A self-balancing aerodynamic layout of solar panels for low-orbit satellites

Abstract

The present invention discloses a self-balancing low-orbit satellite solar wing aerodynamic layout, including a satellite body and two groups of self-balancing solar wings; the satellite body is a rectangular parallelepiped, and the two groups of self-balancing solar wings are respectively located on both sides of the satellite body to form a symmetrical structure, and the self-balancing solar wing includes a first folding solar wing, which is connected to the satellite body to form an adjustable deflection angle, and the deflection angle is adjusted to achieve the aerodynamic moment balance of the low-orbit satellite. The present invention can eliminate the aerodynamic interference torque to which the satellite is subjected when operating in a low-orbit near-Earth orbit, greatly reduce the influence of the thin atmosphere in the near-Earth space on the satellite orbit, and also reduce the aerodynamic interference torque of the satellite attitude control, while being able to meet the requirements of the side heat protection system of the satellite body, thereby improving the satellite reliability.

Description

Technical Field

[0001] The invention relates to a self-balancing low-orbit satellite solar wing aerodynamic layout, belonging to the technical field of satellite aerodynamic balancing. Background Art

[0002] The low earth orbit (LEO) environment, 200 to 1000 km from the ground, is the operating area of ​​earth observation satellites. Satellites operating in this earth orbit are low-orbit satellites. Ultra-low-orbit satellites have low orbital altitudes and can quickly reach the predetermined orbit and start work; their orbital period is short, and they can conduct reconnaissance and observation of the same target in a very short time. The reconnaissance work is more frequent than traditional satellites, and more information is obtained; and the indicators of the space cameras or radars carried by them do not need to be too high, and they can achieve the effect of traditional satellites, or even better. The payload alone can greatly reduce the cost of satellites; the return point of ultra-low-orbit satellites can also be obtained with relatively high accuracy. Therefore, the research and design of ultra-low-orbit satellites can produce huge social and economic effects and have great engineering application value. Ultra-low-orbit satellites are becoming the focus of attention of major aerospace powers, and famous universities and research institutes have also begun to carry out research on low-orbit satellites.

[0003] However, in ultra-low orbit environments, especially at altitudes of 200-300km, the thin atmosphere in the near-Earth space will cause greater aerodynamic drag (the atmospheric density near an altitude of 200km is more than 1000 times that of an altitude of 700km). This drag can cause Earth observation satellites to quickly deviate from their target orbits, requiring speed compensation and frequent orbital maneuvers. At the same time, if the satellite configuration is asymmetric or the center of mass deviates too much from the satellite's centroid, the thin atmosphere near the Earth will also produce a large aerodynamic disturbance torque on the satellite's attitude control, especially for near-Earth satellites equipped with single-wing solar panels for two-dimensional sun-oriented control. The aerodynamic disturbance torque amplitude is very large, up to 100g·cm. Although the force of the thin atmosphere near the Earth is very small compared to the Earth's gravity, this force accumulates during the satellite's operating time, and the long-term perturbations have a huge impact on the satellite's orbit and attitude, causing the satellite's actuators to frequently unload.

[0004] Normally, the control part of the satellite always uses the aerodynamic torque as the interference torque and uses the reaction flywheel or jet to balance it. For example, the Japan Aerospace Exploration Agency (JAXA) launched the world's first earth observation satellite SLATS (Super Low Altitude Test Satellite) in 2017, which operates in an ultra-low orbit below 300km (such as Figure 1Its orbit eventually dropped to an altitude of 268km to 150km. In order to overcome the atmospheric resistance in space, it used the world's first ion engine to maintain the predetermined orbit. The solar panels used by the satellite are arranged horizontally and side by side, as shown in Figure 1. Figure 1 As shown, this satellite design will be affected by a large aerodynamic disturbance torque when operating in a low-Earth orbit, which will affect the stability of the satellite's attitude and pointing accuracy, and generate large energy consumption.

[0005] Therefore, if measures are taken to reduce or eliminate the aerodynamic disturbance torque on the satellite when it is operating in a low-Earth orbit, not only can sufficient attitude stability and attitude pointing accuracy be obtained, but also the satellite energy consumption can be effectively reduced, thereby reducing the cost of satellite development. It has great economic value and provides a new idea and strategy for the research on control systems for future low-cost and low-power satellites. Summary of the invention

[0006] The technical problem solved by the present invention is: to overcome the shortcomings of the existing satellite aerodynamic balancing technology, and to provide a self-balancing low-orbit satellite solar wing aerodynamic layout, which can effectively reduce or even eliminate the aerodynamic torque of the low-orbit satellite, while ensuring the requirements of the satellite body side thermal protection system, thereby improving the reliability of the satellite control system.

