Offset sun alignment method suitable for DRO satellite measurement and control constraint, electronic equipment and medium

Abstract

The invention discloses an offset sun alignment method suitable for DRO satellite measurement and control constraint, which comprises the following steps: firstly, determining a maximum allowable offset angle phi, then determining an offset sun vector S'b based on the size of a sun-earth included angle Ase and the maximum allowable offset angle phi, determining attitude angles # imgabs0 # theta in X-axis and Y-axis directions based on the offset sun vector S 'b, and finally, determining the attitude angles # imgabs1 # theta in the X-axis and Y-axis directions based on the attitude angles # imgabs1 # theta in the X-axis and Y-axis directions. And the attitude of the DRO satellite is controlled, so that the sun-pointing attitude angle is [0, 0, 0], and offset sun-pointing is realized. The method can reduce the maneuvering time of a single machine and reduce the satellite attitude adjustment frequency.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular to a biased solar-focusing method, electronic equipment, and medium applicable to DRO satellite measurement and control constraints. Background Art

[0002] DRO satellites operate in a Distant Retrograde Orbit (DRO). DRO orbits are located near the gravitational equilibrium point between the Earth and the Moon, providing excellent orbital stability. Furthermore, DRO allows spacecraft to easily travel to Earth, deep space, and the Moon. Therefore, DRO satellites are currently being used for space science exploration and as deep space probe transit.

[0003] When a DRO-orbit satellite is oriented toward the Sun, due to the characteristics of the DRO orbit and the constraints of the antenna beam angle, the communication antenna cannot reach the ground, making ground-based measurement and control impossible. Therefore, to meet the multiple constraints of energy and measurement and control, a time-sharing scheduling method is currently used on-orbit. By rotating time slices, energy optimization and communication needs are alternately met. However, this method places high demands on attitude maneuvers, requiring the satellite to make multiple attitude adjustments. Alternatively, a multi-dimensional SADA can be installed on the satellite to separately drive the sail panels toward the Sun, but this method places high demands on the SADA configuration. Summary of the Invention

[0004] In order to address some or all of the problems in the prior art, the present invention provides, in a first aspect, a biased solar tracking method applicable to DRO satellite tracking and control constraints, comprising:

[0005] Determine the maximum allowable offset angle;

[0006] Determining a biased solar vector based on the angle between the sun and the earth and the maximum allowable bias angle;

[0007] Determining attitude angles in the X-axis and Y-axis directions based on the biased sun vector; and

[0008] Based on the X-axis and Y-axis attitude angles, the DRO satellite attitude is controlled so that the attitude angle towards the sun is [0, 0, 0].

[0009] Furthermore, the maximum allowable offset angle is determined according to energy balance constraints and the output current of the windsurfing panel.

[0010] Furthermore, determining the offset solar vector based on the angle between the sun and the earth and the maximum allowable offset angle includes:

[0011] If the sun-earth angle is less than or equal to 90°, rotating the projection of the sun vector in the present system around the normal of a first plane by a first angle, wherein the first plane is a plane where the projection of the geocentric vector and the sun vector in the present system are located, and the first angle is equal to the maximum allowable offset angle; and

[0012] If the sun-earth angle is greater than or equal to 90°, the projection of the sun vector in this system is rotated around the normal of the first plane by a second angle, wherein the second angle is equal to the opposite value of the maximum allowable offset angle.

[0013] Furthermore, the sun-earth angle is equal to the angle between the geocentric vector and the solar vector projected in this system.

[0014] Furthermore, the solar vector is determined according to the Julian century and the position of the inertial system.

[0015] Further, based on the biased sun vector S' b =[S' bx S' by S' bz Determining the X-axis and Y-axis orientation angles includes:

[0016] If the Y-axis coordinate value of the offset sun vector S' by If it is greater than or equal to 0, the attitude angle in the X-axis direction Calculated according to the following formula:

[0017]

[0018] If the Y-axis coordinate value of the offset sun vector S' by If it is less than 0, the X-axis attitude angle Calculated according to the following formula:

[0019]

[0020] If the X-axis coordinate value of the offset sun vector S' bx If it is greater than or equal to 0, the Y-axis attitude angle θ is calculated according to the following formula:

[0021] θ=π-atan2(S' bx ,S' bz ),

[0022] If the X-axis coordinate value of the offset sun vector S' bx If it is less than 0, the Y-axis attitude angle θ is calculated according to the following formula:

[0023] θ=-π-atan2(S' bx ,S' bz ).

