Solar wing capable of being repeatedly unfolded according to different preset angles and design method

Abstract

The invention discloses a solar wing capable of being repeatedly unfolded according to different preset angles and a design method. The solar wing comprises a fixed wing, a left unfolding wing, a right unfolding wing and a pressing mechanism. The fixed wing is located in the middle and is in a horizontal state, the left unfolding wing and the right unfolding wing are fixed to the two sides of the fixed wing through a pressing mechanism and pressed on the outer wall of a spacecraft, the left unfolding wing and the right unfolding wing conduct releasing and unfolding actions in the working process, and the left unfolding wing and the right unfolding wing keep a fixed included angle beta with the fixed wing in the pressed state; during working, two-section unfolding is performed according to a preset angle 1 and a preset angle 2. According to the solar wing capable of being repeatedly unfolded according to the different preset angles and the design method, the advantages of high reliability and low cost of a fixed solar wing and high power generation efficiency of a solar wing with a driving mechanism are combined, the power generation efficiency of the solar wing is reliably improved with low cost, and the solar wing adapts to different illumination conditions; the method is particularly suitable for emergency launching of spacecrafts providing services for high-value space resources such as space stations.

Description

Technical Field

[0001] The present invention belongs to the technical field of spacecraft solar wings, and in particular relates to a solar wing that can be deployed multiple times at different predetermined angles and a design method thereof. Background Art

[0002] Most spacecraft in orbit rely mainly on solar panels to obtain energy. Due to the limitations of the launch window and the influence of the spacecraft's orbit, the power generation efficiency of fixed solar panels is not high. A solar-directional drive mechanism is needed to further improve the power generation efficiency. However, the solar-directional drive mechanism is expensive, and the rotating components reduce the reliability of the spacecraft. Therefore, it is of great significance to design a simple and reliable solar panel deployment method that can achieve high power generation efficiency under different lighting conditions. Summary of the Invention

[0003] The technical problem solved by the present invention is: in response to the shortcomings of existing fixed solar wing technology, a solar wing that can be deployed multiple times at different predetermined angles and a design method are provided. The advantages of high reliability and low cost of fixed solar wings are combined with the high power generation efficiency of solar wings with drive mechanisms. The power generation efficiency of solar wings is reliably improved at a lower cost, and the system can adapt to different lighting conditions. It is particularly suitable for emergency launches of spacecraft that provide services for high-value space resources such as space stations.

[0004] In order to solve the above technical problems, the present invention discloses a solar wing that can be unfolded multiple times at different predetermined angles, including a fixed wing, a left unfolding wing, a right unfolding wing, and a clamping mechanism; the fixed wing is pressed and fixed on the outer wall of a spacecraft in a horizontal state; the left unfolding wing and the right unfolding wing are both pressed on the outer wall of the spacecraft by the clamping mechanism, and are respectively located on the left and right sides of the fixed wing; the left unfolding wing and the right unfolding wing perform unfolding and releasing actions when working; the left unfolding wing and the right unfolding wing maintain a fixed angle β with the fixed wing in the compressed state, and perform two-stage unfolding according to a first predetermined unfolding angle and a second predetermined unfolding angle when working.

[0005] Furthermore, the first predetermined deployment angle and the second predetermined deployment angle are determined according to the variation characteristics of the solar illumination angle α of the orbital plane on which the spacecraft is located during its on-orbit operation.

[0006] Furthermore, the solar illumination angle α varies in the range of -θ to +θ, and the first predetermined deployment angle is β+θ / 2.

[0007] Furthermore, the second predetermined deployment angle is an average change value θ / 2 of the sunlight angle α, and the second deployment direction is opposite to the first deployment direction.

[0008] Further, after the left deployment wing and the right deployment wing are deployed and released, they have three states, namely the first deployment state, the second deployment state, and the third deployment state; the first deployment state is that the left deployment wing is deployed once to a first predetermined deployment angle and the right deployment wing is not deployed; the second deployment state is that the left deployment wing is not deployed and the right deployment wing is deployed once to a first predetermined deployment angle; the third deployment state is that the left deployment wing is deployed twice to a second predetermined deployment angle and the right deployment wing is deployed twice to a second predetermined deployment angle.

[0009] Further, the three states are determined according to the characteristics of the solar illumination angle α of the orbital plane during the on-orbit operation of the spacecraft; when the solar illumination angle α is in the range of (θ / 2, θ], it is the first deployment state; when the solar illumination angle α is in the range of [-θ, -θ / 2), it is the second deployment state; when the solar illumination angle α is in the range of [-θ / 2, θ / 2], it is the third deployment state.

