellite flies in a U-turn manner; when the sun vector is below the orbital plane, i.e., when the sun vector is opposite to the normal direction of the orbital plane, and the sun altitude changes from 0 ° → -90 °, the satellite returns to the original flight state.
Preferably, the determining suitability step comprises:
shadow area judging method and shadow area turning flight scheme:
and calculating an included angle alpha between the satellite, the earth and the sun, wherein the relation of alpha is calculated as follows:
Figure BDA0002366698240000041
wherein,
Figure BDA0002366698240000042
obtaining satellite inertial system lower position vectors for the satellite orbit parameters,
Figure BDA0002366698240000043
is the sun vector, dot represents the vector
Figure BDA0002366698240000044
Sum vector
Figure BDA0002366698240000045
Dot product of (1);
calculating the judgment angle x, wherein the relation of x is as follows:
χ＝acos(r0/(r0+a0))+π/2
r0is the radius of the earth, a0For track height, π is typically 3.1415926;
such as satellite orbital altitude a0Is 510km, r0The radius of the earth is 6378.14km
χ＝acos(6378.14/(6378.14+510))+π/2
＝1.9580
The judgment basis is as follows: when the judgment angle x is larger than an included angle alpha between the star, the earth and the sun, the judgment angle x is an illumination area, and when the judgment angle x is smaller than the included angle alpha between the star, the earth and the sun, the judgment angle x is a shadow area;
shadow zone u-turn flight scheme: in the shadow area, the sun sensor can not obtain the sun vector by measuring, the sun vector can be obtained by the orbit parameter recursion on the satellite, and the influence of the orbit recursion error can be ignored in the short time of the shadow area; the solar altitude is still applicable to representing the whole satellite supply change and the heat dissipation surface change by taking the solar altitude as a parameter.
Preferably, the step of selecting according to comprises:
selecting the maneuvering time: in order to minimize the influence of the turning flight on the application task executed by the whole satellite, the turning flight time is selected in a south latitude area with relatively less entering task areas, and the maximum latitude is limited by a satellite orbit inclination angle i.
Compared with the prior art, the invention has the following beneficial effects:
the invention provides a U-turn flight scheme taking the solar altitude as a parameter basis from the overall optimization design angle of the whole satellite, so that the satellite has a fixed illumination surface and a radiating surface, and the thermal control design of the whole satellite is facilitated; enough illumination time is ensured, and the whole satellite energy is balanced; the fixed lighting surface also solves the constraint of sensor and load layout design.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
fig. 1 is a schematic view of a sun vector and a satellite flight direction provided by the present invention.
FIG. 2 is a schematic flow chart of the steps provided by the present invention.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
The invention provides a total optimization design method based on illumination conditions of an inclined orbit satellite, which is characterized by comprising the following steps:
a calculation step: obtaining a sun vector by satellite orbit parameter recursion or sun sensor measurement, and obtaining a sun altitude angle by calculation, wherein the whole satellite energy supply change and the radiating surface change are represented by taking the sun altitude angle as a parameter;
determining a flight scheme: the satellite U-turn flight scheme which takes the whole satellite energy and the fixed radiating surface into consideration;
determining the flight polarity of the satellite: determining the flight polarity of the satellite according to the change direction of the sun vector;
and determining applicability: adaptability of shadow area judging method and shadow area turning flight scheme;
according to the selection step: and selecting basis for turning around flight time.
