Satellite solar cell array deployment mechanism

Abstract

Disclosed is a satellite solar cell array deployment mechanism, comprising: a servo motor fixed on a satellite structural frame; a first fixed rod fixedly connected to the satellite structural frame; a first movable rod, a first end of which is pivotally connected to the first end of the first fixed rod; a second movable rod, a first end of which is pivotally connected to the second end of the first fixed rod; the second fixed rod is fixedly connected to the satellite solar cell array; a third movable rod, a first end of which is pivotally connected to the first end of the second fixed rod; a fourth movable rod, a first end of which is pivotally connected to the second end of the second fixed rod; wherein, the second end of the first movable rod is pivotally connected to the second end of the third movable rod; the second end of the second movable rod is pivotally connected to the second end of the fourth movable rod; the second movable rod is pivotally connected to the middle part of the third movable rod; the output shaft of the servo motor drives the first movable rod to rotate around the first pin shaft, thereby driving the satellite solar cell array to deploy and fold relative to the satellite structural frame.

Description

Technical Field

[0001] The present invention relates to the field of aerospace technology, and in particular to a satellite solar cell array deployment mechanism. Background Art

[0002] There are many ways to deploy satellite solar cell arrays, such as extension arms, memory alloys, hinges, etc. However, satellites of different specifications correspond to different deployment methods. Taking into account factors such as cost, reliability, and the overall structural deployment of the satellite, large satellites (generally weighing more than 500kg) usually use extension arms and memory alloys for deployment, while micro satellites (generally weighing 10kg to 100kg) usually use hinges for deployment.

[0003] Among the existing satellite solar cell array deployment methods, the extension arm deployment method has a complex structure, a large volume, and is difficult to design; the memory alloy deployment method has extremely high technical requirements for thermal control, and if there is an error in the control, the deployment will fail; the hinge deployment method has limited load-bearing capacity, the deployment angle is difficult to control, and the deployment process has a large impact.

[0004] Therefore, there is a need in the art for a satellite solar cell array deployment technology solution that can overcome the shortcomings of the prior art. Summary of the Invention

[0005] To this end, according to an embodiment of the present invention, a satellite solar cell array deployment mechanism is proposed, comprising:

[0006] Servo motor, fixed on the satellite structure frame 9;

[0007] A first fixing rod 4 is fixedly connected to the satellite structure frame 9;

[0008] A first movable rod 1, a first end of which is pivotally connected to a first end of a first fixed rod 4 via a first pin 21;

[0009] A second movable rod 3, a first end of which is pivotally connected to the second end of the first fixed rod 4 via a second pin 22;

[0010] A second fixing rod 6 is fixedly connected to the satellite solar cell array 8;

[0011] A third movable rod 5, a first end of which is pivotally connected to the first end of the second fixed rod 6 via a third pin 23;

[0012] a fourth movable rod 7, a first end of which is pivotally connected to the second end of the second fixed rod 6 via a fourth pin 24;

[0013] The second end of the first movable rod 1 is pivotally connected to the second end of the third movable rod 5 via a fifth pin 25;

[0014] The second end of the second movable rod 3 is pivotally connected to the second end of the fourth movable rod 7 via a sixth pin 26;

[0015] The middle parts of the second movable rod 3 and the third movable rod 5 are pivotally connected via a seventh pin 27;

[0016] The output shaft of the servo motor drives the first movable rod 1 to rotate around the first pin shaft 21 , thereby driving the satellite solar cell array 8 to unfold and fold relative to the satellite structure frame 9 .