[0010] The present invention also provides a design method for a solar wing that is deployed multiple times at different predetermined angles, and the specific steps are as follows:

[0011] S1. According to the spacecraft orbit design, determine the variation range of the solar illumination angle α;

[0012] S2. According to the spacecraft configuration design, determine the included angle β between the compressed state of the deployment wing and the fixed wing;

[0013] S3. Conduct the design of the deployment wing to determine the first predetermined deployment angle and the second predetermined deployment angle;

[0014] S4. According to the design situation of the launch window, determine the deployment strategy of the deployment wing based on the solar altitude angle at the time of launch.

[0015] Further, the deployment strategy of the deployment wing is as follows: when the solar illumination angle α at the time of launch is in the range of (θ / 2, θ], the left deployment wing is deployed once according to the first predetermined deployment angle and the right deployment wing is not deployed, being in the first deployment state. After entering the orbit, according to the change of the solar illumination angle α, switch the deployment state to the third deployment state according to the deployment state determination method; when the solar illumination angle α at the time of launch is in the range of [-θ, -θ / 2), the right deployment wing is deployed once according to the first predetermined deployment angle and the left deployment wing is not deployed, being in the second deployment state. After entering the orbit, according to the change of the solar illumination angle α, switch the deployment state to the third deployment state according to the deployment state determination method; when the solar illumination angle α at the time of launch is in the range of [-θ / 2, θ / 2], the left deployment wing is deployed twice to the horizontal state and the right deployment wing is deployed twice to the horizontal state, remaining in the third deployment state.

[0016] The present invention has the following advantages:

[0017] (1) A design method of a solar wing that unfolds multiple times at different predetermined angles. Three solar wings, namely a fixed wing, a left unfolding wing, and a right unfolding wing, are designed. In the launch state, they are all pressed tightly against the outer wall of the spacecraft, with a compact structure, making full use of the space in the fairing of the launch vehicle.

[0018] (2) A design method of a solar wing that unfolds multiple times at different predetermined angles. The primary and secondary directional unfolding of the unfolding wing is achieved by the grouped release of the pressing mechanism, without the need to configure a sun-pointing drive mechanism, with a simple structure and high reliability.

[0019] (3) A design method of a solar wing that unfolds multiple times at different predetermined angles. Under different lighting conditions, the spacecraft can obtain a high power generation efficiency. Description of the Drawings

[0020] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings. It should be noted that for clarity and ease of illustration, the drawings are not necessarily drawn to scale.

[0021] Figure 1 Schematic diagram of the pressing state of the solar wing that unfolds multiple times at different predetermined angles of the present invention;

[0022] Figure 2 Front view of the unfolded state 1 of the solar wing that unfolds multiple times at different predetermined angles of the present invention;

[0023] Figure 3 Schematic diagram of the unfolded state 2 of the solar wing that unfolds multiple times at different predetermined angles of the present invention;

[0024] Figure 4 Schematic diagram of the unfolded state 3 of the solar wing that unfolds multiple times at different predetermined angles of the present invention;

[0025] Figure 5 Schematic diagram of the design flow of the design method of the solar wing that unfolds multiple times at different predetermined angles of the present invention.

[0026] Reference signs in the figures: 1 - left unfolding wing, 2 - fixed wing, 3 - right unfolding wing, 4 - included angle β between the pressing state of the unfolding wing and the fixed wing, 5 - outer wall of the spacecraft, 11 - predetermined unfolding angle 1, 12 - predetermined unfolding angle 2, 13 - solar illumination angle, 14 - direction of solar rays.

[0027] Detailed Embodiments

[0028] The present invention will be further described below in conjunction with the drawings.

[0029] In this embodiment, Figure 1 It is a schematic diagram of the compressed state of the solar wing that unfolds multiple times at different predetermined angles. The solar wing includes a left deployment wing 1, a fixed wing 2, and a right deployment wing 3;

[0030] The fixed wing 2 is tightly pressed and fixed on the outer wall 5 of the spacecraft in a horizontal state; the left deployment wing 1 is pressed on the outer wall 5 of the spacecraft by a pressing mechanism and is located on the left side of the fixed wing 2, the right deployment wing 3 is pressed on the outer wall 5 of the spacecraft by a pressing mechanism and is located on the right side of the fixed wing 2, and the left deployment wing 1 and the right deployment wing 3 perform unfolding and releasing actions during operation; the left deployment wing 1, the fixed wing 2, and the right deployment wing 3 all adapt to the space environment and provide energy for the spacecraft;

[0031] In the compressed state, both the left deployment wing 1 and the right deployment wing 3 maintain a fixed angle β4 with the fixed wing 2, and perform two-stage unfolding according to the first predetermined unfolding angle 11 and the second predetermined unfolding angle 12 during operation.

