CN115258204B - Method for increasing energy of medium-low orbit communication satellite - Google Patents
Method for increasing energy of medium-low orbit communication satellite Download PDFInfo
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- CN115258204B CN115258204B CN202211077635.2A CN202211077635A CN115258204B CN 115258204 B CN115258204 B CN 115258204B CN 202211077635 A CN202211077635 A CN 202211077635A CN 115258204 B CN115258204 B CN 115258204B
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- 238000004891 communication Methods 0.000 title claims abstract description 36
- 238000000034 method Methods 0.000 title claims abstract description 18
- 230000005855 radiation Effects 0.000 claims abstract description 36
- 229910052684 Cerium Inorganic materials 0.000 claims abstract description 24
- GWXLDORMOJMVQZ-UHFFFAOYSA-N cerium Chemical compound [Ce] GWXLDORMOJMVQZ-UHFFFAOYSA-N 0.000 claims abstract description 24
- 239000011521 glass Substances 0.000 claims abstract description 24
- 238000000576 coating method Methods 0.000 claims abstract description 9
- 238000005286 illumination Methods 0.000 claims abstract description 9
- 239000011248 coating agent Substances 0.000 claims abstract description 8
- 230000007246 mechanism Effects 0.000 claims description 4
- WHXSMMKQMYFTQS-UHFFFAOYSA-N Lithium Chemical compound [Li] WHXSMMKQMYFTQS-UHFFFAOYSA-N 0.000 abstract description 2
- 229910052744 lithium Inorganic materials 0.000 abstract description 2
- 230000017525 heat dissipation Effects 0.000 description 9
- 238000010586 diagram Methods 0.000 description 4
- 230000008859 change Effects 0.000 description 3
- 238000001816 cooling Methods 0.000 description 3
- 230000004907 flux Effects 0.000 description 3
- 230000000694 effects Effects 0.000 description 2
- 239000003973 paint Substances 0.000 description 2
- 206010034719 Personality change Diseases 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
Classifications
-
- 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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- Engineering & Computer Science (AREA)
- Life Sciences & Earth Sciences (AREA)
- Sustainable Development (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Photovoltaic Devices (AREA)
Abstract
The invention particularly relates to a medium-low orbit communication satellite energy source increasing method. The energy of the medium-low orbit communication satellite is improved by combining with the design of a thermal control radiating surface; in the satellite in-orbit motion, the radiating surface is used as an energy supply device at the same time, and the specific method comprises the following steps: and the Z-surface is used as a radiating surface, a cerium glass secondary surface reflector is used as a thermal control coating of the Z-radiating surface, and the cerium glass secondary surface reflector reflects solar radiation to a sailboard of the medium-low orbit communication satellite, so that the energy source of the medium-low orbit communication satellite is increased. The novel energy source device has the advantages that the radiating surface with single original function has a novel function of providing energy sources for medium-low orbit communication satellites, different functions are combined onto the same component by utilizing the illumination reflection principle, and under the premise that the area of a sailboard is not increased, when the beta angle of the satellite exceeds the rotation range of the SADA B axis, the satellite can acquire additional energy sources, the discharge depth of a lithium battery is reduced, and the time required by energy balance is reduced. The increase of the whole star quality and the cost caused by a method for expanding the area of the sailboard battery piece is avoided.
Description
Technical Field
The invention relates to a spacecraft thermal control design and an energy design, in particular to a medium-low orbit communication satellite energy increasing method.
Background
The medium-low orbit communication satellite has obvious advantages in the military field and the base station-free area due to low communication delay and low construction cost, and is actively constructed at home and abroad. The middle-low orbit communication satellites are distributed in the inclined orbits to acquire multi-target return visits, the satellite in-orbit beta angle (the included angle between the sun vector and the orbit plane) has a large variation range, the satellite posture is oriented to the ground, the three axes are stable, and a two-axis sailboard driving mechanism (SADA) is usually adopted to drive the solar sailboard to track the sun for acquiring energy, so that satellite business development is ensured.
The SADA B axis rotation angle is limited by the field of view of the communication satellite antenna and is less than-90 degrees to +90 degrees, and when the beta angle exceeds the B axis rotation range, the effective irradiation of the satellite sailboard is reduced, and the acquired energy is weakened. In order to ensure the full energy supply of the satellite in orbit, the effective irradiation amount of the sailboard is increased by expanding the area of the battery plate of the sailboard, but the attaching base plate and the cable are also increased, the quality and the cost of the whole satellite are increased, and the control difficulty of the sailboard is increased.
