CN115180179A - Self-leveling low-orbit satellite solar wing pneumatic layout - Google Patents
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Abstract
The invention discloses a self-balancing low-orbit satellite solar wing pneumatic layout, which comprises a satellite body and two groups of self-balancing solar wings, wherein the satellite body is provided with a plurality of solar wings; the satellite body is a cuboid, two sets of self-balancing solar wings are respectively located on two sides of the satellite body to form a symmetrical structure, each self-balancing solar wing comprises a first folding solar wing, the first folding solar wings are connected with the satellite body to form an adjustable deflection angle, and the size of the deflection angle is adjusted to realize low-orbit satellite aerodynamic moment balance. The invention can eliminate the pneumatic interference torque when the satellite operates in low orbit near the earth, greatly reduce the influence of the thin atmosphere near the earth space on the satellite orbit, reduce the pneumatic interference torque of the satellite attitude control, and simultaneously can ensure the requirement of a heat protection system on the side surface of the satellite body, thereby improving the satellite reliability.
Description
Technical Field
The invention relates to a self-leveling low-orbit satellite solar wing pneumatic layout, and belongs to the technical field of satellite pneumatic leveling.
Background
A Low Earth Orbit (LEO) environment 200-1000 km away from the ground is an Earth observation satellite operation area, and a satellite operating in the LEO environment is a Low Orbit satellite. The ultra-low orbit satellite has low orbit height, can quickly reach a preset orbit and can be unfolded to work; the orbit period is short, the same target can be detected and observed in a short time, the detection work is more frequent than that of the traditional satellite, and more information is obtained; moreover, the indexes of equipment such as a space camera or a radar and the like carried by the satellite are not too high, the effect of the traditional satellite can be achieved, even better, and the cost of the satellite can be greatly reduced by the effective load; the return landing point of the ultra-low orbit satellite can also obtain higher precision. Therefore, the research and design of the ultra-low orbit satellite can generate huge social and economic effects and have greater engineering application value. Ultra-low orbit satellites are becoming the focus of attention of all aerospace major countries, and low orbit satellite research is started in various universities and research institutes.
However, in an ultra-low orbit environment, particularly at 200-300km altitude, the thin atmosphere of the near-earth space causes large aerodynamic resistance (the density of the atmosphere near 200km altitude is more than 1000 times of 700km altitude), which enables the earth observation satellite to rapidly deviate from the target orbit, and the earth observation satellite must be speed-compensated and frequently orbit maneuvers are performed; meanwhile, if the satellite configuration design is asymmetric or the centroid deviates from the satellite centroid too much, the near-earth rarefied atmosphere can generate a large pneumatic interference torque for controlling the satellite attitude, and particularly, the amplitude of the pneumatic interference torque is large and can reach 100g cm for a near-earth satellite provided with a single-wing solar sailboard for two-dimensional sun-to-sun directional control. Although the force of the near earth rarefied atmosphere is small compared with the gravity of the earth, the force is accumulated in the satellite operation time, and long-term perturbation has great influence on the orbit and the attitude of the satellite, so that the satellite actuator can carry out frequency unloading.
Usually, the control part of the satellite always uses the aerodynamic moment as disturbing moment, balanced by a reaction flywheel or jet. For example, the first geostationary Satellite, SLATS (Super Low altitute Test Satellite), launched by the Japanese aerospace research and development agency (JAXA) in 2017, which runs at ultra-Low orbits below 300km (as shown in FIG. 1), eventually falls in orbit to an Altitude of 268km to 150km, and employs the first ion engine in the world to maintain a predetermined orbit in order to overcome the air resistance in space. The solar wings adopted by the satellite are horizontally arranged in parallel, as shown in figure 1, the satellite design can be influenced by large aerodynamic interference moment when running at a low earth orbit, and meanwhile, the stability and pointing accuracy of the satellite attitude are influenced, and large energy consumption is generated.