[0032] The working condition of this embodiment is: a cargo transportation spacecraft transports goods to the space station, with an on-orbit lifespan of 3 months. The target orbit is the operating orbit of the space station at an altitude of 400 km and an inclination of 42°. The range of the solar illumination angle α13 in this orbital plane is [-66°, 66°]. In the entire mission profile, the flight stage with the greatest energy demand is the rendezvous and docking section with the space station after entering the orbit. Taking the solar illumination angle of 40° in the orbital plane after entering the orbit and the solar illumination angle of 20° in the orbital plane after departing from the space station as examples, a solar wing that unfolds twice at a predetermined angle in orbit is designed, and the angle β4 between the compressed state of the deployment wing and the fixed wing 2 is 55°.

[0033] Figure 5 It is a schematic diagram of the process flow of the design method of the solar wing that unfolds multiple times at different predetermined angles. The following presents a design method of the solar wing that unfolds multiple times at different predetermined angles, including: First, according to the spacecraft orbit design, determine the range of change of the solar illumination angle α; then, according to the spacecraft configuration design, determine the angle β between the compressed state of the deployment wing and the fixed wing 2; then, conduct the design of the deployment wing to determine the first predetermined unfolding angle and the second predetermined unfolding angle; then, according to the design situation of the launch window, determine the deployment strategy of the deployment wing based on the solar altitude angle at the time of launch; finally, according to the change of the illumination angle in orbit, the deployment wing unfolds further to improve the power generation efficiency of the entire solar wing.

[0034] Figure 2 It is a schematic diagram of the first deployment state of the solar wing that unfolds multiple times at different predetermined angles, Figure 3 It is a schematic diagram of the second deployment state of the solar wing that unfolds multiple times at different predetermined angles, Figure 4Schematic diagram of the third deployment state of the solar wing that unfolds multiple times at different predetermined angles. The first deployment state is that the left deployment wing unfolds once to the first predetermined deployment angle, and the right deployment wing does not unfold; the second deployment state is that the left deployment wing does not unfold, and the right deployment wing unfolds once to the first predetermined deployment angle; the third deployment state is that the left deployment wing unfolds twice to the horizontal state, and the right deployment wing unfolds twice to the horizontal state.

[0035] The two predetermined deployment angles, the first predetermined deployment angle 11 and the second predetermined deployment angle 12, are determined according to the variation characteristics of the solar illumination angle α of the orbital plane during the on-orbit operation of the spacecraft; in this embodiment, the included angle β between the deployment wing in the compressed state and the fixed wing 2 is 55°, the first predetermined deployment angle 11 is β + θ / 2 = 88°, the second predetermined deployment angle 12 is the average change value θ / 2 = 33° of the solar illumination angle α13, and the two deployment directions are opposite.

[0036] The different deployment states of the deployment wing are determined according to the characteristics of the solar illumination angle α of the orbital plane during the on-orbit operation of the spacecraft. When the solar illumination angle α is within the range of (θ / 2, θ], it is the first deployment state; when the solar illumination angle α is within the range of [-θ, -θ / 2), it is the second deployment state; when the solar illumination angle α is within the range of [-θ / 2, θ / 2], it is the third deployment state.

[0037] In this embodiment, after entering the orbit, the solar illumination angle α of the orbital plane is 40°, the deployment wing is in the first deployment state, the left deployment wing 1 unfolds 88° at once, and the right deployment wing 3 does not unfold; after evacuating from the space station, the solar illumination angle α of the orbital plane is 20°, the deployment wing is in the third deployment state, the left deployment wing 1 unfolds twice to the horizontal state, and the right deployment wing 3 unfolds twice to the horizontal state.

[0038] According to the solar illumination angle α of the orbital plane after entering the orbit being 40°, if a fixed solar wing is adopted and the left solar wing does not unfold after entering the orbit, only the fixed wing and the right solar wing have the power generation ability; if a solar wing that unfolds multiple times at different predetermined angles is adopted, as described in the embodiment, the left deployment wing 1 unfolds 88° at once, and the overall power generation power is increased by 57% without increasing the area of the solar wing;

[0039] According to the solar illumination angle α of the orbital plane after evacuating from the space station being 20°, if a fixed solar wing is adopted and the left solar wing does not unfold, only the fixed wing and the right solar wing have the power generation ability; if the design method of a solar wing that unfolds multiple times at different predetermined angles is adopted, as described in the embodiment, the left deployment wing 1 further unfolds twice to the horizontal state, and the right deployment wing 3 unfolds twice to the horizontal state, and the overall power generation power is increased by 60% without increasing the area of the solar wing.