Disclosure of Invention
Aiming at the problem that in the prior art, when the rotation of the two SADA B shafts is limited, the acquired energy of the whole satellite is weakened, and the satellite acquires additional energy by combining a thermal control design on the premise of not increasing the area of a solar sailboard.
The invention provides a method for improving the energy of a medium-low orbit communication satellite, which combines the design of a heat control radiating surface to improve the energy of the medium-low orbit communication satellite; when the satellite moves in orbit, the radiating surface is used as an energy supply device at the same time, and the specific method comprises the following steps: and the Z-surface is used as a radiating surface, a cerium glass secondary surface reflector is used as a thermal control coating of the Z-radiating surface, and the cerium glass secondary surface reflector reflects solar radiation to a sailboard of the medium-low orbit communication satellite, so that the energy source of the medium-low orbit communication satellite is increased.
Further, the medium-low orbit communication satellites are distributed in an inclined orbit, the satellite attitude is +Z axis to earth orientation and flies along +X direction, and the triaxial is stable.
Further, the medium-low orbit communication satellite adopts a two-axis solar array driving mechanism, the axis A is used for sun tracking, and the rotation angle is-180 degrees to +180 degrees; the B axis is used for controlling the relation between the sailboard and the incident angle of the sun, and the rotation angle is-theta to +theta, wherein theta is less than 45 degrees.
Further, the change range of the angle beta of the middle-low orbit communication satellite in orbit, namely the included angle between the solar vector and the plane of orbit, is close to-90 degrees to +90 degrees.
Further, the satellite-Z plate has a width W, a length L, a sailboard length M, a single-side sailboard width total length X, and an effective area of a single-side sailboard battery piece S; the length of the B-axis supporting rod is T, the included angle between solar irradiation and the Z plate is phi, and the included angle between the satellite and the Z plate as well as the Y plate is phiThe linear distance from the axis A to the Z plate is Q, and the cerium glass secondary surface reflecting mirror performs specular reflection on solar irradiation, and is->When the satellite is in orbit |beta|<And the included angle between solar illumination and the sailboard is 90 degrees, the solar radiation is reflected into space by the Z-plate cerium glass secondary surface reflector, and the effective area of the sailboard for receiving the solar radiation is 2S.
Further, when the satellite is in orbitWhen the solar panel is used, the control angles of the B axes at the two sides are theta and-theta, and the included angle between solar illumination and the sailboard is +.>-the Z-plate cerium glass secondary surface mirror reflects solar radiation into space, the sailboard receiving an effective solar radiation area of 2Scos (|β| - θ).
Further, when the satellite is in orbitWhen the solar radiation angle is controlled to be theta and-theta, the solar radiation height H=Wsin phi reflected by the secondary surface mirror of the-Z-plate cerium glass, and the included angle phi=theta-phi between the solar radiation reflected by the secondary surface mirror of the-Z-plate cerium glass and the sailboard.
Further, whenWhen the solar radiation is reflected into space by the Z-plate cerium glass secondary surface reflecting mirror, the effective radiation area of the sailboard for receiving the sun is 2Scos (beta-theta).
Further, whenWhen the solar radiation reflected by the Z-plate cerium glass secondary surface reflecting mirror falls to the width on the sailboardThe effective irradiation area of the sailboard for receiving the sun is
Further, whenWhen the solar radiation reflected by the Z-plate cerium glass secondary surface reflecting mirror falls to the width of the sailboard +.>The effective irradiation area of the sailboard for receiving the sun is 2Scos (beta-theta) +WL cos beta.
The beneficial effects of the invention are as follows: the novel solar energy power supply device has the advantages that the radiating surface with single original function has a new function of providing energy sources for medium-low orbit communication satellites, different functions are combined onto the same component by utilizing the illumination reflection principle, and when the beta angle exceeds the rotation range of the B axis on the premise of not increasing the area of a sailboard, the satellite can acquire additional energy sources, the discharge depth of a lithium battery is reduced, and the time required by energy balance is reduced. The increase of the whole star quality and the cost caused by a method for expanding the area of the sailboard battery piece is avoided.