[0017] The satellite solar cell array deployment mechanism involved in this patent has a simple structure, is stable and reliable, has a controllable deployment angle, has a small deployment impact, and is easy to install. It can ensure that the satellite solar cell array is deployed to a set angle at a constant angular velocity and maintained at a set position. The gap between the satellite solar cell array and the structural frame can be very small (under the same conditions, the solar cell array can be ensured to have a larger area), effectively avoiding interference with the satellite structural frame during the deployment process, resulting in a large part of the solar cell array being cut off, and also avoiding the influence of vibration impact generated during the deployment of the solar cell array on the satellite attitude and orbital position. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] Figure 1 A schematic structural diagram of a satellite solar cell array deployment mechanism in a closed state according to an embodiment of the present invention is shown.

[0019] Figure 2 A schematic structural diagram of a satellite solar cell array deployment mechanism in a deployed state according to an embodiment of the present invention is shown.

[0020] Figure 3 A schematic perspective view showing a satellite solar cell array deployment mechanism according to an embodiment of the present invention installed on a satellite is shown. DETAILED DESCRIPTION

[0021] Embodiments of the present invention are described in detail below with reference to the accompanying drawings. In the following description, many specific details are set forth so that those skilled in the art can more fully understand the present invention. However, it will be apparent to those skilled in the art that the present invention may be implemented without some of these specific details. In addition, it should be understood that the present invention is not limited to the specific embodiments described. On the contrary, any combination of the following features and elements may be considered to implement the present invention, regardless of whether they relate to different embodiments. Therefore, the following aspects, features, embodiments, and advantages are for illustrative purposes only and should not be considered as elements or limitations of the claims, unless expressly set forth in the claims.

[0022] The meanings of the terms involved in this specification are generally the usual meanings in the art, or the meanings normally understood by those skilled in the art after reading this specification. The terms "include" and "comprising" in this specification are open-ended, that is, in addition to the elements mentioned, other elements not mentioned may also be included. The terms "connect", "connected" and other similar terms in this specification generally include mechanical connections, electrical connections or a combination thereof, and generally include both direct connections and indirect connections or connections via other components. The terms "first", "second" and the like in this specification are only used to distinguish different components of the same type, and do not indicate any order in terms of importance, structure, function, etc.

[0023] Now refer to Figure 1-3 ,in, Figure 1 FIG2 shows a schematic structural diagram of a satellite solar cell array deployment mechanism in a closed state according to an embodiment of the present invention. Figure 2 A schematic structural diagram showing a satellite solar cell array deployment mechanism in a deployed state according to an embodiment of the present invention is shown. Figure 3 A schematic perspective view showing a satellite solar cell array deployment mechanism according to an embodiment of the present invention installed on a satellite is shown.

[0024] like Figure 1-3 As shown in FIG, the satellite solar cell array deployment mechanism 10 according to an embodiment of the present invention includes:

[0025] A servo motor (not shown) is fixed to the satellite structure frame 9;

[0026] A first fixing rod 4 is fixedly connected to the satellite structure frame 9;

[0027] A first movable rod 1, a first end of which is pivotally connected to a first end of a first fixed rod 4 via a first pin 21;

[0028] A second movable rod 3, a first end of which is pivotally connected to the second end of the first fixed rod 4 via a second pin 22;

[0029] A second fixing rod 6 is fixedly connected to the satellite solar cell array 8;

[0030] A third movable rod 5, a first end of which is pivotally connected to the first end of the second fixed rod 6 via a third pin 23;

[0031] a fourth movable rod 7, a first end of which is pivotally connected to the second end of the second fixed rod 6 via a fourth pin 24;

[0032] The second end of the first movable rod 1 is pivotally connected to the second end of the third movable rod 5 via a fifth pin 25;

[0033] The second end of the second movable rod 3 is pivotally connected to the second end of the fourth movable rod 7 via a sixth pin 26;

[0034] The middle parts of the second movable rod 3 and the third movable rod 5 are pivotally connected via a seventh pin 27;

[0035] The output shaft of the servo motor drives the first movable rod 1 to rotate around the first pin shaft 21 , thereby driving the satellite solar cell array 8 to unfold and fold relative to the satellite structure frame 9 .