[0007] The technical solution of the present invention is:

[0008] A self-balancing low-orbit satellite solar wing aerodynamic layout comprises a satellite body and two groups of self-balancing solar wings; the satellite body is a rectangular parallelepiped, the two groups of self-balancing solar wings are respectively located on both sides of the satellite body to form a symmetrical structure, the self-balancing solar wings comprise a first folding solar wing, the first folding solar wing is connected to the satellite body to form an adjustable deflection angle, and the deflection angle is adjusted to achieve aerodynamic moment balance of the low-orbit satellite.

[0009] In the above-mentioned solar wing aerodynamic layout, with the connection line between the first folded solar wing and the satellite body as the rotation axis, the first folded solar wing deflects downward according to aerodynamic balancing to form the deflection angle θ along the extension surface of the upper surface of the satellite body.

[0010] In the above-mentioned solar wing aerodynamic layout, the self-balancing solar wing also 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 with their sides connected sequentially. The second folding solar wing and the third folding solar wing are on the same horizontal plane.

[0011] In the above-mentioned aerodynamic layout of the solar wing, the first folding solar wing, the second folding solar wing and the third folding solar wing have the same shape, and the gap size Lp between the first folding solar wing and the second folding solar wing is the same as the gap size between the second folding solar wing and the third folding solar wing.

[0012] In the above-mentioned solar wing aerodynamic layout, the length L of the self-balancing solar wing is 3 to 5 times the width Lw.

[0013] In the above-mentioned solar wing aerodynamic layout, the length of the satellite body is equal to the length L of the self-balancing solar wing, the length L is 4 to 8 times the height H, and the height H is 0.8 to 1.5 times the width W.

[0014] In the above-mentioned solar wing aerodynamic layout, the deflection angle θ ranges from 0° to 50°.

[0015] In the above-mentioned solar wing aerodynamic layout, the method for adjusting the deflection angle to achieve aerodynamic torque balance of low-orbit satellites comprises the following specific steps:

[0016] (1) Obtain the satellite reference center of mass coordinates (x g ,y g 、z g );

[0017] (2) Meshing the satellite body and solar wing surfaces to generate grid cells, and outputting the normal and position parameters of each grid cell surface;

[0018] (3) Based on the satellite flight altitude and flight date, select the solar activity parameter F10.7 and the geomagnetic parameter Ap;

[0019] (4) Based on the solar activity parameter F10.7 and the geomagnetic parameter Ap in step (3), the atmospheric density ρ of the satellite flight environment is calculated using the atmospheric model;

[0020] (5) According to the normal and position parameters of the surface of each grid unit in step (2) and the atmospheric density ρ in step (4), a test particle is selected, and the interaction between the gas molecules and the entire satellite surface is simulated by using the satellite aerodynamic fast calculation method TPMC, and the momentum exchange amount after the test particle collides with the satellite surface is calculated;

[0021] (6) According to the satellite reference centroid coordinates (x g ,y g 、z g ) and the value of the deflection angle θ, and calculate the aerodynamic force and torque of the entire satellite according to the momentum exchange amount in step (5);

[0022] (7) According to the aerodynamic force and torque of the whole satellite, adjust the value of the deflection angle θ, and repeat step (6) until the aerodynamic torque of the whole satellite is 0, and obtain the value of the deflection angle θ at this time.

[0023] In the above solar wing aerodynamic layout, the calculation formula for calculating the aerodynamic force and torque of the entire satellite in step (6) is:

[0024]

[0025] Where F is the aerodynamic force of the entire satellite, M is the moment of the entire satellite, and N is E is the total number of grid cells on the entire satellite surface, F l 、M l are the aerodynamic force and moment on the lth grid unit respectively.

[0026] The aerodynamic force F on the lth grid unit l and moment M l The calculation formula is:

[0027]

[0028] Where N is the total number of collisions between all test particles and the satellite surface, r j , ΔP j are the vector and momentum exchange from the satellite reference center of mass to the collision position during the jth collision; A is the weight factor, which indicates the number of real gas molecules represented by one test particle.