Therefore, if measures are taken to reduce or eliminate the pneumatic interference moment received by the satellite during the low-earth-orbit operation, not only can enough attitude stability and attitude pointing accuracy be obtained, but also the energy consumption of the satellite can be effectively reduced, so that the satellite development cost is reduced, the economic value is higher, and a new idea and strategy are provided for the future research on the control system of the low-cost low-power-consumption satellite.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the solar wing aerodynamic layout of the self-leveling low-orbit satellite can effectively reduce or even eliminate the aerodynamic moment of the low-orbit satellite, and can meet the requirement of a heat protection system on the side surface of a satellite body, thereby improving the reliability of a satellite control system.
The technical solution of the invention is as follows:
a self-balancing low-orbit satellite solar wing pneumatic layout comprises a satellite body and two groups of self-balancing solar wings; the satellite body is a cuboid, two groups of self-balancing solar wings are respectively located on two sides of the satellite body to form symmetrical structures, each self-balancing solar wing comprises a first folding solar wing, the first folding solar wing is connected with the satellite body to form an adjustable deflection angle, and the size of the deflection angle is adjusted to realize low-orbit satellite aerodynamic moment balance.
In the aerodynamic layout of the sun wing, the first folding sun wing is deflected downwards according to the aerodynamic trim and forms the deflection angle theta with the extended surface of the upper surface of the satellite body by taking a connecting line of the first folding sun wing and the satellite body as a rotating shaft.
In the above pneumatic layout of solar wings, the self-balancing solar wing further includes a second folding solar wing and a third folding solar wing, the first folding solar wing, the second folding solar wing and the third folding solar wing are all flat plate structures, the side edges are connected in sequence, and the second folding solar wing and the third folding solar wing are on the same horizontal plane.
In the aerodynamic layout of the solar wings, the first, second and third solar wings have the same shape, and a gap size Lp between the first and second solar wings is the same as a gap size between the second and third solar wings.
In the aerodynamic layout of the solar wing, the length L of the self-balancing solar wing is 3 to 5 times the width Lw.
In the aerodynamic layout of the solar wings, the length of the satellite body is equal to the length L of the self-leveling solar wings, the length L is 4-8 times of the height H, and the height H is 0.8-1.5 times of the width W.
In the aerodynamic layout of the solar wing, the deflection angle θ ranges from 0 ° to 50 °.
In the aerodynamic layout of the solar wing, the method for realizing the aerodynamic moment balance of the low-orbit satellite by adjusting the deflection angle specifically comprises the following steps:
(1) Obtaining a satellite reference centroid coordinate (x) from a geometric model of the satellite g 、y g 、z g )；
(2) Meshing the satellite body and the surface of the solar wing to generate grid units, and outputting normal and position parameters of the surface of each grid unit;
(3) Selecting a solar activity parameter F10.7 and a geomagnetic parameter Ap according to the flight altitude and the flight date of the satellite;
(4) Calculating the atmospheric density rho of the satellite flight environment by using an atmospheric model according to the solar activity parameter F10.7 and the geomagnetic parameter Ap in the step (3);
(5) Selecting test particles according to the normal and position parameters of the surfaces of the grid cells in the step (2) and the atmospheric density rho in the step (4), simulating the interaction between gas molecules and the surfaces of the whole satellites by utilizing a satellite aerodynamic force rapid calculation method TPMC, and calculating the momentum exchange quantity after the test particles collide with the surfaces of the satellites;
(6) According to the satellite reference centroid coordinate (x) in the step (1) g 、y g 、z g ) And the value of the deflection angle theta, and calculating the pneumatic force and the moment of the whole satellite of the satellite according to the momentum exchange quantity in the step (5);
(7) And (5) adjusting the value of the deflection angle theta according to the result of the aerodynamic force and moment of the whole satellite of the satellite, and repeating the step (6) until the aerodynamic moment of the whole satellite of the satellite is 0, so as to obtain the value of the deflection angle theta at the moment.
In the aerodynamic layout of the solar wing, the calculation formula for calculating the aerodynamic force and moment of the whole satellite of the satellite in the step (6) is as follows:
wherein F is aerodynamic force of the whole satellite, M is moment of the whole satellite, and N is E The total number of grid cells for the whole satellite surface, F l 、M l Respectively the pneumatic force and the moment applied to the ith grid unit.
