Experiment module solar wing double-shaft coupling sun-alignment fixed angle servo control method and system

Abstract

The invention discloses an experiment module solar wing double-shaft coupling sun-facing fixed angle servo control method and system, and aims to reduce mechanical wear by reducing rotation of a solar wing, reduce collision risk by reducing the projection area of the solar wing on an orbital plane orthogonal plane, reduce physical damage on the basis of meeting the power requirement, and improve the reliability of the solar wing double-shaft coupling sun-facing fixed angle servo control system. The task requirement for long-life operation of the space station is met, and the on-orbit life of the solar wing of the space station is prolonged. Namely, sun orientation control is completed by adopting a fixed rotation angle value and an intermittent rotation mode for the two degrees of freedom of the solar wing. The design process comprises the steps that firstly, a beta axis is designed to be in a fixed angle servo control mode, the rotation fixed angle of the beta axis is totally 2n + 1 gears, and the value range is {-n delta,...,-3 delta,-2 delta,-delta, 0, delta, 2 delta, 3 delta,..., n delta}; when the beta axis does not return to zero, setting the alpha axis to return to zero horizontally; when the beta axis returns to zero, if the alpha axis returns to zero horizontally and cannot meet the power requirement, the alpha axis is designed to be in a fixed angle servo control mode, the number of fixed rotation angles of the alpha axis is three, and the value range is {-angle alpha0 ', 0 and angle alpha0'}.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft on-orbit management technology, and in particular to a method and system for servo control of a dual-axis coupled fixed-angle sun-facing solar wing in an experimental cabin. Background Art

[0002] Xie Pengru (Xie Pengru, Gu Shaojing, Ma Jijun, et al. Configuration and layout design of split-and-assembled dual-degree-of-freedom solar orientation system [J]. Shanghai Aerospace (Chinese and English), 2023, 40(05): 94-101.), Zhao Zhen (Zhao Zhen, Wang Bi, Qian Zhiyuan, et al. Solar tracking control of flexible solar cell wing of space station [J]. Shanghai Aerospace (Chinese and English), 2021, 38(05): 31-40.) and others pointed out that the dual-degree-of-freedom solar orientation system of the space station experimental module adopts a split-and-assembled solar orientation configuration scheme, and realizes the dual-degree-of-freedom solar orientation of the solar cell wing through two axially orthogonal single-degree-of-freedom solar orientation subsystems, such as Figure 1 During the space station's orbital operation, the solar panels rotate around the α axis to compensate for the impact of orbital angle changes on the solar incidence angle, and the solar panels rotate around the β axis to compensate for the impact of solar altitude angle changes on the solar incidence angle.

[0003] Prior to this, the solar panels of large domestic spacecraft such as the Tiangong space laboratory, Shenzhou spacecraft and cargo spacecraft all adopted a single-degree-of-freedom continuous solar orientation method. The only large low-orbit spacecraft abroad that uses dual-degree-of-freedom solar orientation technology is the International Space Station in the United States.

[0004] When the experimental module is flying alone, the β-axis solar orientation subsystem realizes continuous rotation of 0° to 360° within the orbital period in the three-axis earth-oriented flight attitude. In the inertial flight attitude, the β-axis solar orientation subsystem rotates intermittently to compensate for the change of the solar altitude angle.

[0005] After the experimental module docked and transferred with the core module in orbit, a three-module configuration of the space station "T" combination was formed. In this configuration, the α solar orientation device was powered on and started, accelerating from a stop state to 0.300° / s, and began to search and capture the sun. After capturing the sun, the α solar orientation device switched to the space station orbital angular velocity of 0.065° / s to continuously track the sun, achieving continuous rotation of 0° to 360° within the orbital cycle; the β axis adopted a fixed angle servo mode for the sun, and its gear adjustment threshold was 14°, which could ensure that the incident angle of the solar cell wing was within 0° to 14°, and the power generation of the solar cell wing was in a relatively stable maximum output state.