Specifically, the calculating step includes:
orbit parameters recursion sun vector on satellite: calculating the number of the julian century, wherein the relative epoch J2000 unit time is T, and obtaining the vector of the sun under an inertial system through 6 solar orbits
Figure BDA0002366698240000068
as＝149598022.589827
es＝0.01670862-0.00004204*T-0.00000124*T2
is＝0.409146133727251-0.000226965524811429*T-2.86040071854626e-009*T2
ws＝4.93818828588621+0.0300115182644598*T+0.00045972*T2
Ms＝6.24005996669206+628.301955132127*T-0.00980875039620813*T2
Wherein as is the solar orbit height, es is the eccentricity, is the inclination angle, ws is the perigee angular distance, and Ms is the mean perigee angle;
Es＝Ms+(es-es3/8)*sin(Ms)+1/2*es2*sin(2*Ms)+3/8*es3*sin(3*Ms)
wherein Es is a near point angle;
Figure BDA0002366698240000061
wherein,
Figure BDA0002366698240000062
is the sun vector;
wherein, the expressions of Ps and Qs are as follows:
Figure BDA0002366698240000063
Figure BDA0002366698240000064
specifically, the calculating step includes:
the sun sensor calculates to obtain a sun vector: the coordinate value of the solar facula on the APS detector is used for obtaining the high and low angles of the sun sensor relative to the incident ray, namely the expression of the sun ray vector under the coordinates of the sun sensor, and the sun vector under the inertial system is obtained through coordinate conversion
Figure BDA0002366698240000065
Is represented by (a);
the solar incident direction angle θ can be obtained according to the following expression:
Figure BDA0002366698240000066
wherein: x is the number ofi，yiAs coordinates of the spot on the detector, x0，y0As coordinates of the center origin of the detector, fiThe distance from the sun sensor diaphragm to the APS image detector imaging surface is obtained;
sun vector in space-sensitive coordinate system
Figure BDA0002366698240000067
Is expressed as:
Figure BDA0002366698240000071
wherein r is the radius of the light spot relative to the calibration origin of the detector;
it should be noted that the sun vector
Figure BDA0002366698240000072
For the representation of the sun sensor in a single-machine coordinate system, the sun vector in an inertial system is obtained through a coordinate conversion matrix according to a single-machine installation mode
Figure BDA0002366698240000073
Is shown.
Specifically, the calculating step includes:
the sun vector can be obtained by recursion of the orbit parameters on the star in the illumination area and calculation of the sun vector obtained by the sun sensor
Figure BDA0002366698240000074
Obtaining sun vector in shadow region by recursion of sun vector only from on-satellite orbit parameter
Figure BDA0002366698240000075
Specifically, the calculating step includes:
the sun altitude angle beta is defined as the earth to sun vector
Figure BDA0002366698240000076
The included angle between the track surface and the track surface can be obtained according to the following formula:
Figure BDA0002366698240000077
wherein,
Figure BDA0002366698240000078
is an orbital momentum vector.
Specifically, the calculating step includes:
the solar altitude represents the whole satellite supply change and the radiating surface change as parameters: defining a satellite body coordinate system ObXbYbZbThe coordinate system is fixedly connected with the star body and is a rectangular coordinate system; xbAxis, YbAxis, ZbThe axes are three geometric axes of the satellite; at nominal attitude, XbThe axis is a rolling axis and is consistent with the flight direction of the satellite, namely positive flight time + XbThe axis is in accordance with the flight direction, inverted flight time-XbThe axis is in line with the flight direction; zbThe axis being the yaw axis, pointing towards the centre of the earth, i.e. + ZbThe axis points to the earth center; y isbThe axis is a pitch axis, and according to the right-hand rule, i.e. perpendicular to the orbital plane and along the direction of the negative normal, the star-Y and + Y planes respectively refer to the body axis-Yb，+YbA corresponding plane;
when the solar altitude angle approaches 90 degrees, the solar panel has good illumination condition and the whole satellite has sufficient energy; when the solar altitude approaches 0 degree, the illumination condition of the solar sailboard is deteriorated, and the whole satellite energy source is insufficient; when the solar altitude is close to minus 90 degrees, the solar sailboard has no illumination, and the whole satellite energy source is insufficient; when the solar altitude changes from 90 degrees to → 0 degrees, the star-Y surface is a radiating surface; when the solar altitude changes from 0 DEG → 90 DEG, the star body + Y surface is a heat dissipation surface.
Specifically, the step of determining the flight plan comprises:
the design of the satellite U-turn flight scheme considering the whole satellite energy and the fixed radiating surface is as follows: when the solar altitude is near 0 degree, the satellite performs turning flight control through the attitude and orbit control subsystem attitude actuating mechanism, and after the satellite turns and flies, the rolling axis-X of the satellite bodybAxial in the direction of flight, star yaw + ZbPointing at the earth; after the satellite turns around and flies, when the solar altitude approaches 90 degrees, the solar sailboard has good illumination condition and the whole satellite has sufficient energy; when the solar altitude is close to 0 degree, the illumination condition of the solar sailboard gradually becomes better from poor along with the change of the solar altitude; when the solar altitude angle approaches minus 90 degrees, the solar sailboard is changed from no illumination to illumination, and the whole satellite has sufficient energy; before the satellite turns around and flies, when the solar altitude changes from 90 degrees to → 0 degrees, the star body and the Y surface are used as heat dissipation surfaces, and the satelliteAfter turning around and flying, when the solar altitude changes from 0 DEG to 90 DEG, the star body and the Y surface can still be used as a radiating surface.