Aerodynamic force F on the l grid cell l Sum moment M l The calculation formula of (2) is as follows:
where N is the total number of collisions of all test particles with the satellite surface, r j 、ΔP j Respectively vector and momentum exchange quantity from a satellite reference mass center to a collision position in the j-th collision; a is a weight factor representing the number of real gas molecules represented by a test particle.
In the aerodynamic layout of the solar wing, a calculation formula of a vector from the satellite reference centroid to the collision position in the j-th collision is as follows:
wherein θ is the deflection angle, (x) g ，y g ，z g ) Referencing centroid coordinates for the satellite, (x) l ，y l ，z l ) And the collision position coordinates are i, j and k respectively are unit vectors in x, y and z directions.
In the aerodynamic layout of the solar wing, the calculation formula of the momentum exchange amount from the satellite reference centroid to the collision position in the j-th collision is as follows:
ΔP j ＝mc i -mc r
in the formula, c i The velocity vector of the incident molecule that is the test particle; c. C r Is the molecular velocity vector of the complete diffuse reflection of the test particle and m is the mass of the incoming gas molecules in the test particle.
Compared with the prior art, the invention has the advantages that:
(1) According to the invention, through the design of the first folding solar wing in the rotary multi-folding solar wing, the aerodynamic interference moment of the thin atmosphere to the satellite can be eliminated. Compared with the existing horizontal solar wing non-deflection design, the method has the advantages that enough attitude stability and attitude pointing accuracy can be obtained, the energy consumption of the satellite can be effectively reduced, the satellite development cost is reduced, and the method has a great economic value.
(2) The design of the self-leveling solar wing only changes the first folding and deflecting angle in the multi-folding solar wing, has small influence on the design appearance of a satellite, and is more reliable and simpler to operate.
(3) The design of the self-leveling solar wing only changes the first folding and deflecting angle in the multi-folding solar wing, and the other solar wings keep the horizontal state, so that the heat dissipation requirement of the side surface of the satellite can be ensured, and the requirement on a heat protection system is reduced.
Description of the drawings:
FIG. 1 is a schematic diagram of a low-orbit satellite SLATS of JAXA of Japan;
FIG. 2 is a schematic diagram of the aerodynamic layout of the self-trim low earth orbit satellite solar wings of the present invention;
FIG. 3 is a self-balancing solar wing low-orbit satellite surface grid of the present invention;
FIG. 4 is a flow chart of the aerodynamic moment balance design of the present invention;
FIG. 5 is a diagram of the self-trim aerodynamic trim characteristics of a 250km pitch direction for a solar wing low-orbit satellite according to the invention.
Detailed description of the preferred embodiments
The following examples are given to illustrate specific embodiments of the present invention.
Fig. 2 is a schematic diagram showing the aerodynamic layout of the self-balancing low-orbit satellite solar wings of the present invention, and one aerodynamic layout of the self-balancing low-orbit satellite solar wings comprises a satellite body 1 and two sets of self-balancing solar wings 2; the satellite body 1 is a cuboid, two groups of self-balancing solar wings 2 are respectively located on two sides of the satellite body 1 to form symmetrical structures, each self-balancing solar wing 2 comprises a first folding solar wing 21, each first folding solar wing 21 is connected with the satellite body 1 to form an adjustable deflection angle, and the size of the deflection angle is adjusted to achieve low-orbit satellite aerodynamic moment balance.
With a connecting line between the first solar folding wing 21 and the satellite body 1 as a rotating shaft, the first solar folding wing 21 deflects downwards according to pneumatic trim and forms a deflection angle θ with the upper surface of the satellite body 1 along the extended surface. The deflection angle theta is in the range of 0-50 degrees.
