CN109159925B - Satellite solar wing configuration design method capable of meeting load-to-sun observation requirements - Google Patents
Satellite solar wing configuration design method capable of meeting load-to-sun observation requirements Download PDFInfo
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- CN109159925B CN109159925B CN201810709164.XA CN201810709164A CN109159925B CN 109159925 B CN109159925 B CN 109159925B CN 201810709164 A CN201810709164 A CN 201810709164A CN 109159925 B CN109159925 B CN 109159925B
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- 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
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
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- 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
- B64G1/66—Arrangements or adaptations of apparatus or instruments, not otherwise provided for
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Abstract
A satellite solar wing configuration design method meeting load-to-sun observation requirements comprises the following steps: step 1, analyzing the field requirement of the load on daily observation, and determining the field envelope of the load according to the load working mode; step 2, analyzing the arrangement requirement of the in-orbit state of the solar wing, and determining the relative position relationship between the solar wing and the satellite body according to the change condition of the sun illumination angle; step 3, analyzing the limitation requirement of the carrier rocket on the overall dimension of the satellite, and determining the optional arrangement form of the load and the solar wing launching state by combining the overall state of the satellite body; and 4, embedding the load sun-facing observation view field into the solar wing according to the constraint conditions determined in the three steps, and defining the layout positions of the load and the solar wing on the satellite body to form the satellite solar wing configuration meeting the load sun-facing observation requirement. The invention breaks through the functional limitation that the solar wing is only used for supplying power to the satellite, and ensures that the solar wing has the function of allowing sunlight to pass through the solar wing so as to facilitate load sun observation.
Description
Technical Field
The invention relates to the technical field of satellite overall design, in particular to a satellite solar wing configuration design method meeting load sun observation requirements.
Background
With the development of satellite technology, the demand of users for solar activity monitoring on satellites is more and more urgent, so that the satellites are required to be loaded with earth observation loads and sun observation loads.
When the traditional satellite configuration design is carried out, the arrangement form of the solar wings is usually determined according to the characteristics of orbit illumination, then the layout design of the load is carried out, the configuration design cannot be carried out on the solar wings and the load in an overall mode, and the problems that the satellite layout is not compact enough, the longitudinal mass center is high, the weight is heavy and the like are often caused.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a satellite solar wing configuration design method meeting the requirement of load sun-sun observation, and the arrangement form of the load and the solar wing is considered in the overall design of the satellite configuration according to the characteristic that the sun-sun observation load and the solar wing are required to be illuminated by the sun.
In order to achieve the purpose, the technical scheme adopted by the invention is as follows:
a satellite solar wing configuration design method meeting load-to-sun observation requirements comprises the following steps:
and 4, embedding the load sun observation view field into the solar wing according to the constraint conditions determined in the three steps, and determining the layout positions of the load and the solar wing on the satellite body, so that the satellite solar wing configuration meeting the load sun observation requirement is formed.
Preferably, the load is a remote sensing camera loaded on a satellite for monitoring the change of the solar energy.
Preferably, the payload field envelope of step 1 is obtained by traversing the range of motion of the field at a certain position of the payload by the array.
Preferably, the specific method for determining the relative position relationship between the solar wing and the satellite body in the step 2 is as follows: according to the variation range of the solar illumination angle, the optimal value of the included angle between the solar wing in-orbit state cell and the satellite body is calculated, and the included angle between the normal of the solar wing in-orbit state cell and the satellite body is finally determined by combining the engineering feasibility of solar wing design and the energy requirement of the whole satellite.
Preferably, the specific method for determining the alternative arrangement form of the load in the step 3 is as follows: according to the requirement of a carrier rocket on the maximum envelope size of a satellite, all the arrangement position schemes of the sun-facing observation loads are given by combining the appearance characteristics of the satellite body and the on-orbit working requirement of the sun-facing observation loads to be measured for sun tracking, then the engineering feasibility of each scheme is analyzed one by one, and finally the optional arrangement position of the sun-facing observation loads is determined.
Preferably, the specific method for determining the alternative arrangement form of the solar wing emission state in the step 3 is as follows: according to the requirement of a carrier rocket on the maximum envelope size of a satellite, all solar wing arrangement position schemes are given by combining the appearance characteristics of a satellite body and the on-orbit working requirement that solar wing cells are irradiated by sunlight as perpendicularly as possible, then the engineering feasibility of each scheme is analyzed one by one, and the optional arrangement position of the solar wings is finally determined.