Specifically, the step of determining the flight polarity of the satellite comprises the following steps:
determining the flight polarity of the satellite according to the change direction of the sun vector: the constraint condition of the turning flight of the satellite in the orbit flight period is the solar altitude, the solar altitude is a slow variable relative to the orbital motion of the satellite, the solar altitude can be turned to fly after changing within a period of time, and the judgment of the change direction of the solar vector is increased. When the sun vector is above the orbital plane, namely when the sun vector is consistent with the normal direction of the orbital plane, and the sun altitude changes from 90 degrees to → 0 degrees, the satellite flies in a U-turn manner; when the sun vector is below the orbital plane, i.e., when the sun vector is opposite to the normal direction of the orbital plane, and the sun altitude changes from 0 ° → -90 °, the satellite returns to the original flight state.
Specifically, the determining the suitability step includes:
shadow area judging method and shadow area turning flight scheme:
and calculating an included angle alpha between the satellite, the earth and the sun, wherein the relation of alpha is calculated as follows:
Figure BDA0002366698240000081
wherein,
Figure BDA0002366698240000082
obtaining satellite inertial system lower position vectors for the satellite orbit parameters,
Figure BDA0002366698240000083
is the sun vector, dot represents the vector
Figure BDA0002366698240000084
Sum vector
Figure BDA0002366698240000085
Dot product of (1);
calculating the judgment angle x, wherein the relation of x is as follows:
χ＝acos(r0/(r0+a0))+π/2
r0is the radius of the earth, a0For track height, π is typically 3.1415926;
such as satellite orbital altitude a0Is 510km, r0The radius of the earth is 6378.14km
χ＝acos(6378.14/(6378.14+510))+π/2
＝1.9580
The judgment basis is as follows: when the judgment angle x is larger than an included angle alpha between the star, the earth and the sun, the judgment angle x is an illumination area, and when the judgment angle x is smaller than the included angle alpha between the star, the earth and the sun, the judgment angle x is a shadow area;
shadow zone u-turn flight scheme: in the shadow area, the sun sensor can not obtain the sun vector by measuring, the sun vector can be obtained by the orbit parameter recursion on the satellite, and the influence of the orbit recursion error can be ignored in the short time of the shadow area; the solar altitude is still applicable to representing the whole satellite supply change and the heat dissipation surface change by taking the solar altitude as a parameter.
Specifically, the step of selecting according to the selection comprises:
selecting the maneuvering time: in order to minimize the influence of the turning flight on the application task executed by the whole satellite, the turning flight time is selected in a south latitude area with relatively less entering task areas, and the maximum latitude is limited by a satellite orbit inclination angle i.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
The present invention will be described more specifically below with reference to preferred examples.
Preferred example 1:
as shown in fig. 2, the present invention provides a general optimization design method based on illumination conditions of an inclined orbit satellite, comprising the following steps:
step 1: obtaining a sun vector by satellite orbit parameter recursion or sun sensor measurement, and obtaining a sun altitude angle by calculation, wherein the whole satellite energy supply change and the radiating surface change are represented by taking the sun altitude angle as a parameter;
step 2: the satellite U-turn flight scheme which takes the whole satellite energy and the fixed radiating surface into consideration;
and step 3: determining the flight polarity of the satellite according to the change direction of the sun vector;
and 4, step 4: adaptability of shadow area judging method and shadow area turning flight scheme;
and 5: and selecting basis for turning around flight time.
As shown in fig. 1, it is a schematic view of the sun vector and the satellite flight direction provided by the present invention;
the step 1 comprises the following steps:
firstly, recursion of solar vectors by orbit parameters on the satellite: calculating the number of the julian century, wherein the relative epoch J2000 unit time is T, and obtaining the vector of the sun under an inertial system through 6 solar orbits
Figure BDA0002366698240000092
as＝149598022.589827
es＝0.01670862-0.00004204*T-0.00000124*T2
is＝0.409146133727251-0.000226965524811429*T-2.86040071854626e-009*T2
ws＝4.93818828588621+0.0300115182644598*T+0.00045972*T2
Ms＝6.24005996669206+628.301955132127*T-0.00980875039620813*T2
Wherein as is the solar orbit height, es is the eccentricity, is the inclination angle, ws is the perigee angular distance, and Ms is the mean perigee angle.