[0006] Yang Yongmiao et al. pointed out (Yang Yongmiao, Ge Maoyan, Zhou Yan, et al. Modeling and analysis of multi-attitude output capability of manned spacecraft solar panels [J]. Shanghai Aerospace (Chinese and English), 2023, 40(06): 30-39) that, according to the needs of the actual on-orbit mission, the solar panels may be in a complex working state of stopping at a specific position of the α-axis in addition to the rotating sun-oriented working state. In working modes such as "radial rendezvous, docking and evacuation of the assembly" and "long-term low-load operation of the power channel", the solar panels will be in a working state of horizontal zeroing of the α-axis, and in working modes such as "axial rendezvous, docking and evacuation of the assembly" and "assembly orbit control", the solar panels will be in a working state of vertical zeroing of the α-axis.

[0007] At the same time, when the α-axis of the solar wing is horizontally zeroed and the β-axis adopts a fixed-angle servo mode facing the sun, as the solar altitude angle increases, the output capacity of the sailboard increases accordingly, and as the solar altitude angle decreases, the output capacity of the sailboard decreases accordingly.

[0008] According to this principle, when the β-axis angle is large, the α-axis can meet the power demand by returning to zero horizontally. When the β-axis angle is small, if the α-axis cannot meet the power demand by returning to zero horizontally, a sun-directed control strategy for the α-axis needs to be designed.

[0009] Daniel Hastings et al. clearly stated in their book (Spacecraft-Space Environment Interaction) that the relative impact velocity (km / s) at low orbital altitude can be estimated from the angle θ between the impact velocity vector and the spacecraft velocity vector: V_ d =15.4cos(±θ). That is, within the range of 0° to 90°, the larger the angle θ, the smaller the impact speed. Therefore, the space debris flying head-on in the forward direction of the space station's orbit will cause the greatest collision damage to the solar panels. The cumulative projection area of ​​the space station's solar panels on the plane orthogonal to the orbital plane should be minimized to reduce the risk of being hit by space debris in the low-orbit space environment.

[0010] The existing Japanese-targeted strategies have the following disadvantages:

[0011] The continuous directional tracking mode of the solar panels of the space station will firstly cause continuous mechanical wear on the rotation mechanism of the solar panels; secondly, the cumulative projection area of ​​the solar panels on the plane orthogonal to the orbital plane is relatively large, and the collision risk in the low-orbit space environment is relatively high. Both mechanical wear and debris impact will reduce the power generation capacity of the solar panels, directly affecting the operating life of the space station. Summary of the invention

[0012] In view of this, the present invention provides a dual-axis coupled fixed-angle servo control method and system for the experimental cabin solar wing. The design purpose is to reduce mechanical wear by reducing the rotation of the solar wing, and to reduce the risk of collision by reducing the projection area of ​​the solar wing on the orthogonal plane of the orbital plane. While meeting the power requirements, physical damage is reduced to meet the mission requirements of long-life operation of the space station and extend the on-orbit life of the space station solar wing.

[0013] To achieve the above-mentioned purpose, the technical solution of the present invention - a method for controlling a fixed angle of servo of a dual-axis coupled solar wing in an experimental cabin, comprises the following steps:

[0014] S1: Clarify the solar altitude angle range of the space station.

[0015] S2: Preliminarily determine the gear difference Δ and angle adjustment threshold δ of the β-axis fixed-angle servo to the sun; set the β-axis of the solar wing in the experimental cabin to the fixed-angle servo control mode.

[0016] S3: Get the maximum power generation W of the solar wing 总 The numerical value of .

[0017] S4: Calculate the current power requirement W of the space station.

[0018] S5: Based on the space station power demand W and the maximum power generation of the solar wing W 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun.

[0019] S6: Determine whether ∠α0' is less than ε. If ∠α0'>ε, the α-axis of the solar wing is set to a continuous sun-orientation control mode; if ∠α0'<ε, the α-axis of the solar wing is set to a fixed-angle servo control mode.