[0029] In the above solar wing aerodynamic layout, the calculation formula of the vector from the satellite reference center of mass to the collision position during the j-th collision is:

[0030]

[0031] Where θ is the deflection angle, (x g ,y g , z g ) is the reference centroid coordinate of the satellite, (x l ,y l , z l ) is the collision position coordinate, i, j, k are the unit vectors in x, y, z directions respectively.

[0032] In the above solar wing aerodynamic layout, the calculation formula for the momentum exchange amount from the satellite reference center of mass to the collision position during the j-th collision is:

[0033] ΔP j =mc i -mc r

[0034] In the formula, c i is the velocity vector of the incident molecule of the test particle; c r is the molecular velocity vector of the complete diffuse reflection of the test particle, and m is the molecular mass of the incoming gas in the test particle.

[0035] The advantages of the present invention compared with the prior art are:

[0036] (1) The present invention can eliminate the aerodynamic interference torque of the rarefied atmosphere on the satellite by rotating the first fold of the multi-fold solar wing. Compared with the existing horizontal solar wing non-deflection design, it can not only obtain sufficient attitude stability and attitude pointing accuracy, but also effectively reduce the energy consumption of the satellite, thereby reducing the cost of satellite development, and has great economic value.

[0037] (2) The self-balancing solar wing design of the present invention only changes the first fold angle of the multi-fold solar wing, which has little impact on the design appearance of the satellite and is more reliable and simple to operate.

[0038] (3) The self-balancing solar wing design of the present invention only changes the deflection angle of the first fold of the multi-fold solar wing, while the other solar wings remain in a horizontal state, which can ensure the heat dissipation requirements of the satellite side, thereby reducing the requirements for the thermal protection system. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1 This is a schematic diagram of SLATS, a low-orbit satellite of Japan's JAXA;

[0040] Figure 2 A schematic diagram of the aerodynamic layout of the self-balancing low-orbit satellite solar wing of the present invention;

[0041] Figure 3 The self-balancing solar wing low-orbit satellite surface grid of the present invention;

[0042] Figure 4 The flow chart of aerodynamic moment balance design of the present invention is as follows;

[0043] Figure 5 This is the aerodynamic balancing characteristic diagram of the self-balancing solar wing low-orbit satellite in the pitch direction of 250km. Specific implementation plan

[0044] The specific implementation scheme of the present invention is described below with reference to examples.

[0045] like Figure 2 The figure shows a schematic diagram of the aerodynamic layout of the self-balancing low-orbit satellite solar wing of the present invention, wherein the self-balancing low-orbit satellite solar wing aerodynamic layout comprises a satellite body 1 and two groups of self-balancing solar wings 2; the satellite body 1 is a rectangular parallelepiped, and the two groups of self-balancing solar wings 2 are respectively located on both sides of the satellite body 1 to form a symmetrical structure, and the self-balancing solar wings 2 comprise a first folding solar wing 21, which is connected to the satellite body 1 to form an adjustable deflection angle, and the deflection angle is adjusted to achieve the aerodynamic moment balance of the low-orbit satellite.

[0046] With the connecting line between the first folded solar wing 21 and the satellite body 1 as the rotation axis, the first folded solar wing 21 deflects downward according to aerodynamic balancing to form a deflection angle θ along the extension surface of the upper surface of the satellite body 1. The value range of the deflection angle θ is 0° to 50°.

[0047] The self-balancing solar wing 2 also includes 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 all flat plate structures, and the sides are connected in sequence. The second folding solar wing 22 and the third folding solar wing 23 are on the same horizontal plane. The first folding solar wing 21, the second folding solar wing 22 and the third folding solar wing 23 have the same shape, and the gap size Lp between the first folding solar wing 21 and the second folding solar wing 22 is the same as the gap size between the second folding solar wing 22 and the third folding solar wing 23. The length L of the self-balancing 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-balancing solar wing 2, the length L is 4 to 8 times the height H, and the height H is 0.8 to 1.5 times the width W.