Drawings
FIG. 1 is a schematic diagram of a medium-low orbit communication satellite;
FIG. 2 is a graph of simulated average solar irradiance heat flux density reaching each face of a satellite at different beta angles;
FIG. 3 is a graph of average heat dissipation capacity versus angle for a Z-plane using different thermal control coatings;
FIG. 4 is a schematic diagram of the solar radiation of a windsurfing board with increased heat dissipation surface when an orbiting satellite is directly under the sun;
FIG. 5 is a schematic diagram of the influence of the heat dissipation surface on the solar radiation received by the sailboard under different beta angle illumination conditions when the satellite is positioned right under the sun; wherein a) is that the reflected solar radiation does not fall on the sailboard, b) is that the reflected solar radiation partially falls on the sailboard, c) is that the reflected solar radiation completely falls on the sailboard;
FIG. 6 is a graph showing current contrast for a satellite windsurfing board for different beta angle illumination conditions at a time when the satellite is directly under the sun in an exemplary embodiment.
Detailed Description
The following description of the embodiments of the present invention will be made apparent and fully in view of the accompanying drawings, in which some, but not all embodiments of the invention are shown. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
In the description of the present invention, it should be noted that the directions or positional relationships indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. are based on the directions or positional relationships shown in the drawings, are merely for convenience of describing the present invention and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In the description of the present invention, it should be noted that, unless explicitly specified and limited otherwise, the terms "mounted," "connected," and "connected" are to be construed broadly, and may be either fixedly connected, detachably connected, or integrally connected, for example; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present invention will be understood in specific cases by those of ordinary skill in the art.
The invention solves the technical problems that: aiming at the problem that in the prior art, when the rotation of the two SADA B shafts is limited, the acquired energy of the whole satellite is weakened, and on the premise of not increasing the area of a solar sailboard, the satellite acquires additional energy by combining a thermal control design.
As shown in fig. 1 and 4, the invention aims at the existing middle-low orbit communication satellites distributed in an inclined orbit, the satellite attitude is oriented to the ground by +Z axis, and the triaxial is stable; the satellite selects a solar sailboard driving mechanism (SADA) for ensuring service capability, and preferably two-axis SADA; the axis A is used for sun tracking, and the rotation angle is-180 degrees to +180 degrees; b axisThe rotation angle of the B axis is limited by the field of view of the outdoor antenna, and the rotation angle is-theta to +theta, and is generally theta<45 deg.. The width of the satellite-Z plate is W, the length of the satellite-Z plate is L, the length of the sailboard is M, the total width length of the unilateral sailboard is X, and the effective area of the unilateral sailboard cell is S; the length of the SADA B-axis supporting rod is T, the included angle between solar irradiation and the Z plate is phi, and the included angle between the satellite and the Z plate is Y plateThe SADA has an A-axis to-Z-plate linear distance Q and the OSR specularly reflects solar radiation.
Fig. 5 is a schematic diagram showing the influence of the heat dissipation surface on the solar radiation received by the sailboard under different beta angle illumination conditions when the satellite is positioned right under the sun; wherein a) is that the reflected solar radiation does not fall on the sailboard, b) is that the reflected solar radiation partially falls on the sailboard, and c) is that the reflected solar radiation completely falls on the sailboard.
The selection of the heat radiating surface and the selection of the heat radiating surface thermal control coating are described below.
Simulation solar constant 1367W/m 2 The earth albedo coefficient is 0.3, the average solar irradiation heat flux density reaching each surface is shown in fig. 2, and the infrared heat flux density reaching each surface is shown in table 1.
The + -Y surface is affected by alternating solar irradiation heat flow, and the earth infrared heat flow of the + -Z surface is larger, so that the earth infrared heat flow is not suitable for being used as a radiating surface.
The heat flow of solar irradiation of the X surface and the Z surface and the heat flow of the earth infrared are smaller, and the solar heat radiator is suitable for being used as a radiating surface.
TABLE 1
The selection of the-Z cooling surface using the thermal control coating compares the five year end of life cerium glass secondary surface mirror (OSR) (α=0.13, epsilon=0.8), S781 (α=0.4, epsilon=0.87) organic white paint, KS-ZA (α=0.22, epsilon=0.92) inorganic white paint cooling capability.
As can be taken from fig. 3, the heat dissipation capacity KS-ZA > OSR > S781, and the stability OSR > KS-ZA > S781.
In consideration of satellite attitude change and heat dissipation stability during the launching process, an OSR is preferably used as the-Z-plane thermal control coating.