Es＝Ms+(es-es3/8)*sin(Ms)+1/2*es2*sin(2*Ms)+3/8*es3*sin(3*Ms)
Es is a near point angle.
Figure BDA0002366698240000091
Wherein
Figure BDA0002366698240000101
Is a sun vector, wherein Ps and Qs are expressed as follows:
Figure BDA0002366698240000102
Figure BDA0002366698240000103
calculating by the sun sensor to obtain a sun vector: the coordinate value of the solar facula on the APS detector is used for obtaining the high and low angles of the sun sensor relative to the incident ray, namely the expression of the sun ray vector under the coordinates of the sun sensor, and the sun vector under the inertial system is obtained through coordinate conversion
Figure BDA0002366698240000104
Is shown.
The solar incident direction angle θ can be obtained according to the following expression:
Figure BDA0002366698240000105
wherein: x is the number ofi，yiAs coordinates of the spot on the detector, x0，y0As coordinates of the center origin of the detector, fiThe distance from the diaphragm of the sun sensor to the imaging surface of the APS image detector.
Sun vector in space-sensitive coordinate system
Figure BDA0002366698240000106
Is expressed as:
Figure BDA0002366698240000107
and r is the radius of the light spot relative to the calibration origin of the detector.
It should be noted that the sun vector
Figure BDA0002366698240000108
For the representation of the sun sensor in a single-machine coordinate system, the sun vector in an inertial system is obtained through a coordinate conversion matrix according to a single-machine installation mode
Figure BDA0002366698240000109
Is shown.
In the illumination areas, the sun vectors can be obtained
Figure BDA00023666982400001010
In shadow region, only the sun vector is obtained by
Figure BDA00023666982400001011
The sun altitude angle beta is defined as the vector from the earth to the sun
Figure BDA00023666982400001012
The included angle between the track surface and the track surface can be obtained according to the following formula:
Figure BDA00023666982400001013
wherein
Figure BDA00023666982400001014
Is an orbital momentum vector.
The solar altitude represents the whole satellite supply change and the heat dissipation surface change as parameters: defining a satellite body coordinate system ObXbYbZbThe coordinate system is fixedly connected with the star body and is a rectangular coordinate system. XbAxis, YbAxis, ZbThe axes are the three geometric axes of the satellite. At nominal attitude, XbThe axis is a rolling axis and is consistent with the flight direction of the satellite, namely positive flight time + XbThe axis is in accordance with the flight direction, inverted flight time-XbShaft andthe flight directions are consistent; zbThe axis being the yaw axis, pointing towards the centre of the earth, i.e. + ZbThe axis points to the earth center; y isbThe axis is a pitch axis, and according to the right-hand rule, i.e. perpendicular to the orbital plane and along the direction of the negative normal, the star-Y and + Y planes respectively refer to the body axis-Yb，+YbThe corresponding plane.
When the solar altitude angle approaches 90 degrees, the solar panel has good illumination condition and the whole satellite has sufficient energy; when the solar altitude approaches 0 degree, the illumination condition of the solar sailboard is deteriorated, and the whole satellite energy source is insufficient; when the solar altitude is close to minus 90 degrees, the solar sailboard has no illumination, and the whole satellite energy source is insufficient; when the solar altitude changes from 90 degrees to → 0 degrees, the star-Y surface is a radiating surface; when the solar altitude changes from 0 DEG → 90 DEG, the star body + Y surface is a heat dissipation surface.