[0048] like Figure 4 The figure shows the aerodynamic moment balance design flow chart of the present invention, and the method of adjusting the deflection angle to achieve the aerodynamic moment balance of the low-orbit satellite, the specific steps are:

[0049] (1) Obtain the satellite reference center of mass coordinates (x g ,y g 、z g );

[0050] (2) Figure 3 The figure shows the surface grid of the self-balancing solar wing low-orbit satellite of the present invention, the satellite body and the solar wing surface are gridded to generate grid cells, and the normal and position parameters of each grid cell surface are output;

[0051] (3) Based on the satellite flight altitude and flight date, select the solar activity parameter F10.7 and the geomagnetic parameter Ap;

[0052] (4) Based on the solar activity parameter F10.7 and the geomagnetic parameter Ap in step 3, the atmospheric density ρ of the satellite flight environment is calculated using the NRLMSISE-00 atmospheric model;

[0053] (5) According to the normal and position parameters of each grid unit surface in step 2 and the atmospheric density ρ in step 4, select the test particle, use the satellite aerodynamic fast calculation method TPMC (Test Particle Monte Carlo), simulate the interaction between gas molecules and the entire satellite surface, and calculate the momentum exchange amount after the test particle collides with the satellite surface;

[0054] (6) According to the satellite reference centroid coordinates (x g ,y g 、z g) and the value of the deflection angle θ, and calculate the aerodynamic force and torque of the entire satellite according to the momentum exchange in step 5;

[0055] (7) According to the aerodynamic force and torque of the whole satellite, adjust the value of the deflection angle θ, and repeat step 6 until the aerodynamic torque of the whole satellite is 0, and obtain the value of the deflection angle θ at this time.

[0056] The calculation formula for the aerodynamic force F and moment M of the entire satellite is:

[0057]

[0058] Where N E is the total number of grid cells on the entire satellite surface, F l 、M l are the aerodynamic force and moment on the lth grid unit respectively, and:

[0059]

[0060] Where N is the total number of collisions between all test particles and the satellite surface, r j , ΔP j are the vector and momentum exchange from the satellite reference center of mass to the collision position during the jth collision; A is the weight factor, which indicates the number of real gas molecules represented by one test particle.

[0061] The calculation formula for the vector from the satellite reference center of mass to the collision position at the jth collision is:

[0062]

[0063] Where θ is the deflection angle, (x g ,y g , z g ) is the reference center of mass coordinate of the satellite, (x l ,y l , z l ) is the collision position coordinate, i, j, k are the unit vectors in x, y, z directions respectively.

[0064] The calculation formula of the momentum exchange from the satellite reference center of mass to the collision position during the jth collision is:

[0065] ΔP j =mc i -mc r

[0066] In the formula, c i is the velocity vector of the incident molecule of the test particle; c r is the molecular velocity vector of the complete diffuse reflection of the test particle, and m is the molecular mass of the incoming gas in the test particle.

[0067] If the aerodynamic interference torque of the entire satellite does not reach the expected target value or is eliminated to zero, continue to optimize the rotation angle of the first folded solar wing, keep the second and third folded solar wings in translation, and ensure that the horizontal gap of the three folded solar wings does not change; if the aerodynamic interference torque of the entire satellite reaches the expected target value or is eliminated to zero, the solar wing is the optimized self-balancing solar wing layout.

[0068] The specific solution example of the present invention is as follows:

[0069] The incoming gas calculated in the example of the method of the present invention is air, such as Figure 2 As shown, in this embodiment, the size of the satellite body is: 2700mm×330mm×330mm; three solar cell wings are arranged on both sides of the satellite, with the size of: 2700mm×330mm×12mm, the satellite flight altitude is set to 250km, the solar activity parameter F10.7 takes the parameter 185.75 on April 28, 2037, and the geomagnetic parameter takes 4.5, as shown in FIG. Figure 5 The figure shows the aerodynamic balancing characteristics of the self-balancing solar wing low-orbit satellite at 250 km in pitch direction. Table 1 shows the aerodynamic balancing characteristics of the satellite at 250 km in pitch direction.

[0070] Table 1

[0071]

[0072] It can be seen from the table that by optimizing the deflection angle of the first folded solar wing, the aerodynamic interference torque in the pitch direction of the entire satellite can be eliminated, and the deflection angle of the solar wing is only 27°, which meets the heat dissipation requirements of the side of the satellite body.

[0073] The undisclosed technologies in the present invention are common knowledge to those skilled in the art.