The self-balancing solar wing 2 further comprises a second folding solar wing 22 and a third folding solar wing 23, the first folding solar wing 21, the second folding solar wing 22 and the third folding solar wing 23 are of flat plate structures, the side edges of the first folding solar wing, the second folding solar wing and the third folding solar wing are connected in sequence, and the second folding solar wing 22 and the third folding solar wing 23 are on the same horizontal plane. The first, second, and third sun flaps 21, 22, 23 have the same shape, and the gap size Lp between the first and second sun flaps 21, 22 is the same as the gap size between the second and third sun flaps 22, 23. The length L of the self-trim solar wing 2 is 3 to 5 times the width Lw. The length of the satellite body 1 is equal to the length L of the self-leveling solar wing 2, the length L is 4-8 times of the height H, and the height H is 0.8-1.5 times of the width W.
As shown in fig. 4, which is a flow chart of aerodynamic moment balance design of the invention, the method for realizing aerodynamic moment balance of a low orbit satellite by adjusting the size of a deflection angle specifically comprises the following steps:
(1) Obtaining a satellite reference centroid coordinate (x) from a geometric model of the satellite g 、y g 、z g )；
(2) As shown in fig. 3, the invention is a self-balancing solar wing low-orbit satellite surface grid, which is used for generating grid units by meshing the satellite body and the solar wing surface and outputting normal and position parameters of each grid unit surface;
(3) Selecting a solar activity parameter F10.7 and a geomagnetic parameter Ap according to the flight altitude and the flight date of the satellite;
(4) Calculating the atmospheric density rho of the satellite flight environment by using an NRLMSISE-00 atmospheric model according to the solar activity parameter F10.7 and the geomagnetic parameter Ap in the step 3;
(5) Selecting Test particles according to the normal and position parameters of the surfaces of the grid units in the step 2 and the atmospheric density rho in the step 4, simulating the interaction between gas molecules and the surfaces of the whole satellites by utilizing a satellite aerodynamic force rapid calculation method TPMC (Test Particle Monte Carlo), and calculating the momentum exchange quantity after the Test particles collide with the surfaces of the satellites;
(6) According to the satellite reference centroid coordinate (x) in step 1 g 、y g 、z g ) And the value of the deflection angle theta, and calculating the pneumatic force and the moment of the whole satellite of the satellite according to the momentum exchange quantity in the step 5;
(7) And (5) adjusting the value of the deflection angle theta according to the result of the aerodynamic force and moment of the whole satellite of the satellite, and repeating the step (6) until the aerodynamic moment of the whole satellite of the satellite is 0, so as to obtain the value of the deflection angle theta.
The calculation formula for calculating the aerodynamic force F and the moment M of the whole satellite of the satellite is as follows:
in the formula, N E The total number of grid cells for the whole satellite surface, F l 、M l Respectively is the aerodynamic force and the moment that receive on the ith grid cell, and has:
in which N is allTotal number of collisions of test particles with the satellite surface, r j 、ΔP j Respectively vector and momentum exchange quantity from a satellite reference mass center to a collision position in the j-th collision; a is a weight factor representing the number of real gas molecules represented by a test particle.
The calculation formula of the vector from the satellite reference centroid to the collision position at the jth collision is as follows:
in the formula, theta is a deflection angle (x) g ，y g ，z g ) Reference to the centroid coordinates for the satellite, (x) l ，y l ，z l ) And the collision position coordinates are i, j and k respectively are unit vectors in x, y and z directions.
The calculation formula of the momentum exchange amount from the satellite reference mass center to the collision position in the j-th collision is as follows:
ΔP j ＝mc i -mc r
in the formula, c i The velocity vector of the incident molecule that is the test particle; c. C r Is the molecular velocity vector of the complete diffuse reflection of the test particle and m is the mass of the incoming gas molecules in the test particle.
If the aerodynamic disturbance moment of the whole satellite of the satellite does not reach the expected target value or is eliminated to be zero, the rotation angle of the first folding solar wing is continuously optimized, the translation of the second folding solar wing and the third folding solar wing is kept, and the horizontal gap of the three folding solar wings is ensured not to be changed; and if the aerodynamic disturbance moment of the whole satellite of the satellite reaches an expected target value or is eliminated to be zero, the solar wing is the optimized self-leveling solar wing layout.