As shown in fig. 4, the solar energy moves in the yellow road plane and the intersection point of the solar energy and the elevation drifts, so that the change range of the in-orbit beta angle (the included angle between the solar vector and the orbit plane) of the medium-low orbit communication satellite is large, and the change range is approximately-90 degrees to +90 degrees.
The communication satellite has high load power density and constant power consumption, and heat dissipation is a prominent problem, and in the orbit motion of the satellite, the Z surface is subjected to small solar irradiation heat flow, and is suitable to be used as a heat dissipation surface, and the Z heat dissipation surface heat control coating is preferably a cerium glass secondary surface reflector (OSR).
The present invention will be described in further detail with reference to specific examples.
The satellite adopts two SADA shafts, the A shaft is arranged on a-Z honeycomb plate, namely, the Q length is 0, the rotation angle of the A shaft is-180 degrees to +180 degrees, and the rotation angle of the B shaft is-40 degrees to +40 degrees.
The satellite-Z board has a width W of 1.5M, a length L of 2M, a sailboard length M of 2.3M, a single-side sailboard width total length X of 3.1M, and a sailboard cell effective area 2S of 13.68M 2 The method comprises the steps of carrying out a first treatment on the surface of the The length T of the SADA B shaft supporting rod is 1.2m.
When the incidence angle of solar irradiation of the sailboard is designed to be 0 DEG, the sailboard area 2S generates current of 65A, namely when |beta| <40 DEG, the sailboard generates current of 65A constantly in the sunlight area.
When (when)When, i.e. when 40 °<When beta is less than 60.86 degrees, the Z board OSR reflects solar illumination into space, the effective solar radiation area of the sailboard is 2Scos (beta is 40 degrees) and the current is 65cos (beta is 40 degrees) A.
When (when)When, i.e. when 60.86 °<|β|<At 68.56 deg., the solar radiation reflected by the-Z panel falls to the width of the windsurfing board +.>The effective irradiation area of the sailboard for receiving the sun is +.>The current is 65cos (|beta| -40 °) -29.46cos (40+|beta|) -11.4 cos|beta|).
When (when)When |beta|>At 68.56 deg., the solar radiation reflected by the-Z panel OSR falls to the width of the windsurfing board +.>The effective irradiation area of the sailboard for receiving the sun is 2Scos (|beta| -theta) +WL cos|beta|, namely the generated current is 65cos (|beta| -40 degrees) +14.25 cos|beta|.
As is known from fig. 6, for this satellite:
when |beta| <60.86 degrees, the-Z plate OSR cooling surface does not affect the whole star energy source.
When 60.86 degrees < beta| <68.56 degrees, as beta| increases, -Z plate OSR radiating surface is the energy of whole star increase and when beta| is 68.56 degrees, -Z plate OSR radiating surface has the most obvious effect on energy increase, when-Z plate adopts OSR radiating surface, whole star current is 62.3A, whole star current is 57.1A by adopting original design (-Z plate non-OSR thermal control coating) scheme, and whole star energy increase effect is obvious.
When 68.56 degrees < beta degrees <90 degrees, the Z plate OSR radiating surface is in a downward trend for the whole star increased energy source along with the increase of beta degrees, and when beta degrees is 90 degrees, the Z plate OSR radiating surface is consistent with the current generated by the original design scheme.
The embodiment can prove that when the beta angle exceeds the rotation range of the B axis, the satellite can acquire additional energy by combining the heat control design on the premise of not increasing the area of the solar sailboard.
Claims (7)
1. The method for increasing the energy of the medium-low orbit communication satellite is characterized in that a-Z surface is taken as a radiating surface, a cerium glass secondary surface reflector is selected as a thermal control coating of the-Z radiating surface, and the cerium glass secondary surface reflector reflects solar radiation to a sailboard of the medium-low orbit communication satellite so as to obtain additional energy;
the medium-low orbit communication satellites are distributed in an inclined orbit, the satellite gestures are oriented to the ground along a +Z axis, and fly along a +X direction;
the medium-low orbit communication satellite adopts a two-axis solar sailboard driving mechanism, wherein the axis A is used for sun tracking, and the rotation angle is-180 degrees to +180 degrees; the B axis is used for controlling the relation between the sailboard and the incident angle of the sun, and the rotation angle is~/>，/><45°；
Medium-low orbit communication satellite in orbitThe angle, i.e. the angle between the sun vector and the plane of the track, is at maximum +.>Satisfy the following requirements。
2. The method for increasing energy of a low-medium orbit communication satellite according to claim 1, wherein the satellite-Z plate has a width W, a length L, a sailboard length M, a single-sided sailboard width total length X, and a single-sided sailboard cell active area S; the length of the B-axis supporting rod is T, and the included angle between solar irradiation and the Z-plate isThe included angle between the satellite plus Z plate and the Y plate is +.>The linear distance from the axis A to the Z plate is Q, and the cerium glass secondary surface reflector performs specular reflection on solar irradiation; />When the satellite is in orbit->The included angle between solar illumination and a sailboard is 90 degrees, the Z-board cerium glass secondary surface reflector reflects solar radiation into space, and the effective area of the sailboard for receiving the solar radiation is 2S.