Preferably, in the step 4, various permutation and combination modes of the arrangement positions of the solar wings and the sun-facing observation loads are sequentially considered, the sun-facing direction requirements of the solar wings and the loads are comprehensively considered according to the principle that the height of the longitudinal center of mass of the whole satellite is reduced as much as possible and the available envelope space of the satellite is utilized as much as possible, the load-facing observation field of view is embedded into the solar wings, and the arrangement positions of the loads and the solar wings on the satellite body are determined.
Preferably, the satellite is provided with only one solar wing.
Compared with the prior art, the invention has the following beneficial effects:
1. the functional limitation that the solar wing is only used for supplying power to the satellite is broken through, so that the solar wing has the function of allowing sunlight to penetrate through the solar wing so as to facilitate load sun observation;
2. on the premise of meeting the requirement of load-to-day observation, the envelope space limited by carrying is fully utilized, and the longitudinal centroid height of the whole satellite is effectively reduced;
3. the method is simple and effective, and has wide application and popularization values.
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 field-of-view envelope for a counterglow observation load;
FIG. 2 is a schematic view of a field of view of a location of a load as observed daily;
FIG. 3 is a schematic diagram of the physical dimension constraint of a launch vehicle on a satellite;
FIG. 4 is a schematic view of the load field envelope and the relative position of the solar wings observed over the day;
FIG. 5 is a schematic view of a satellite launch state configuration;
fig. 6 is a schematic view of a satellite flight state configuration.
In the figure: the sun observation load is 1, the view field of the sun observation load is 2, the solar wing is 3, and the satellite body is 4.
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 various changes and modifications can be made by those skilled in the art without departing from the spirit of the invention, and these changes and modifications are all within the scope of the invention.
The satellite layout coordinate system (O-XYZ) referred to in the following examples is defined as:
origin of coordinates O: the lower end frame of the satellite-rocket connecting ring and the theoretical center of the satellite-rocket separating surface;
OX axis: the direction of the satellite body is perpendicular to the satellite-rocket separation surface and points to the direction of the satellite body along the origin of coordinates;
OZ axis: in the satellite and rocket separation surface, the satellite body is vertically pointed to the ground;
OY axis: and a right-hand rectangular coordinate system is formed by the axis OZ and the axis OX.
The invention provides a satellite solar wing configuration design method meeting load-to-sun observation requirements, which comprises the following steps of:
and 2, analyzing the arrangement requirement of the solar wing in the on-orbit state. According to the characteristic that the sun illumination angle is within the range of 59-89 degrees, the optimal value of the included angle between the normal of the solar wing in-orbit state cell piece and the-Y axis of the satellite body is 1-31 degrees, and the included angle between the normal of the solar wing in-orbit state cell piece and the-Y axis of the satellite body is finally determined to be 0 degrees by combining the engineering feasibility of solar wing design and the whole satellite energy demand, namely the solar wing in-orbit state cell piece surface is parallel to the XOZ surface of the satellite;
and 3, analyzing the limit requirements of the carrier rocket on the overall dimension of the satellite. As shown in figure 3, the constraint size of the carrier rocket for the satellite transverse envelope is phi 3310mm, the sun observation load can be arranged on the + Z surface, -Z surface, + X surface, -X surface and-Y surface of the satellite by combining the characteristic that the satellite body is hexahedron shape and the on-orbit work requirement of the sun observation load to be measured, but the earth observation load is arranged on the + Z surface of the satellite, the satellite-X surface is in butt joint with the carrier, and the satellite-Y surface is sunny surface and is not beneficial to load heat dissipation, so the arrangement position of the sun observation load can be the-Z surface and the + X surface of the satellite. The solar wing can be arranged on the-Y surface, + Z surface and-Z surface of the satellite according to the characteristic that the satellite body is hexahedron and the solar wing cell slice is required to be irradiated by sunlight on the orbit as vertically as possible, but the arrangement position of the folded solar wing can only be the-Y surface and the-Z surface of the satellite considering that the + Z surface of the satellite is used for arranging the earth observation load.