[0036] The first fixing rod 4 and the second fixing rod 6 can be fixedly connected to the satellite structure frame 9 and the satellite solar cell array 8 respectively by means of bolts, etc., which can be as follows: Figure 1-2 The rectangular frame shown in the figure is a right-angle connector structure, and its length can be measured from the center points of the first pin 21 and the second pin 22 and the third pin 23 and the fourth pin 24 on both sides thereof.

[0037] Each movable rod can rotate about a corresponding pin. For example, each pin can be fixed to a movable rod, and the pin can rotate in a pin hole provided on another movable rod or fixed rod, so that the two movable rods (or one movable rod and one fixed rod) can rotate about the pin relative to each other. In this way, the two fixed rods and four movable rods form a six-bar mechanism. The first movable rod 1 can be directly or indirectly connected to the output shaft of the servo motor at the first pin 21. For example, the first movable rod 1 can be fixedly connected to or integral with the first pin 21. The first pin 21 can rotate in a pin hole on the first fixed rod 4 and is connected to the output shaft of the servo motor. In this way, the first movable rod 21, driven by the servo motor, drives the third movable rod 5, the second movable rod 3, and the fourth movable rod 7 to move, and drives the solar cell array 8 to expand (or close) through the second fixed rod 6, while ensuring that the solar cell array 8 does not interfere with the satellite structure frame 9. It rotates about each connecting pin from the starting set position to the set end position, thereby driving the entire mechanism to achieve the solar cell array to expand to the set angle at a constant angular velocity. When the servo motor stops running, the solar array 8 remains in the unfolded state, providing a continuous supply of energy to the entire satellite system.

[0038] In some embodiments, the first movable rod 1 and the fourth movable rod 7 are straight rods, the second movable rod 3 and the third movable rod 5 are bent rods, and the length relationship between the rods satisfies the following conditions:

[0039] The sum of the length of the first fixed rod 4 and the length of the first movable rod 1 is less than or equal to the sum of the length of the second movable rod 3 from the second pin 22 to the seventh pin 27 and the length of the third movable rod 5 from the seventh pin 27 to the fifth pin 25; and

[0040] The sum of the length of the second fixed rod 6 and the length of the fourth movable rod 7 is less than or equal to the sum of the length of the third movable rod 5 from the third pin 23 to the seventh pin 27 and the length of the second movable rod 3 from the seventh pin 27 to the sixth pin 26.

[0041] The length of each rod or segment can be calculated from the center point of the pin at both ends of each rod or segment, and the length of the third movable rod 5 from the third pin 23 to the seventh pin 27 refers to the sum of the lengths of the two broken lines it contains.

[0042] In some embodiments, the second movable rod 3 is bent once at the seventh pin 27 toward the satellite solar array 8, and the third movable rod 5 is bent twice near the third pin 23 and at the seventh pin 27 toward the satellite structure frame 9, and the angle of each bend enables the satellite solar array 8 to be unfolded from a closed state at a 90° angle to the satellite structure frame 9 to an unfolded state at a 180° angle to the satellite structure frame 9. Specifically, Figure 1-2 As shown in the figure, the angles of each bend and the lengths of each segment ensure that during the deployment of the satellite solar cell array 8, the angle formed by the first movable rod 1 and the third movable rod 5 at the fifth pin 25, and the angle formed by the second movable rod 3 and the third movable rod 5 at the seventh pin 27 are always less than 90°.

[0043] By appropriately bending the second movable rod 3 and the third movable rod 5, the above-mentioned rod length condition can be easily met, and the satellite solar cell array 8 can be easily unfolded from a closed state at an angle of 90° to the satellite structural frame 9 to an unfolded state at an angle of 180° to the satellite structural frame 9.