[0024] Based on the biased solar method as described above, a second aspect of the present invention provides an electronic device for controlling the biased solar positioning of a DRO satellite, which includes a memory and a processor, wherein the memory is configured to store a computer program, and the computer program executes the biased solar method as described above when the processor is running.

[0025] A third aspect of the present invention further provides a computer-readable storage medium for controlling a DRO satellite's biased solar alignment, which stores a computer program. When the computer program is run on a processor, the biased solar alignment method as described above is executed.

[0026] This invention provides a biased solar alignment method suitable for DRO satellite tracking and control constraints. This method ensures that DRO satellites meet ground communication and tracking requirements when tracking the sun without significantly sacrificing sailboard conversion energy. This method selects the optimal attitude bias path, effectively reducing maneuvering time. Furthermore, it rationally plans the bias angle, reducing the frequency of attitude adjustments and avoiding frequent satellite adjustments. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] To further illustrate the above and other advantages and features of various embodiments of the present invention, a more detailed description of various embodiments of the present invention will be presented with reference to the accompanying drawings. It will be understood that these drawings depict only typical embodiments of the present invention and are not to be considered as limiting the scope thereof. In the drawings, for clarity, identical or corresponding components will be represented by the same or similar reference numerals.

[0028] Figure 1 Shows a schematic diagram of the DRO satellite measurement and control blind area;

[0029] Figure 2a and Figure 2b A schematic diagram showing the DRO satellite bias to the sun is shown;

[0030] Figure 3 A schematic flow chart illustrating a biased solar alignment method applicable to DRO satellite tracking and control constraints according to an embodiment of the present invention is shown;

[0031] Figure 4 A schematic diagram illustrating a process of determining a biased sun vector according to an embodiment of the present invention is shown;

[0032] Figure 5 A schematic diagram showing changes in the three-axis attitude angles under bias control using a biased solar tracking method applicable to DRO satellite measurement and control constraints according to one embodiment of the present invention is shown;

[0033] Figure 6 A schematic diagram showing changes in the three-axis attitude angular velocity under bias control using a biased solar tracking method applicable to DRO satellite measurement and control constraints according to one embodiment of the present invention is shown;

[0034] Figure 7 A schematic diagram showing the change in the angle between the sun vector and the -Z axis of the system under bias control of a biased solar method applicable to DRO satellite tracking and control constraints according to one embodiment of the present invention is shown; and

[0035] Figure 8 A schematic diagram showing the change in the angle between the earth vector and the -Z axis of the system under bias control of a biased solar method applicable to DRO satellite tracking and control constraints according to an embodiment of the present invention is shown. DETAILED DESCRIPTION

[0036] In the following description, the present invention is described with reference to various embodiments. However, those skilled in the art will recognize that the various embodiments can be implemented without one or more of the specific details or with other alternative and / or additional methods or components. In other cases, well-known structures or operations are not shown or described in detail to avoid obscuring the inventive aspects of the present invention. Similarly, for the purpose of explanation, specific numbers and configurations are set forth in order to provide a comprehensive understanding of the embodiments of the present invention. However, the present invention is not limited to these specific details. In addition, it should be understood that the various embodiments shown in the drawings are illustrative representations and are not necessarily drawn to scale.

[0037] In this specification, reference to "one embodiment" or "the embodiment" means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the present invention. The appearances of the phrase "in one embodiment" in various places in this specification are not necessarily all referring to the same embodiment.

[0038] It should be noted that the embodiments of the present invention describe the method steps in a specific order, but this is only for the purpose of illustrating the specific embodiment and does not limit the order of the steps. On the contrary, in different embodiments of the present invention, the order of the steps can be adjusted according to actual needs.

[0039] The inventors have found through research that when a satellite is orientated towards the sun, its -Z axis usually points to the sun in order to ensure that the sailboard is facing the sun. However, satellite communication antennas are generally installed on the ±Z axis. During the satellite's orientation towards the sun, ground measurement and control constraints must be met. Figure 1 As shown, only when the Earth vector falls in region 101 relative to the Sun vector in the plane where the Sun vector and the Earth vector are located, the satellite points to the Sun vector. At this time, there is no bias and the tracking and control antenna can cover the Earth. However, when the Sun-Earth angle is between α and 180-α, that is, when the Earth vector falls in region 102 relative to the Sun vector in the plane where the Sun vector and the Earth vector are located, if the satellite is oriented toward the Sun so that the -Z axis points to the Sun, then the communication antenna cannot cover the Earth and the satellite needs to be maneuvered to achieve satellite bias toward the Sun, where α is the satellite antenna beam angle.