[0020] S7: When ∠β' is just adjusted to Δ or -Δ, the power generation capacity of the solar wing is calculated when the α axis is horizontally zeroed to evaluate whether it meets the mission requirements. If it meets the mission requirements, the current values ​​of Δ and δ are adopted and the process ends.

[0021] If the task requirements cannot be met, execute S8.

[0022] S8: Increase the gear difference Δ of the β-axis angle value and the value of the adjustment threshold δ, so as to postpone the timing of the β-axis gear adjustment, delay the time for the α-axis to return to zero horizontally, increase the solar altitude angle to ensure the power output of the solar wing, and return to execute step S7 until the solar wing's power generation capacity meets the mission requirements when the α-axis returns to zero horizontally.

[0023] Further, S2: preliminarily determining the gear difference Δ and angle adjustment threshold δ of the β-axis fixed angle servo to the sun, and setting the β-axis of the solar wing of the experimental cabin to a fixed angle servo control mode, specifically including the following steps:

[0024] S21: Determine the value of Δ, where the fixed angle value range of the β-axis rotation is {-nΔ,…,-3Δ,-2Δ,-Δ,0,Δ,2Δ,3Δ,…,nΔ}. Since the β-axis is used to compensate for the influence of the solar altitude angle on the solar incident angle, and the solar altitude angle range of the space station is -66° to 66°, nΔ≤66°.

[0025] S22: Determine the value of δ, δ<Δ.

[0026] S23: The servo control strategy for the fixed angle of the β axis of the solar wing of the experimental cabin is as follows: when the solar altitude angle is within the range of (-δ, δ), the β axis is rotated so that ∠β'=0°; as the solar altitude angle increases, when ∠β=δ, the β axis is rotated so that ∠β'=Δ; when ∠β=Δ+δ, the β axis is rotated so that ∠β'=2Δ; and so on, until ∠β'=nΔ. As the solar altitude angle decreases, when ∠β=-δ, the β axis is rotated so that ∠β'=-Δ; when ∠β=-Δ-δ, the β axis is rotated so that ∠β'=-2Δ; and so on, until ∠β'=-nΔ.

[0027] Further, S4: calculate the current power demand W of the space station, specifically: the power generated by the space station's solar panels is used to power its own platform, charge the battery pack, power the application payload, and power the docked aircraft. After calculating each power demand, the total is calculated and a margin is left to obtain W.

[0028] Further, S5: according to the power demand W of the space station and the maximum power generation power W of the solar wing 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun, specifically:

[0029] When ∠α'=0 and ∠β'=0, if the solar wing power generation meets the power demand of the space station, the α-axis will be kept in the horizontal zero state; in this mission state, the solar wing sailboard always adopts the solar orientation control strategy of "α-axis horizontal zeroing, β-axis fixed angle servo mode to the sun".

[0030] If ∠α'=0, ∠β'=0, W0<W, and the α-axis is defined as three fixed angles: ∠α0', 0°, -∠α0'; ∠α0' is:

[0031]

[0032] Furthermore, if ∠α0'<ε, the α-axis of the solar wing is set to a fixed angle servo control mode, and the following steps are performed:

[0033] When ∠β'≠0, the α-axis keeps the horizontal zero state. When ∠β'=0, the α-axis fixed angle servo to the sun is controlled, and its fixed rotation angle ∠α0' is calculated. The control strategy of the α-axis fixed angle servo to the sun is as follows: when the space station enters the sunlit area, the α-axis is rotated so that ∠α'=∠α0'; when the solar incident angle ∠θ=90°-2∠α0', the α-axis is rotated so that ∠α'=0°; when the solar incident angle ∠θ=-(90°-∠α0'), the α-axis is rotated so that ∠α'=-∠α0'; after entering the shadow area, the α-axis is controlled to return to the horizontal zero state; the α-axis sun orientation device rotates at a speed of 0.300° / s, and the rotation direction is based on the principle of proximity.