3. The method for increasing energy of low-medium orbit communication satellite according to claim 2, wherein when the satellite is in orbitWhen the control angle of the B shafts at two sides is +.>And->The included angle between the sunlight and the sailboard is +.>The Z-plate cerium glass secondary surface reflector reflects solar radiation into space, and the effective solar radiation receiving area of the sailboard is as follows。
4. The method for increasing energy of low-medium orbit communication satellite according to claim 3, wherein when the satellite is in orbitWhen the control angle of the B shafts at two sides is +.>And->-Z-plate cerium glass secondary surface mirror reflecting solar irradiance heightSolar radiation reflected by the Z-plate cerium glass secondary surface reflecting mirror forms an included angle +.>。
5. The method for increasing energy of low-medium orbit communication satellite according to claim 4, wherein whenWhen in use, the Z-plate cerium glass secondary surface reflector reflects solar radiation into space, and the effective radiation area of the sailboard for receiving the sun is +.>。
6. The method for increasing energy of low-medium orbit communication satellite according to claim 4, wherein whenWhen the solar radiation reflected by the Z-plate cerium glass secondary surface reflecting mirror falls to the width of the sailboard +.>The effective irradiation area of the sailboard for receiving the sun is +.>。
7. The medium-low orbit communication satellite energy according to claim 4A source adding method, characterized in that whenWhen the solar radiation reflected by the Z-plate cerium glass secondary surface reflecting mirror falls to the width of the sailboard +.>The effective irradiation area of the sailboard for receiving the sun is +.>。
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| CN202211077635.2A CN115258204B (en) | 2022-09-05 | 2022-09-05 | Method for increasing energy of medium-low orbit communication satellite |
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| CN115258204B true CN115258204B (en) | 2024-04-05 |
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4371135A (en) * | 1979-07-30 | 1983-02-01 | Rca Corporation | Solar array spacecraft reflector |
| US5618012A (en) * | 1995-01-12 | 1997-04-08 | Space Systems/Loral, Inc. | Satellite stabilization system |
| JPH11217100A (en) * | 1998-02-03 | 1999-08-10 | Nec Corp | Power recovery structure of spacecraft |
| KR20000007679A (en) * | 1998-07-06 | 2000-02-07 | 추호석 | Solar light reflecting device of satellite and power generating method using device thereof |
| CN110002010A (en) * | 2019-04-24 | 2019-07-12 | 中国人民解放军战略支援部队航天工程大学 | A kind of method of satellite optical camouflage |
| CN114564035A (en) * | 2022-03-09 | 2022-05-31 | 上海鉴创科技有限公司 | Double-shaft solar sailboard driving control method |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2012183855A (en) * | 2011-03-03 | 2012-09-27 | Japan Aerospace Exploration Agency | Apparatus and method for generating flash of light toward the earth using solar light reflection |
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- 2022-09-05 CN CN202211077635.2A patent/CN115258204B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4371135A (en) * | 1979-07-30 | 1983-02-01 | Rca Corporation | Solar array spacecraft reflector |
| US5618012A (en) * | 1995-01-12 | 1997-04-08 | Space Systems/Loral, Inc. | Satellite stabilization system |
| JPH11217100A (en) * | 1998-02-03 | 1999-08-10 | Nec Corp | Power recovery structure of spacecraft |
| KR20000007679A (en) * | 1998-07-06 | 2000-02-07 | 추호석 | Solar light reflecting device of satellite and power generating method using device thereof |
| CN110002010A (en) * | 2019-04-24 | 2019-07-12 | 中国人民解放军战略支援部队航天工程大学 | A kind of method of satellite optical camouflage |
| CN114564035A (en) * | 2022-03-09 | 2022-05-31 | 上海鉴创科技有限公司 | Double-shaft solar sailboard driving control method |
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