And 4, according to the constraint conditions determined in the previous three steps, sequentially considering various arrangement and combination modes of the solar wing and the arrangement position of the sun-facing observation load, comprehensively considering the sun-facing direction requirements of the solar wing and the load according to the principle of reducing the height of the longitudinal center of mass of the whole satellite as much as possible and utilizing the available envelope space of the satellite as much as possible, embedding the load sun-facing observation field into the solar wing (as shown in figure 4), determining the layout mode that the solar wing is arranged on the Y surface of the satellite when folded and the sun-facing observation load is arranged on the Z surface of the satellite, and forming the satellite solar wing configuration meeting the load sun-facing observation requirements, wherein the satellite configuration in the transmitting state is shown in figure 5, and the satellite configuration in the flying state is shown in figure 6.
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 (8)
1. A satellite solar wing configuration design method meeting load-to-sun observation requirements is characterized by comprising the following steps:
step 1, analyzing the field requirement of the load on daily observation, and determining the field envelope of the load according to the load working mode;
step 2, analyzing the arrangement requirement of the in-orbit state of the solar wing, and determining the relative position relationship between the solar wing and the satellite body according to the change condition of the sun illumination angle;
step 3, analyzing the limitation requirement of the carrier rocket on the overall dimension of the satellite, and determining the optional arrangement form of the load and the solar wing launching state by combining the overall state of the satellite body;
and 4, embedding the load sun observation view field into the solar wing according to the constraint conditions determined in the three steps, and determining the layout positions of the load and the solar wing on the satellite body, so that the satellite solar wing configuration meeting the load sun observation requirement is formed.
2. The satellite solar wing configuration design method meeting the requirement of load-to-sun observation according to claim 1, wherein the load is a remote sensing camera loaded on a satellite and used for monitoring the change condition of solar energy.
3. The method for designing the satellite solar wing configuration meeting the requirement of load-on-day observation according to claim 1, wherein the load view field envelope of the step 1 is obtained by traversing the motion range of a view field at a certain position of the load through an array.
4. The satellite solar wing configuration design method meeting the requirement of load-to-sun observation according to claim 1, wherein the specific method for determining the relative position relationship between the solar wing and the satellite body in the step 2 is as follows: according to the variation range of the solar illumination angle, the optimal value of the included angle between the solar wing in-orbit state cell and the satellite body is calculated, and the included angle between the solar wing in-orbit state cell and the satellite body is finally determined by combining the engineering feasibility of solar wing design and the energy requirement of the whole satellite.
5. The method for designing the satellite solar wing configuration meeting the requirement of load-to-sun observation according to claim 1, wherein the specific method for determining the optional arrangement form of the load in the step 3 is as follows: according to the requirement of a carrier rocket on the maximum envelope size of a satellite, all the arrangement position schemes of the sun-facing observation loads are given by combining the appearance characteristics of the satellite body and the on-orbit working requirement of the sun-facing observation loads to be measured for sun tracking, then the engineering feasibility of each scheme is analyzed one by one, and finally the optional arrangement position of the sun-facing observation loads is determined.
6. The satellite solar wing configuration design method meeting the load-on-day observation requirement of claim 5, wherein the specific method for determining the optional arrangement form of the solar wing emission state in the step 3 is as follows: according to the requirement of a carrier rocket on the maximum envelope size of a satellite, all solar wing arrangement position schemes are given by combining the appearance characteristics of a satellite body and the on-orbit working requirement that solar wing cells are irradiated by sunlight as perpendicularly as possible, then the engineering feasibility of each scheme is analyzed one by one, and the optional arrangement position of the solar wings is finally determined.
7. The satellite solar wing configuration design method meeting the load-sun observation requirement according to claim 6, characterized in that in step 4, various arrangement and combination modes of solar wings and sun-sun observation load arrangement positions are sequentially considered, the sun-sun pointing requirements of the solar wings and the load are comprehensively considered according to the principle of reducing the whole satellite longitudinal mass center height as much as possible and utilizing the available envelope space of the satellite as much as possible, the load-sun observation view field is embedded into the solar wings, and the arrangement positions of the load and the solar wings on the satellite body are clarified.
8. The method for designing a satellite solar wing configuration meeting the requirement of load-on-day observation according to claim 1, wherein only one solar wing is configured for the satellite.
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