[0040] The inventors further discovered that, based on the energy constraints of the entire satellite, while not significantly sacrificing the solar-direction performance, the angle between the satellite's desired Earth axis and the Earth's direction is ensured to be smaller than the antenna beam angle, that is, while ensuring a small -Z-axis angle deviation from the Sun, the antenna can complete Earth coverage. The offset angle φ can be selected according to the following strategy, where the offset angle φ is greater than 0: Figure 2a As shown, if the earth vector is located in region 201, it rotates around the normal of the plane where the sun vector and the earth vector are located, so that the sun vector points to Figure 2a The dotted line in the figure can ensure that the satellite is facing the sun while the antenna covers the earth. Figure 2b As shown, if the earth vector is located in region 202, it rotates around the normal of the plane where the sun vector and the earth vector are located, so that the sun vector points to Figure 2b The dotted line in the figure can ensure that the satellite is facing the sun while the antenna covers the earth.

[0041] However, in actual operation, in addition to considering energy constraints, the panel output current must also be considered, that is, selecting an appropriate offset angle under the dual constraints. In one embodiment of the present invention, the maximum light incident angle, i.e., the maximum offset angle φ, under the energy balance constraint of the entire satellite is determined by combining the energy balance constraint and the panel output current. According to the principle of solar panel power generation, its output voltage is a constant value, and the power attenuation characteristic is an analysis of the output current. Therefore, the panel output current formula is:

[0042] P=S*Ki*i0*cosδ,

[0043] Among them, S is the sail panel area, Ki is the current conversion coefficient, i0 is the solar constant, and δ is the light incident angle. Based on this, it can be obtained that when the sail panel area is fixed, due to the orbital characteristics, the current conversion coefficient and the solar constant are also relatively fixed. In order not to significantly sacrifice the solar orientation efficiency and ensure the energy supply of the entire satellite, it is necessary to minimize the light incident angle, that is, in the case of offset solar orientation, the angle between the selected offset solar vector and the actual solar vector is as small as possible. At the same time, in order to ensure the balance of energy demand for each orbit, the sail panel power demand is usually split into battery charging power and whole satellite discharge power during design. At the same time, since the sail panel efficiency will decay to a certain extent with the time on orbit, the design retains a power margin of 5% to 10% to meet the energy balance at the end of its life, that is, the energy balance constraint must meet the following conditions:

[0044] P=P1+P2+P0,

[0045] Among them, P is the required charging power of the sailboard, P1 is the required charging power of the battery, P2 is the discharge power of the entire satellite, and P0 is the energy margin design.

[0046] Based on the above-mentioned bias angle selection method, the present invention further provides a satellite bias-to-sun method suitable for DRO orbit, which can solve the problems of long cycle, long distance and insufficient communication capability of deep space exploration missions.

[0047] Figure 3 FIG. 1 is a flow chart showing a method for biasing the sun with DRO satellite tracking and control constraints according to an embodiment of the present invention. Figure 3 As shown, a biased solar alignment method applicable to DRO satellite tracking and control constraints includes:

[0048] First, in step 301, the maximum allowable offset angle is determined. According to the aforementioned method, under dual constraints, the maximum allowable offset angle φ is determined, i.e., the energy balance constraint and the panel output current formula are combined to obtain the maximum light incident angle under the energy balance constraint of the entire satellite, which is used as the maximum allowable offset angle φ.

[0049] Next, in step 302, the biased sun vector is determined. se The magnitude of and the maximum allowable offset angle φ determine the offset sun vector S' b In one embodiment of the present invention, in addition to satisfying the double-constrained paranoid sun pointing, in order to ensure a smooth transition of the satellite attitude, the shortest path planning is also used to complete the satellite attitude rotation pointing. Figure 4 A schematic diagram of a process for determining a biased solar vector according to an embodiment of the present invention is shown as follows: Figure 4 As shown, in one embodiment of the present invention, determining the biased sun vector includes:

[0050] First, determine the inertial solar vector. The inertial solar vector is determined based on the Julian century T:

[0051]

[0052] in,

[0053] L is the geometric mean ecliptic longitude of the mean equinox, which is calculated based on the current Julian century T and the mean anomaly M: L = 4.8951 + 628.3076*T + 0.03342*sin(M), where the mean anomaly M is determined based on the current Julian century T: M = 6.2401 + 628.30196*T; and

[0054] EPS is the obliquity of the mean ecliptic, and its value is 0.4090928;

[0055] At the same time, determine the inertial system center vector. i The inertial system position can be calculated by conventional technical means in this field, which will not be described in detail here;