The specific solution example of the present invention is as follows:
the incoming flow gas calculated by the method example of the present invention is air, as shown in fig. 2, in the present embodiment, the satellite body size is: 2700mm X330 mm; two sides of the satellite are respectively provided with 3 solar cell wings, and the size is as follows: the satellite flight height is set to be 250km, the solar activity parameter F10.7 is 185.75 of 2037 years, 4 months and 28 days, the geomagnetic parameter is 4.5, a 250km pitching direction pneumatic balancing characteristic diagram of the self-balancing solar wing low-orbit satellite is shown in fig. 5, and the 250km pitching direction pneumatic balancing characteristic diagram of the satellite is shown in table 1.
TABLE 1
As can be seen from the table, the aerodynamic interference moment in the pitching direction of the whole satellite of the satellite can be eliminated by optimizing the deflection angle of the first folding solar wing, the deflection angle of the solar wing is only 27 degrees, and the heat dissipation requirement of the side face of the satellite body is met.
The present invention is not disclosed in the technical field of the common general knowledge of the technicians in this field.
Claims (12)
1. The utility model provides a low orbit satellite solar wing aerodynamic configuration from balancing which characterized in that: comprises a satellite body (1) and two groups of self-balancing solar wings (2); the satellite body (1) is a cuboid, two sets of self-balancing solar wings (2) are respectively located on two sides of the satellite body (1) to form a symmetrical structure, each self-balancing solar wing (2) comprises a first folding solar wing (21), the first folding solar wings (21) are connected with the satellite body (1) to form an adjustable deflection angle, and the size of the deflection angle is adjusted to realize low-orbit satellite aerodynamic moment balance.
2. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 1, wherein: and the connecting line of the first folding solar wing (21) and the satellite body (1) is used as a rotating shaft, and the first folding solar wing (21) deflects downwards according to pneumatic trim and forms the deflection angle theta with the extended surface of the upper surface of the satellite body (1).
3. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 1, wherein: the self-balancing solar wing (2) further comprises a second folding solar wing (22) and a third folding solar wing (23), the first folding solar wing (21), the second folding solar wing (22) and the third folding solar wing (23) are flat plate structures, the side edges are connected in sequence, and the second folding solar wing (22) and the third folding solar wing (23) are on the same horizontal plane.
4. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 1, wherein: the shape of the first folding solar wing (21), the shape of the second folding solar wing (22) and the shape of the third folding solar wing (23) are the same, and the size Lp of a gap between the first folding solar wing (21) and the second folding solar wing (22) is the same as the size Lp of a gap between the second folding solar wing (22) and the third folding solar wing (23).
5. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 1, wherein: the length L of the self-balancing solar wing (2) is 3-5 times of the width Lw.
6. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 1, wherein: the length of the satellite body (1) is equal to the length L of the self-leveling solar wing (2), the length L is 4-8 times of the height H, and the height H is 0.8-1.5 times of the width W.
7. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 1, wherein: the value range of the deflection angle theta is 0-50 degrees.
8. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 1, wherein: the method for realizing the low-orbit satellite aerodynamic moment balance by adjusting the size of the deflection angle comprises the following specific steps:
(1) Obtaining a satellite reference centroid coordinate (x) from a geometric model of the satellite g 、y g 、z g )；
(2) Carrying out mesh subdivision on the surfaces of the satellite body and the solar wings to generate mesh units, and outputting normal and position parameters of the surfaces of the mesh units;
(3) Selecting a solar activity parameter F10.7 and a geomagnetic parameter Ap according to the flight altitude and the flight date of the satellite;
(4) Calculating the atmospheric density rho of the satellite flight environment by using an atmospheric model according to the solar activity parameter F10.7 and the geomagnetic parameter Ap in the step (3);
(5) Selecting test particles according to the normal and position parameters of the surfaces of the grid cells in the step (2) and the atmospheric density rho in the step (4), simulating the interaction between gas molecules and the surfaces of the whole satellites by utilizing a satellite aerodynamic force rapid calculation method TPMC, and calculating momentum exchange quantity after the test particles collide with the surfaces of the satellites;
(6) According to the satellite reference centroid coordinate (x) in the step (1) g 、y g 、z g ) And the value of the deflection angle theta, and calculating the pneumatic force and the moment of the whole satellite of the satellite according to the momentum exchange quantity in the step (5);
(7) And (5) adjusting the value of the deflection angle theta according to the result of the aerodynamic force and moment of the whole satellite of the satellite, and repeating the step (6) until the aerodynamic moment of the whole satellite of the satellite is 0, so as to obtain the value of the deflection angle theta at the moment.
