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 PDFInfo
- Publication number
- 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
- Authority
- CN
- China
- Prior art keywords
- satellite
- sun
- flight
- vector
- solar
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 238000005286 illumination Methods 0.000 title claims abstract description 56
- 238000000034 method Methods 0.000 title claims abstract description 30
- 238000013461 design Methods 0.000 title claims abstract description 25
- 238000005457 optimization Methods 0.000 title claims abstract description 14
- 239000013598 vector Substances 0.000 claims abstract description 119
- 230000008859 change Effects 0.000 claims abstract description 42
- 238000005259 measurement Methods 0.000 claims abstract description 5
- 238000013459 approach Methods 0.000 claims description 16
- 230000017525 heat dissipation Effects 0.000 claims description 15
- 230000014509 gene expression Effects 0.000 claims description 11
- 238000006243 chemical reaction Methods 0.000 claims description 8
- 238000005096 rolling process Methods 0.000 claims description 8
- 230000007246 mechanism Effects 0.000 claims description 5
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 claims description 4
- 238000003384 imaging method Methods 0.000 claims description 4
- 238000009434 installation Methods 0.000 claims description 4
- 239000011159 matrix material Substances 0.000 claims description 4
- 241000127225 Enceliopsis nudicaulis Species 0.000 claims 1
- 230000009286 beneficial effect Effects 0.000 abstract description 2
- 238000011161 development Methods 0.000 description 4
- 241001424688 Enceliopsis Species 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 230000004048 modification Effects 0.000 description 3
- 238000001514 detection method Methods 0.000 description 2
- 230000007547 defect Effects 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/223—Modular spacecraft systems
Landscapes
- Engineering & Computer Science (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Physics & Mathematics (AREA)
- Astronomy & Astrophysics (AREA)
- General Physics & Mathematics (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
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
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
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;
wherein, the expressions of Ps and Qs are as follows:
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 conversionIs represented by (a);
the solar incident direction angle θ can be obtained according to the following expression:
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;
wherein r is the radius of the light spot relative to the calibration origin of the detector;
it should be noted that the sun vectorFor 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 modeIs 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 sensorObtaining sun vector in shadow region by recursion of sun vector only from on-satellite orbit parameter
Preferably, the calculating step comprises:
the sun altitude angle beta is defined as the earth to sun vectorThe included angle between the track surface and the track surface can be obtained according to the following formula:
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:
wherein,obtaining satellite inertial system lower position vectors for the satellite orbit parameters,is the sun vector, dot represents the vectorSum vectorDot 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
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;
wherein, the expressions of Ps and Qs are as follows:
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 conversionIs represented by (a);
the solar incident direction angle θ can be obtained according to the following expression:
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;
wherein r is the radius of the light spot relative to the calibration origin of the detector;
it should be noted that the sun vectorFor 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 modeIs 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 sensorObtaining sun vector in shadow region by recursion of sun vector only from on-satellite orbit parameter
Specifically, the calculating step includes:
the sun altitude angle beta is defined as the earth to sun vectorThe included angle between the track surface and the track surface can be obtained according to the following formula:
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:
wherein,obtaining satellite inertial system lower position vectors for the satellite orbit parameters,is the sun vector, dot represents the vectorSum vectorDot 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
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.
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 conversionIs shown.
The solar incident direction angle θ can be obtained according to the following expression:
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.
and r is the radius of the light spot relative to the calibration origin of the detector.
It should be noted that the sun vectorFor 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 modeIs shown.
In the illumination areas, the sun vectors can be obtainedIn shadow region, only the sun vector is obtained by
The sun altitude angle beta is defined as the vector from the earth to the sunThe included angle between the track surface and the track surface can be obtained according to the following formula:
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:
obtaining satellite inertial system lower position vectors for the satellite orbit parameters,is the sun vector, dot represents the vectorSum vectorDot 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
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;
wherein,
the expressions Ps and Qs are as follows:
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 conversionIs represented by (a);
the solar incident direction angle θ can be obtained according to the following expression:
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;
wherein r is the radius of the light spot relative to the calibration origin of the detector;
4. The method of claim 1, wherein the calculating step comprises:
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:
wherein,obtaining satellite inertial system lower position vectors for the satellite orbit parameters,is the sun vector, dot represents the vectorSum vectorDot 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