[0056] Next, determine the projection of this system. Get the position vector r of the satellite inertial system i , calculate the geocentric vector e i and the sun vector S i Projection e in this system b 、S b :

[0057]

[0058] Among them, R bi is the conversion matrix between the inertial system and the satellite system, which is based on the inertial system quaternion q i =q1i+q2j+q3k+q4 Determine:

[0059]

[0060] in,

[0061]

[0062] l 21 =2(q1q2-q3q4),

[0063] l 31 =2(q1q3+q2q4),

[0064] l 12 =2(q1q2+q3q4),

[0065]

[0066] l 32 =2(q2q3-q1q4),

[0067] l 13 =2(q1q3-q2q4),

[0068] l 23 =2(q2q3+q1q4),

[0069]

[0070] Next, determine the angle between the sun and the earth. se The magnitude is equal to the geocentric vector e i Projection e in this system b With the sun vector S i Projection S in this system b Angle:

[0071] A se =acos(dot(e b ,S b));as well as

[0072] Finally, the offset sun vector is determined. se The size of the offset angle and the maximum allowable offset angle φ can determine the offset sun vector S' b ,include:

[0073] If the angle between the sun and the earth is A se Less than or equal to 90°, then the sun vector S i Projection S in this system b Rotate the first plane by a first angle around the normal N of the first plane, wherein the first plane is the center of the earth vector e i Projection under this system b With the sun vector S i Projection S in this system b The plane where it is located has a normal N = cross (S b ,e b ), and the first angle is equal to the maximum allowable offset angle φ; and

[0074] If the angle between the sun and the earth is A se If it is greater than or equal to 90°, the sun vector S i Projection S in this system b Rotating a second angle around the normal N of the first plane, wherein the second angle is equal to the inverse of the maximum allowable offset angle φ;

[0075] Next, in step 303, the attitude angles in the X-axis and Y-axis directions are determined. b Determine the attitude angles in the X-axis and Y-axis directions θ, the offset solar angle pointing to the -Z axis In one embodiment of the present invention, based on the biased solar vector S' b =[S' bx S' by S' bz ]Determine the attitude angles in the X and Y axes θ includes:

[0076] If the Y-axis coordinate value of the offset sun vector S' by If it is greater than or equal to 0, the attitude angle in the X-axis direction Calculated according to the following formula:

[0077]

[0078] If the Y-axis coordinate value of the offset sun vector S' by If it is less than 0, the X-axis attitude angle Calculated according to the following formula:

[0079]

[0080] If the X-axis coordinate value of the offset sun vector S' bx If it is greater than or equal to 0, the Y-axis attitude angle θ is calculated according to the following formula:

[0081] θ=π-atan2(S' bx ,S' bz ),

[0082] If the X-axis coordinate value of the offset sun vector S' bx If it is less than 0, the Y-axis attitude angle θ is calculated according to the following formula:

[0083] θ=-π-atan2(S' bx ,S' bz );as well as

[0084] Finally, in step 304, the sun is offset. Based on the X-axis and Y-axis attitude angles θ, controls the DRO satellite attitude so that the solar attitude angle is [0, 0, 0]. Since the -Z axis needs to remain facing the sun, the Z axis direction is not controlled. At this point, the satellite -Z axis can be pointed in the direction of the offset solar vector.

[0085] Figures 5 to 8 Schematic diagrams showing the changes in the three-axis attitude angle, the three-axis attitude angular velocity, the angle between the sun vector of the system and the -Z axis, and the angle between the earth vector of the system and the -Z axis under bias control of a biased solar method suitable for DRO satellite measurement and control constraints based on an embodiment of the present invention are shown respectively, wherein the orbit of the satellite is a DRO orbit (2:1), the sun-satellite-earth angle varies between 60° and 120°, the satellite antenna beam angle is 55°, and the effective area of ​​the satellite sailboard is 2.25m 2 The total satellite power consumption is 551W. First, by considering the satellite beam angle limitations and the sailboard area, based on the offset angle selection method described above, it is possible to calculate that the maximum allowable satellite offset angle is 30°. Then, based on the offset solar orientation method described above, satellite attitude is determined using star sensors, satellite angular velocity is determined using a gyroscope, and offset solar orientation is achieved in conjunction with flywheel PD control. It can be seen that the satellite is oriented toward the sun while meeting the antenna coverage of the Earth and satisfying energy requirements.

[0086] Based on the above-mentioned biased solar tracking method, the present invention further provides a DRO satellite, which adopts the above-mentioned biased solar tracking method to perform biased solar tracking.