[0044] While the satellite solar array deployment mechanism 10 according to an embodiment of the present invention has been described above with reference to the accompanying drawings, it should be noted that the above description and illustrations are merely illustrative and not limiting of the present invention. In other embodiments of the present invention, the device may include more, fewer, or different components, and the connections, inclusions, and functional relationships between the components may differ from those described and illustrated. For example, a component may typically include other subcomponents in addition to those illustrated and described; multiple components may be combined into a larger component; and so on. All such variations are within the spirit and scope of the present invention.

[0045] The satellite solar cell array deployment mechanism according to the embodiment of the present invention has a simple structure, is stable and reliable, has a controllable deployment angle, has a small deployment impact, and is easy to install. It can ensure that the satellite solar cell array is deployed to a set angle at a constant angular velocity and maintained at the set position. The gap between the satellite solar cell array and the structural frame can be very small (under the same conditions, the solar cell array can have a larger area), effectively avoiding interference with the satellite structural frame during the deployment process, resulting in a large portion of the solar cell array being cut off. At the same time, it also avoids the influence of vibration impact generated during the deployment of the solar cell array on the satellite attitude and orbital position.

[0046] Although the present invention has been disclosed above through the embodiments, the present invention is not limited thereto. Various changes and modifications made by those skilled in the art without departing from the spirit and scope of the present invention should be included in the scope of protection of the present invention, and the scope of protection of the present invention shall be subject only to the scope defined by the language of the claims and their equivalents.

Claims

1. A satellite solar cell array deployment mechanism, comprising: A servo motor is fixed on the satellite structure frame (9); A first fixing rod (4) is fixedly connected to the satellite structure frame (9); A first movable rod (1), a first end of which is pivotally connected to a first end of a first fixed rod (4) via a first pin (21); A second movable rod (3), a first end of which is pivotally connected to the second end of the first fixed rod (4) via a second pin (22); A second fixing rod (6) is fixedly connected to the satellite solar cell array (8); A third movable rod (5), a first end of which is pivotally connected to the first end of the second fixed rod (6) via a third pin (23); a fourth movable rod (7), a first end of which is pivotally connected to the second end of the second fixed rod (6) via a fourth pin (24); The second end of the first movable rod (1) and the second end of the third movable rod (5) are pivotally connected via a fifth pin (25); The second end of the second movable rod (3) is pivotally connected to the second end of the fourth movable rod (7) via a sixth pin (26); The middle parts of the second movable rod (3) and the third movable rod (5) are pivotally connected via a seventh pin (27); The output shaft of the servo motor drives the first movable rod (1) to rotate around the first pin shaft (21), thereby driving the satellite solar cell array (8) to unfold and fold relative to the satellite structural frame (9).

2. The satellite solar cell array deployment mechanism according to claim 1, wherein: The first movable rod (1) and the fourth movable rod (7) are straight rods, the second movable rod (3) and the third movable rod (5) are bent rods, and the length relationship between the rods satisfies the following conditions: The sum of the length of the first fixed rod (4) and the length of the first movable rod (1) is less than or equal to the sum of the length of the second movable rod (3) from the second pin shaft (22) to the seventh pin shaft (27) and the length of the third movable rod (5) from the seventh pin shaft (27) to the fifth pin shaft (25); as well as The sum of the length of the second fixed rod (6) and the length of the fourth movable rod (7) is less than or equal to the sum of the length of the third movable rod (5) from the third pin shaft (23) to the seventh pin shaft (27) and the length of the second movable rod (3) from the seventh pin shaft (27) to the sixth pin shaft (26).

3. The satellite solar cell array deployment mechanism according to claim 2, wherein: The second movable rod (3) is bent once at the seventh pin (27) toward the satellite solar cell array (8), and the third movable rod (5) is bent twice near the third pin (23) and at the seventh pin (27) toward the satellite structural frame (9), and the angles of each bend enable the satellite solar cell array (8) to be unfolded from a closed state at an angle of 90° to the satellite structural frame (9) to an unfolded state at an angle of 180° to the satellite structural frame (9).

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