[0034] Another embodiment of the present invention further provides a dual-axis coupled sun-facing fixed angle servo control system for an experimental cabin solar wing, comprising the following eight modules, namely:

[0035] The solar altitude angle range determination module is used to determine the solar altitude angle range of the space station.

[0036] The β-axis preliminary setting module is used to preliminarily determine the gear difference Δ and angle adjustment threshold δ of the β-axis fixed-angle servo to the sun; and set the β-axis of the solar wing of the experimental cabin to the fixed-angle servo control mode.

[0037] Power generation acquisition module, used to obtain the maximum power generation W of the solar wing 总 The numerical value of .

[0038] The power demand calculation module of the space station is used to calculate the power demand W of the current space station.

[0039] The α-axis fixed angle calculation module is used to calculate the power demand W of the space station and the maximum power generation power W of the solar wing. 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun.

[0040] The α-axis control mode setting module is used to determine whether ∠α0' is less than ε. If ∠α0'>ε, the solar wing α-axis is set to a continuous sun-orientation control mode; if ∠α0'<ε, the solar wing α-axis is set to a fixed-angle servo control mode.

[0041] The evaluation module is used to calculate the power generation capacity of the solar wing when the α-axis is horizontally zeroed when ∠β' is just adjusted to Δ or -Δ, and evaluate whether it meets the mission requirements. If it meets the mission requirements, the current values ​​of Δ and δ are adopted and the process ends; if it does not meet the mission requirements, the β-axis adjustment module is executed.

[0042] The β-axis adjustment module is used to increase the gear difference Δ of the β-axis angle value and the value of the adjustment threshold δ to postpone the timing of the β-axis gear adjustment, delay the time for the α-axis to return to zero horizontally, increase the solar altitude angle to ensure the power output of the solar wing, and return to execution module seven until the solar wing's power generation capacity meets the mission requirements when the α-axis returns to zero horizontally.

[0043] Furthermore, the β-axis preliminary setting module specifically performs the following steps:

[0044] S21: Determine the value of Δ, where the fixed angle value range of the β-axis rotation is {-nΔ,…,-3Δ,-2Δ,-Δ,0,Δ,2Δ,3Δ,…,nΔ}. Since the β-axis is used to compensate for the influence of the solar altitude angle on the solar incident angle, and the solar altitude angle range of the space station is -66° to 66°, nΔ≤66°.

[0045] S22: Determine the value of δ, δ<Δ.

[0046] S23: The servo control strategy for the fixed angle of the β axis of the solar wing of the experimental cabin is as follows: when the solar altitude angle is within the range of (-δ, δ), the β axis is rotated so that ∠β'=0°; as the solar altitude angle increases, when ∠β=δ, the β axis is rotated so that ∠β'=Δ; when ∠β=Δ+δ, the β axis is rotated so that ∠β'=2Δ; and so on, until ∠β'=nΔ. As the solar altitude angle decreases, when ∠β=-δ, the β axis is rotated so that ∠β'=-Δ; when ∠β=-Δ-δ, the β axis is rotated so that ∠β'=-2Δ; and so on, until ∠β'=-nΔ.

[0047] Furthermore, in the power demand calculation module of the space station, the current power demand W of the space station is calculated. Specifically, the power generated by the solar panels of the space station is used to power its own platform, charge the battery pack, consume electricity for application loads, and power the docked aircraft and connect to the grid. After calculating each power demand, the total is found and a margin is left to obtain W.

[0048] Furthermore, the α-axis fixed angle calculation module specifically performs the following process:

[0049] When ∠α'=0 and ∠β'=0, if the solar wing power generation meets the power demand of the space station, the α-axis will be kept in the horizontal zero state; in this mission state, the solar wing sailboard always adopts the solar orientation control strategy of "α-axis horizontal zeroing, β-axis fixed angle servo mode to the sun".