Claims

1. A self-balancing low-orbit satellite solar wing aerodynamic layout, characterized by: The invention comprises a satellite body (1) and two groups of self-balancing solar wings (2); the satellite body (1) is a rectangular parallelepiped, the two groups of self-balancing solar wings (2) are respectively located on both sides of the satellite body (1) to form a symmetrical structure, the self-balancing solar wings (2) comprise a first folding solar wing (21), the first folding solar wing (21) is connected to the satellite body (1) to form an adjustable deflection angle, and the deflection angle is adjusted to achieve aerodynamic moment balance of the low-orbit satellite; The method for adjusting the deflection angle to achieve aerodynamic torque balance of a low-orbit satellite comprises the following specific steps: (1) Obtain the satellite reference center of mass coordinates (x g ,y g 、z g ); (2) Meshing the satellite body and solar wing surfaces to generate grid cells, and outputting the normal and position parameters of each grid cell surface; (3) Based on the satellite flight altitude and flight date, select the solar activity parameter F10.7 and the geomagnetic parameter Ap; (4) Based on the solar activity parameter F10.7 and the geomagnetic parameter Ap in step (3), the atmospheric density ρ of the satellite flight environment is calculated using the atmospheric model; (5) According to the normal and position parameters of the surface of each grid unit in step (2) and the atmospheric density ρ in step (4), a test particle is selected, and the interaction between the gas molecules and the entire satellite surface is simulated by using the satellite aerodynamic fast calculation method TPMC, and the momentum exchange amount after the test particle collides with the satellite surface is calculated; (6) According to the satellite reference centroid coordinates (x g ,y g 、z g ) and the value of the deflection angle θ, and calculate the aerodynamic force and torque of the entire satellite according to the momentum exchange amount in step (5); (7) According to the aerodynamic force and torque of the whole satellite, adjust the value of the deflection angle θ, and repeat step (6) until the aerodynamic torque of the whole satellite is 0, and obtain the value of the deflection angle θ at this time.

2. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 1 is characterized by: With the connection line between the first folding solar wing (21) and the satellite body (1) as the rotation axis, the first folding solar wing (21) deflects downward according to aerodynamic balancing to form the deflection angle θ along the extension surface of the upper surface of the satellite body (1).

3. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 1 is characterized by: The self-leveling 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 all flat plate structures, the sides of which are sequentially connected; the second folding solar wing (22) and the third folding solar wing (23) are on the same horizontal plane.

4. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 1 is characterized by: The first folding solar wing (21), the second folding solar wing (22) and the third folding solar wing (23) have the same shape, and the gap size Lp between the first folding solar wing (21) and the second folding solar wing (22) is the same as the gap size between the second folding solar wing (22) and the third folding solar wing (23).

5. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 1 is characterized by: The length L of the self-balancing solar wing (2) is 3 to 5 times the width Lw.

6. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 1 is characterized by: The length of the satellite body (1) is equal to the length L of the self-balancing solar wing (2), the length L is 4 to 8 times the height H, and the height H is 0.8 to 1.5 times the width W.

7. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 1 is characterized by: The deflection angle θ has a value range of 0° to 50°.

8. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 1 is characterized by: The calculation formula for calculating the aerodynamic force F and moment M of the entire satellite in step (6) is: Where F is the aerodynamic force of the entire satellite, M is the moment of the entire satellite, and N is E is the total number of grid cells on the entire satellite surface, F l 、M l are the aerodynamic force and moment on the lth grid unit respectively.

9. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 8 is characterized by: The aerodynamic force F on the lth grid unit l and moment M l The calculation formula is: Where N is the total number of collisions between all test particles and the satellite surface, r j , ΔP j are the vector and momentum exchange from the satellite reference center of mass to the collision position during the jth collision; A is the weight factor, which indicates the number of real gas molecules represented by one test particle.

10. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 9 is characterized by: The calculation formula of the vector from the satellite reference center of mass to the collision position during the j-th collision is: Where θ is the deflection angle, (x g ,y g , z g ) is the reference centroid coordinate of the satellite, (x l ,y l , z l ) is the collision position coordinate, i, j, k are the unit vectors in x, y, z directions respectively.

11. The self-balancing low-orbit satellite solar wing aerodynamic layout according to claim 10 is characterized by: The momentum exchange amount ΔP from the satellite reference mass center to the collision position during the jth collision j The calculation formula is: ΔP j =mc i -mc r In the formula, c i is the velocity vector of the incident molecule of the test particle; c r is the molecular velocity vector of the complete diffuse reflection of the test particle, and m is the molecular mass of the incoming gas in the test particle.

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