The step 2 comprises the following steps:
the design of the satellite U-turn flight scheme which takes the whole satellite energy and the fixed radiating surface into consideration. When the solar altitude is near 0 degree, the satellite performs turning flight control through the attitude and orbit control subsystem attitude actuating mechanism, and after the satellite turns and flies, the rolling axis-X of the satellite bodybAxial in the direction of flight, star yaw + ZbPointing towards the earth. After the satellite turns around and flies, when the solar altitude approaches 90 degrees, the solar sailboard has good illumination condition and the whole satellite has sufficient energy; when the solar altitude is close to 0 degree, the illumination condition of the solar sailboard gradually becomes better from poor along with the change of the solar altitude; when the solar altitude angle approaches minus 90 degrees, the solar sailboard is changed from no illumination to illumination, and the whole satellite has sufficient energy; before the satellite turns around and flies, when the solar altitude angle changes from 90 degrees to → 0 degrees, the star body and the Y surface are used as heat dissipation surfaces, and after the satellite turns around and flies, when the solar altitude angle changes from 0 degrees to 90 degrees, the star body and the Y surface can still be used as heat dissipation surfaces.
The step 3 comprises the following steps:
and determining the flight polarity of the satellite according to the change direction of the sun vector. The constraint condition of the turning flight of the satellite in the orbit flight period is the solar altitude, the solar altitude is a slow variable relative to the orbital motion of the satellite, the solar altitude can be turned to fly after changing within a period of time, and the judgment of the change direction of the solar vector is increased. When the sun vector is above the orbital plane, namely when the sun vector is consistent with the normal direction of the orbital plane, and the sun altitude changes from 90 degrees to → 0 degrees, the satellite flies in a U-turn manner; when the sun vector is below the orbital plane, i.e., when the sun vector is opposite to the normal direction of the orbital plane, and the sun altitude changes from 0 ° → -90 °, the satellite returns to the original flight state.
The step 4 comprises the following steps:
shadow area judging method and shadow area turning flight scheme.
Calculating an included angle alpha between the satellite, the earth and the sun. The α relationship is calculated as follows:
Figure BDA0002366698240000111
Figure BDA0002366698240000112
obtaining satellite inertial system lower position vectors for the satellite orbit parameters,
Figure BDA0002366698240000113
is the sun vector, dot represents the vector
Figure BDA0002366698240000114
Sum vector
Figure BDA0002366698240000115
Dot product of (c).
② calculating the judgment angle χ. The χ relationship is calculated as follows:
χ＝acos(r0/(r0+a0))+π/2
r0is the radius of the earth, a0For track height, π is typically 3.1415926.
Such as satellite orbital altitude a0Is 510km, r0The radius of the earth is 6378.14km
χ＝acos(6378.14/(6378.14+510))+π/2
＝1.9580
Third, judgment basis
When the judgment angle x is larger than the included angle alpha between the star, the earth and the sun, the judgment angle x is an illumination area, and when the judgment angle x is smaller than the included angle alpha between the star, the earth and the sun, the judgment angle x is a shadow area.
Shadow zone turning flight scheme
In the shadow area, the sun vector cannot be obtained by the sun sensor in the step II in the claim 2, the sun vector can be obtained by the step I in the claim 2, and the influence of the orbit recursion error can be ignored in the short time of the shadow area. Step (iv) of claim 2 is still applicable.
The step 5 comprises the following steps:
and selecting the maneuvering time. In order to minimize the influence of the turning flight on the application task executed by the whole satellite, the turning flight time is selected in a south latitude area with relatively less entering task areas, and the maximum latitude is limited by a satellite orbit inclination angle i.
In the description of the present application, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience in describing the present application and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present application.
Those skilled in the art will appreciate that, in addition to implementing the systems, apparatus, and various modules thereof provided by the present invention in purely computer readable program code, the same procedures can be implemented entirely by logically programming method steps such that the systems, apparatus, and various modules thereof are provided in the form of logic gates, switches, application specific integrated circuits, programmable logic controllers, embedded microcontrollers and the like. Therefore, the system, the device and the modules thereof provided by the present invention can be considered as a hardware component, and the modules included in the system, the device and the modules thereof for implementing various programs can also be considered as structures in the hardware component; modules for performing various functions may also be considered to be both software programs for performing the methods and structures within hardware components.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.