[0040] As described in this embodiment, under the condition of a 40° light incidence angle, the power generation efficiency is increased by 57% in the first unfolded state; under the condition of a 20° light incidence angle, the power generation efficiency is increased by 60% in the third unfolded state.

[0041] Although the present invention has been disclosed above with the preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present invention without departing from the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention. The content not detailedly described in the specification of the present invention belongs to the well-known technology of those skilled in the art.

Claims

1. A solar panel that unfolds multiple times at different predetermined angles, characterized in that: The solar panel includes a fixed wing, a left deployable wing, a right deployable wing, and a pressing mechanism; The fixed wing is tightly pressed and fixed on the outer wall of the spacecraft in a horizontal state; When launched, the left deployable wing and the right deployable wing are both pressed on the outer wall of the spacecraft by the pressing mechanism, and are respectively located on the left and right sides of the fixed wing; when working, the left deployable wing and the right deployable wing perform deployment and release actions; When in the pressed state, the left deployable wing and the right deployable wing maintain a fixed angle β with the fixed wing, and perform two-stage deployment according to the first predetermined deployment angle and the second predetermined deployment angle when working.

2. A solar wing that unfolds multiple times at different predetermined angles as described in claim 1, characterized in that: The first predetermined deployment angle and the second predetermined deployment angle are determined according to the variation characteristics of the solar illumination angle α of the orbital plane during the on-orbit operation of the spacecraft.

3. A solar wing that unfolds multiple times at different predetermined angles as described in claim 2, characterized in that: The variation range of the solar illumination angle α is -θ to +θ, and the first predetermined deployment angle is β + θ / 2.

4. A solar wing that unfolds multiple times at different predetermined angles, characterized in that: The second predetermined deployment angle is the average variation value θ / 2 of the solar illumination angle α, and the second deployment direction is opposite to the first deployment direction.

5. A solar wing that unfolds multiple times at different predetermined angles, characterized in that: After the left deployable wing and the right deployable wing are deployed and released, there are three states, namely the first deployment state, the second deployment state, and the third deployment state; The first deployment state is that the left deployable wing is deployed once to the first predetermined deployment angle and the right deployable wing is not deployed; The second deployment state is that the left deployable wing is not deployed and the right deployable wing is deployed once to the first predetermined deployment angle; The third deployment state is that the left deployable wing is deployed twice to the horizontal state and the right deployable wing is deployed twice to the horizontal state.

6. A solar wing that unfolds multiple times at different predetermined angles as described in claim 5, characterized in that: The three states are determined according to the characteristics of the solar illumination angle α of the orbital plane during the on-orbit operation of the spacecraft; When the solar illumination angle α is in the range of (θ / 2, θ], it is the first deployment state; When the solar illumination angle α is in the range of [-θ, -θ / 2), it is the second deployment state; When the solar illumination angle α is in the range of [-θ / 2, θ / 2], it is the third deployment state.

7. A design method for a solar wing that unfolds multiple times at different predetermined angles, characterized in that, The specific steps of the method are as follows: S1. According to the spacecraft orbit design, determine the variation range of the solar illumination angle α; S2. According to the spacecraft configuration design, determine the angle β between the pressed state of the deployable wing and the fixed wing; S3. Conduct the design of the deployable wing to determine the first predetermined deployment angle and the second predetermined deployment angle; S4. According to the launch window design situation, determine the deployment strategy of the deployable wing based on the solar altitude angle at the time of launch.

8. A design method for a solar array that unfolds multiple times at different predetermined angles, characterized in that, The deployment strategy of the deployable wing is as follows: At the time of launch, when the solar illumination angle α is in the range of (θ / 2, θ], the left deployable wing is deployed once according to the first predetermined deployment angle and the right deployable wing is not deployed, being in the first deployment state; After entering the orbit, according to the variation of the solar illumination angle α, switch the deployment state to the third deployment state according to the deployment state determination method described in claim 6; At the time of launch, when the solar illumination angle α is in the range of [-θ, -θ / 2), the right deployable wing is deployed once according to the first predetermined deployment angle and the left deployable wing is not deployed, being in the second deployment state; After entering the orbit, according to the variation of the solar illumination angle α, switch the deployment state to the third deployment state according to the deployment state determination method described in claim 6; During launch, the solar illumination angle α is within the range of [-θ / 2, θ / 2], the left deployment wing is secondarily deployed to the horizontal state, and the right deployment wing is secondarily deployed to the horizontal state, maintaining the third deployment state.

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