9. The aerodynamic layout of a self-levelling low-earth-orbit satellite solar wing according to claim 8, characterized in that: the calculation formula for calculating the aerodynamic force F and the moment M of the whole satellite in the step (6) is as follows:
wherein F is aerodynamic force of the whole satellite, M is moment of the whole satellite, and N is E For the total number of grid cells on the entire satellite surface, F l 、M l Respectively the pneumatic force and the moment applied to the first grid unit.
10. The aerodynamic layout of a self-leveling low-earth satellite solar wing according to claim 9, wherein: aerodynamic force F on the l grid cell l Sum moment M l The calculation formula of (2) is as follows:
wherein N is the total number of collisions between all test particles and the satellite surface, r j 、ΔP j Respectively vector and momentum exchange quantity from a satellite reference center of mass to a collision position during the jth collision; a is a weight factor representing the number of real gas molecules represented by a test particle.
11. The aerodynamic layout of a self-leveling low-earth satellite solar wing as claimed in claim 10, wherein: the calculation formula of the vector from the satellite reference centroid to the collision position in the j-th collision is as follows:
wherein θ is the deflection angle, (x) g ，y g ，z g ) Referencing centroid coordinates for the satellite, (x) l ，y l ，z l ) And the collision position coordinates are i, j and k respectively are unit vectors in x, y and z directions.
12. A self-levelling low-earth satellite solar wing aerodynamic layout according to claim 8 or 10, characterized in that: the momentum exchange quantity delta P from the reference mass center of the satellite to the collision position in the j collision j The calculation formula of (c) is:
ΔP j ＝mc i -mc r
in the formula, c i The velocity vector of the incident molecule that is the test particle; c. C r Is the molecular velocity vector of the complete diffuse reflection of the test particle and m is the mass of the incoming gas molecules in the test particle.
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0544241A1 (en) * | 1991-11-27 | 1993-06-02 | Hughes Aircraft Company | Method and apparatus for dynamic precompensation of solar wing stepping motions of a satellite |
| US20060038080A1 (en) * | 2002-12-12 | 2006-02-23 | Bernard Polle | Solar control method for spacecraft |
| CN201172484Y (en) * | 2008-01-31 | 2008-12-31 | 航天东方红卫星有限公司 | Satellite bias solar panel adapting for local times of different southbound nodes |
| CN111232246A (en) * | 2020-01-14 | 2020-06-05 | 上海卫星工程研究所 | Overall optimization design method based on illumination conditions of inclined orbit satellite |
| CN111792058A (en) * | 2020-06-28 | 2020-10-20 | 深圳航天东方红海特卫星有限公司 | Method and system for driving solar wing to face sun by low-inclination-angle track single-axis SADA |
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- 2022-06-22 CN CN202210715376.5A patent/CN115180179A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0544241A1 (en) * | 1991-11-27 | 1993-06-02 | Hughes Aircraft Company | Method and apparatus for dynamic precompensation of solar wing stepping motions of a satellite |
| US20060038080A1 (en) * | 2002-12-12 | 2006-02-23 | Bernard Polle | Solar control method for spacecraft |
| CN201172484Y (en) * | 2008-01-31 | 2008-12-31 | 航天东方红卫星有限公司 | Satellite bias solar panel adapting for local times of different southbound nodes |
| CN111232246A (en) * | 2020-01-14 | 2020-06-05 | 上海卫星工程研究所 | Overall optimization design method based on illumination conditions of inclined orbit satellite |
| CN111792058A (en) * | 2020-06-28 | 2020-10-20 | 深圳航天东方红海特卫星有限公司 | Method and system for driving solar wing to face sun by low-inclination-angle track single-axis SADA |
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