Claims (9)

1. A total optimization design method based on illumination conditions of an inclined orbit satellite is characterized by comprising the following steps:
a calculation step: obtaining a sun vector by satellite orbit parameter recursion or sun sensor measurement, and obtaining a sun altitude angle by calculation, wherein the whole satellite energy supply change and the radiating surface change are represented by taking the sun altitude angle as a parameter;
determining a flight scheme: the satellite U-turn flight scheme which takes the whole satellite energy and the fixed radiating surface into consideration;
determining the flight polarity of the satellite: determining the flight polarity of the satellite according to the change direction of the sun vector;
and determining applicability: adaptability of shadow area judging method and shadow area turning flight scheme;
according to the selection step: selecting the time of turning around flight;
the calculating step includes:
the solar altitude represents the whole satellite supply change and the radiating surface change as parameters: defining a satellite body coordinate system ObXbYbZbThe coordinate system is fixedly connected with the star body and is a rectangular coordinate system; xbAxis, YbAxis, ZbThe axes are three geometric axes of the satellite; at nominal attitude, XbThe axis is a rolling axis and is consistent with the flight direction of the satellite, namely positive flight time + XbThe axis is in accordance with the flight direction, inverted flight time-XbThe axis is in line with the flight direction; zbThe axis being the yaw axis, pointing towards the centre of the earth, i.e. + ZbThe axis points to the earth center; y isbThe axis is a pitch axis, and according to the right-hand rule, i.e. perpendicular to the orbital plane and along the direction of the negative normal, the star-Y and + Y planes respectively refer to the body axis-Yb，+YbA corresponding plane;
when the solar altitude angle approaches 90 degrees, the solar panel has good illumination condition and the whole satellite has sufficient energy; when the solar altitude approaches 0 degree, the illumination condition of the solar sailboard is deteriorated, and the whole satellite energy source is insufficient; when the solar altitude is close to minus 90 degrees, the solar sailboard has no illumination, and the whole satellite energy source is insufficient; when the solar altitude changes from 90 degrees to 0 degrees, the star-Y surface is a radiating surface; when the solar altitude changes from 0 degree to-90 degrees, the star body and the Y surface are heat dissipation surfaces.
2. The method of claim 1, wherein the calculating step comprises:
orbit parameters recursion sun vector on satellite: calculating the number of the julian century, wherein the relative epoch J2000 unit time is T, and obtaining the vector of the sun under an inertial system through 6 solar orbits
Figure FDA0002990365470000011
as＝149598022.589827
es＝0.01670862-0.00004204*T-0.00000124*T2
is＝0.409146133727251-0.000226965524811429*T-2.86040071854626e-009*T2
ws＝4.93818828588621+0.0300115182644598*T+0.00045972*T2
Ms＝6.24005996669206+628.301955132127*T-0.00980875039620813*T2
Wherein as is the solar orbit height, es is the eccentricity, is the inclination angle, ws is the perigee angular distance, and Ms is the mean perigee angle;
Es＝Ms+(es-es3/8)*sin(Ms)+1/2*es2*sin(2*Ms)+3/8*es3*sin(3*Ms)
wherein Es is a near point angle;
Figure FDA0002990365470000021
wherein,
Figure FDA0002990365470000022
is the sun vector;
the expressions Ps and Qs are as follows:
Figure FDA0002990365470000023
Figure FDA0002990365470000024
3. the method of claim 1, wherein the calculating step comprises:
the sun sensor calculates to obtain a sun vector: the coordinate value of the solar facula on the APS detector is used for obtaining the high and low angles of the sun sensor relative to the incident ray, namely the expression of the sun ray vector under the coordinates of the sun sensor, and the sun vector under the inertial system is obtained through coordinate conversion
Figure FDA0002990365470000025
Is represented by (a);
the solar incident direction angle θ can be obtained according to the following expression:
Figure FDA0002990365470000026
wherein: x is the number ofi，yiAs coordinates of the spot on the detector, x0，y0As coordinates of the center origin of the detector, fiThe distance from the sun sensor diaphragm to the APS image detector imaging surface is obtained;
sun vector in space-sensitive coordinate system
Figure FDA0002990365470000027
Is expressed as:
Figure FDA0002990365470000028
wherein r is the radius of the light spot relative to the calibration origin of the detector;
it should be noted that the sun vector
Figure FDA0002990365470000029
For the representation of the sun sensor in a single-machine coordinate system, the sun vector in an inertial system is obtained through a coordinate conversion matrix according to a single-machine installation mode
Figure FDA00029903654700000210
Is shown.