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.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN202010037946.0A CN111232246B (en) | 2020-01-14 | 2020-01-14 | Overall optimization design method based on illumination conditions of inclined orbit satellite |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN202010037946.0A CN111232246B (en) | 2020-01-14 | 2020-01-14 | Overall optimization design method based on illumination conditions of inclined orbit satellite |
Publications (2)
Publication Number | Publication Date |
---|---|
CN111232246A CN111232246A (en) | 2020-06-05 |
CN111232246B true CN111232246B (en) | 2021-05-28 |
Family
ID=70873143
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
CN202010037946.0A Active CN111232246B (en) | 2020-01-14 | 2020-01-14 | Overall optimization design method based on illumination conditions of inclined orbit satellite |
Country Status (1)
Country | Link |
---|---|
CN (1) | CN111232246B (en) |
Families Citing this family (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN112857306B (en) * | 2020-12-31 | 2022-12-13 | 航天东方红卫星有限公司 | Method for determining continuous solar altitude angle of video satellite at any view direction point |
CN112896554A (en) * | 2021-02-18 | 2021-06-04 | 航天科工空间工程发展有限公司 | Satellite attitude control method |
CN114715430B (en) * | 2021-03-31 | 2022-11-08 | 中国科学院国家空间科学中心 | System for multi-satellite automatic linear formation and time-varying baseline generation |
CN115180179A (en) * | 2022-06-22 | 2022-10-14 | 中国航天空气动力技术研究院 | Self-leveling low-orbit satellite solar wing pneumatic layout |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP1059231A2 (en) * | 1999-06-08 | 2000-12-13 | Space Systems / Loral, Inc. | Solar array control for a satellite with an electric propulsion system |
CN104090612A (en) * | 2014-07-08 | 2014-10-08 | 上海新跃仪表厂 | Inclined orbit spacecraft energy obtaining method based on yaw steering |
CN107085634A (en) * | 2017-04-12 | 2017-08-22 | 上海航天控制技术研究所 | It is quick to calculate sunshine and the method for sun synchronous satellite star sensor minimum angle |
CN108791955A (en) * | 2018-06-14 | 2018-11-13 | 上海卫星工程研究所 | Static remote sensing satellite camera sun bypassing method |
CN110450980A (en) * | 2019-08-14 | 2019-11-15 | 上海卫星工程研究所 | Satellite solar battery array closed loop is to day tracking and its tracking system |
-
2020
- 2020-01-14 CN CN202010037946.0A patent/CN111232246B/en active Active
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP1059231A2 (en) * | 1999-06-08 | 2000-12-13 | Space Systems / Loral, Inc. | Solar array control for a satellite with an electric propulsion system |
CN104090612A (en) * | 2014-07-08 | 2014-10-08 | 上海新跃仪表厂 | Inclined orbit spacecraft energy obtaining method based on yaw steering |
CN107085634A (en) * | 2017-04-12 | 2017-08-22 | 上海航天控制技术研究所 | It is quick to calculate sunshine and the method for sun synchronous satellite star sensor minimum angle |
CN108791955A (en) * | 2018-06-14 | 2018-11-13 | 上海卫星工程研究所 | Static remote sensing satellite camera sun bypassing method |
CN110450980A (en) * | 2019-08-14 | 2019-11-15 | 上海卫星工程研究所 | Satellite solar battery array closed loop is to day tracking and its tracking system |
Non-Patent Citations (1)
Title |
---|
倾斜轨道航天器太阳翼对日跟踪方法探讨;王颖等;《航天器工程》;20090531;第18卷(第3期);第36-40页 * |
Also Published As
Publication number | Publication date |
---|---|
CN111232246A (en) | 2020-06-05 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
CN111232246B (en) | Overall optimization design method based on illumination conditions of inclined orbit satellite | |
CN111427002B (en) | Azimuth angle calculation method for ground measurement and control antenna pointing satellite | |
US7221317B2 (en) | Space-based lever arm correction in navigational systems employing spot beams | |
CN111897357A (en) | Attitude tracking control method for satellite earth scanning | |
CN110555250B (en) | Method for determining optimal bias angle of non-sun-oriented solar cell array | |
CN113641182B (en) | High-precision aiming and pointing method and system for inter-satellite laser communication system | |
US7221316B2 (en) | Control segment-based lever-arm correction via curve fitting for high accuracy navigation | |
Wolf et al. | Toward improved landing precision on Mars | |
CN112319857B (en) | Combined attitude control method and system for remote distributed satellite | |
CN111505608B (en) | Laser pointing on-orbit calibration method based on satellite-borne laser single-chip footprint image | |
CN110104210B (en) | Multi-satellite-sensitivity layout method for low-orbit sun-tracking satellite | |
CN110641741A (en) | Double-freedom-degree solar panel control method and control system thereof | |
Yoo et al. | Solar tracking system experimental verification based on GPS and vision sensor fusion | |
CN111121788B (en) | Spacecraft attitude singularity determination method and system based on double-vector attitude reference | |
Pensado et al. | Deep Learning-Based Target Pose Estimation Using LiDAR Measurements in Active Debris Removal Operations | |
CN112977889B (en) | Satellite attitude capturing method based on sun sensor and earth sensor | |
CN112925708B (en) | Static orbit microwave star load and platform collaborative scanning imaging simulation method and system | |
Wang et al. | Optimization method for star tracker orientation in the sun-pointing mode | |
CN114802818A (en) | Morning and evening orbit satellite and sun attitude calculation method and guidance method thereof | |
CN117742389B (en) | Heliostat cleaning posture calibration method based on image recognition technology | |
CN115320891B (en) | Near-circle nominal orbit control method based on virtual satellite | |
Zahran et al. | A solar cell based coarse sun sensor for a small leo satellite attitude determination | |
CN116295018B (en) | Target pose measurement method and system | |
CN117538958A (en) | Static track microwave detector electric axis pointing on-orbit calibration and correction method and system | |
Rao et al. | Star tracker alignment determination for resourcesat-I |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PB01 | Publication | ||
PB01 | Publication | ||
SE01 | Entry into force of request for substantive examination | ||
SE01 | Entry into force of request for substantive examination | ||
GR01 | Patent grant | ||
GR01 | Patent grant |