[0087] The present invention also provides an electronic device for controlling the biased solar alignment of a DRO satellite, which includes a memory and a processor, wherein the memory is configured to store a computer program, and the computer program executes the biased solar alignment method as described above when the processor is running.

[0088] The present invention also provides a computer-readable storage medium for controlling a DRO satellite to bias the sun, which stores a computer program. When the computer program is run on a processor, the biased sun method as described above is executed.

[0089] Although various embodiments of the present invention have been described above, it should be understood that they are presented by way of example only and not limitation. It will be apparent to those skilled in the relevant art that various combinations, modifications, and variations may be made thereto without departing from the spirit and scope of the present invention. Therefore, the breadth and scope of the present invention disclosed herein should not be limited by the exemplary embodiments disclosed above, but should be defined solely in accordance with the appended claims and their equivalents.

Claims

1. A biased solar tracking method suitable for DRO satellite tracking and control constraints, characterized in that: include: Determine the maximum allowable offset angle φ; Based on the sun-earth angle A se The magnitude of and the maximum allowable offset angle φ determine the offset sun vector S' b ; Based on the biased sun vector S' b Determine the attitude angles in the X-axis and Y-axis directions θ; and Based on the X-axis and Y-axis attitude angles θ, controls the DRO satellite attitude so that the attitude angle to the sun is [0, 0, 0].

2. The biased sun method according to claim 1, wherein: The maximum allowable offset angle φ is determined according to the energy balance constraint and the output current of the windsurfing panel.

3. The biased sun method according to claim 1, wherein: The sun-earth angle A se Equal to the geocentric vector e i Projection e in this system b With the sun vector S i Projection S in this system b Angle: A se =acos(dot(e b ,S b ))。 4. The biased sun method according to claim 1, wherein: The angle A between the sun and the earth se The magnitude of and the maximum allowable offset angle φ determine the offset sun vector S' b include: If the angle between the sun and the earth is A se Less than or equal to 90°, then the sun vector S i Projection S in this system b Rotate the first plane by a first angle around the normal N of the first plane, wherein the first plane is the center of the earth vector e i Projection under this system b With the sun vector S i Projection S in this system b The plane where the first angle is located is equal to the maximum allowable offset angle φ; and If the angle between the sun and the earth is A se If it is greater than or equal to 90°, the sun vector S i Projection S in this system b A second angle is rotated around the normal N of the first plane, wherein the second angle is equal to the inverse of the maximum allowable offset angle φ.

5. The biased sun method according to claim 3, wherein: The sun vector S i Determined based on the Julian century T and the position of the inertial system: Wherein, L is the geometric mean ecliptic longitude of the mean vernal equinox, eps is the mean ecliptic obliquity, and the value of eps is 0.4090928, L = 4.8951 + 628.3076*T + 0.03342*sin(M), where M is the mean anomaly, and M = 6.2401 + 628.30196*T.

6. The biased sun-facing method according to claim 5, wherein: The geocentric vector e i Projection under this system b With the sun vector S i Projection S in this system b Calculated according to the following formula: Among them, r i is the satellite inertial system position vector, R bi It is the transformation matrix between the inertial system and the satellite system, which is determined according to the inertial system quaternion.

7. The biased sun-facing method according to claim 1, wherein: Based on the biased sun vector S' b =[S′ bx S′ by S′ bz ]Determine the attitude angles in the X and Y axes θ includes: If the Y-axis coordinate value S′ of the offset sun vector by If it is greater than or equal to 0, the attitude angle in the X-axis direction Calculated according to the following formula: If the Y-axis coordinate value S′ of the offset sun vector by If it is less than 0, the X-axis attitude angle Calculated according to the following formula: If the X-axis coordinate value S′ of the offset sun vector bx If it is greater than or equal to 0, the Y-axis attitude angle θ is calculated according to the following formula: θ=π-atan2(S′ bx ,S′ bz ), If the X-axis coordinate value S′ of the offset sun vector bx If it is less than 0, the Y-axis attitude angle θ is calculated according to the following formula: θ=-π-atan2(S′ bx ,S′ bz )。 8. An electronic device for controlling the sun-biasing of a DRO satellite, characterized in that: The device comprises a memory and a processor, wherein the memory is configured to store a computer program, and when the computer program is executed by the processor, the biased sun method according to any one of claims 1 to 7 is executed.

9. A computer-readable storage medium for controlling a DRO satellite's sun bias, characterized in that: A computer program is stored, and when the computer program is run on a processor, the biased sun method according to any one of claims 1 to 7 is executed.

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