4. The method of claim 1, wherein the calculating step comprises:
the sun vector can be obtained by recursion of the orbit parameters on the star in the illumination area and calculation of the sun vector obtained by the sun sensor
Figure FDA0002990365470000031
Obtaining sun vector in shadow region by recursion of sun vector only from on-satellite orbit parameter
Figure FDA0002990365470000032
5. The method of claim 1, wherein the calculating step comprises:
the sun altitude angle beta is defined as the earth to sun vector
Figure FDA0002990365470000033
The included angle between the track surface and the track surface can be obtained according to the following formula:
Figure FDA0002990365470000034
wherein,
Figure FDA0002990365470000035
is an orbital momentum vector.
6. The method of claim 1, wherein the step of determining the flight plan comprises:
the design of the satellite U-turn flight scheme considering the whole satellite energy and the fixed radiating surface is as follows: when the solar altitude is near 0 degree, the satellite performs turning flight control through the attitude and orbit control subsystem attitude actuating mechanism, and after the satellite turns and flies, the rolling axis-X of the satellite bodybAxial in the direction of flight, star yaw + ZbPointing at the earth; after the satellite turns around and flies, when the solar altitude approaches 90 degrees, the solar sailboard has good illumination condition and the whole satellite has sufficient energy; when the solar altitude is close to 0 degree, the illumination condition of the solar sailboard gradually becomes better from poor along with the change of the solar altitude; when the solar altitude angle approaches minus 90 degrees, the solar sailboard is changed from no illumination to illumination, and the whole satellite has sufficient energy; before the satellite turns around and flies, when the solar altitude is changed from 90 degrees to 0 degrees, the star body and the Y surface are used as heat dissipation surfaces, and after the satellite turns around and flies, when the solar altitude is changed from 0 degrees to-90 degrees, the star body and the Y surface can still be used as heat dissipation surfaces.
7. The method of claim 1, wherein the step of determining the flight polarity of the satellite comprises:
determining the flight polarity of the satellite according to the change direction of the sun vector: the constraint condition of the turning flight of the satellite in the orbit flight period is the solar altitude, the solar altitude is a slow variable relative to the orbital motion of the satellite, the solar altitude can be turned to fly after changing within a period of time, and the judgment of the change direction of the solar vector is increased; when the sun vector is above the orbital plane, namely when the sun vector is consistent with the normal direction of the orbital plane, and the sun altitude is changed from 90 degrees to 0 degrees, the satellite performs turning flight; when the sun vector is under the orbit surface, namely the sun vector is opposite to the normal direction of the orbit surface, and the sun altitude is changed from 0 degree to-90 degrees, the satellite recovers the original flight state.
8. The method of claim 1, wherein the determining applicability step comprises:
shadow area judging method and shadow area turning flight scheme:
and calculating an included angle alpha between the satellite, the earth and the sun, wherein the relation of alpha is calculated as follows:
Figure FDA0002990365470000041
wherein,
Figure FDA0002990365470000042
obtaining satellite inertial system lower position vectors for the satellite orbit parameters,
Figure FDA0002990365470000043
is the sun vector, dot represents the vector
Figure FDA0002990365470000044
Sum vector
Figure FDA0002990365470000045
Dot product of (1);
calculating the judgment angle x, wherein the relation of x is as follows:
χ＝arccos(r0/(r0+a0))+π/2
r0is the radius of the earth, a0For track height, π is typically 3.1415926;
such as satellite orbital altitude a0Is 510km, r0The radius of the earth is 6378.14km
Figure FDA0002990365470000046
The judgment basis is as follows: when the judgment angle x is larger than an included angle alpha between the star, the earth and the sun, the judgment angle x is an illumination area, and when the judgment angle x is smaller than the included angle alpha between the star, the earth and the sun, the judgment angle x is a shadow area;
shadow zone u-turn flight scheme: in the shadow area, the sun sensor can not obtain the sun vector by measuring, the sun vector can be obtained by the orbit parameter recursion on the satellite, and the influence of the orbit recursion error can be ignored in the short time of the shadow area; the solar altitude is still applicable to representing the whole satellite supply change and the heat dissipation surface change by taking the solar altitude as a parameter.
9. The method of claim 1, wherein the step of selecting according to the total optimization design based on the illumination condition of the inclined orbit satellite comprises:
selecting the maneuvering time: in order to minimize the influence of the turning flight on the application task executed by the whole satellite, the turning flight time is selected in a south latitude area with relatively less entering task areas, and the maximum latitude is limited by a satellite orbit inclination angle i.
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