[0050] If ∠α'=0, ∠β'=0, W0<W, and the α-axis is defined as three fixed angles: ∠α0', 0°, -∠α0'. ∠α0' is:

[0051]

[0052] Furthermore, in the α-axis control mode setting module, if ∠α0'<ε, the α-axis of the solar wing is set to a fixed angle servo control mode, and the following steps are performed:

[0053] When ∠β'≠0, the α-axis keeps the horizontal zero state. When ∠β'=0, the α-axis fixed angle servo to the sun is controlled, and its fixed rotation angle ∠α0' is calculated. The control strategy of the α-axis fixed angle servo to the sun is as follows: when the space station enters the sunlit area, the α-axis is rotated so that ∠α'=∠α0'; when the solar incident angle ∠θ=90°-2∠α0', the α-axis is rotated so that ∠α'=0°; when the solar incident angle ∠θ=-(90°-∠α0'), the α-axis is rotated so that ∠α'=-∠α0'; after entering the shadow area, the α-axis is controlled to return to the horizontal zero state; the α-axis sun orientation device rotates at a speed of 0.300° / s, and the rotation direction is based on the principle of proximity.

[0054] Beneficial effects:

[0055] The present invention designs a solar orientation control strategy for a space station experimental module solar wing dual-axis coupled fixed-angle servo to the sun. On the basis of the β-axis fixed-angle servo control mode, if the α-axis horizontal zeroing cannot meet the power demand, the α-axis also adopts a fixed-angle servo control mode. The present invention innovatively designs a control mode in which the two degrees of freedom of the solar wing of the space station experimental module simultaneously adopt a fixed-angle servo to the sun, that is, a solar orientation control strategy of dual-axis coupled fixed-angle servo to the sun. The present invention innovatively designs a method for calculating the fixed angle value ∠α0' in the α-axis fixed-angle servo control mode. The present invention achieves the following three effects:

[0056] The first is the power demand, that is, the present invention aims to "ensure that the power generated by the solar wing can meet the power demand of the entire station", and the design purpose is not to maximize power.

[0057] The second is to reduce mechanical wear, that is, to reduce mechanical wear by reducing the rotation of the solar panels, thereby extending the life of the solar panels on the space station.

[0058] In addition, the risk of collision is reduced. The low-orbit environment of the space station is becoming increasingly complex. The solar panels of the Tianhe core module have been damaged by debris impact. The on-orbit management of the space station must consider the risk of reduced power generation capacity due to collisions with space debris. Since the solar panels must be exposed to sunlight and cannot be protected by reinforcement and other means, the on-orbit control of the solar panels must be precisely and cleverly designed to reduce physical damage while meeting power requirements, so as to meet the mission requirements of the long-life operation of the space station. BRIEF DESCRIPTION OF THE DRAWINGS

[0059] Figure 1 Schematic diagram of the dual-degree-of-freedom of the solar wing of the space station experimental module (the A axis is the α axis in the present invention, and the B axis is the β axis in the present invention);

[0060] Figure 2 This is the flow chart of the dual-axis coupling fixed-angle servo strategy. DETAILED DESCRIPTION

[0061] The present invention is described in detail below with reference to the accompanying drawings and embodiments.

[0062] Embodiment 1:

[0063] The present invention provides a solar orientation control strategy for a space station solar wing dual-axis coupled fixed-angle servo to the sun, that is, the solar wing dual degrees of freedom are both controlled by fixed rotation angle values ​​and intermittent rotation.

[0064] The design process of the present invention is as follows: first, the β-axis is designed to be in a fixed-angle servo control mode, and the fixed angle of rotation thereof is in a total of 2n+1 gears, and the value range is {-nΔ, ..., -3Δ, -2Δ, -Δ, 0, Δ, 2Δ, 3Δ, ..., nΔ} (n is the value of n determined according to the mission state, and is positive or negative 66 degrees according to the change of the solar altitude angle, to prevent frequent rotation of the β-axis. In consideration of the power demand of the space station, it can be set to n≤6; Δ is the gear difference, which is a multiple of 10). When the β-axis is not zeroed, the α-axis is set to zero horizontally; when the β-axis is zeroed, if the horizontal zeroing of the α-axis cannot meet the power demand, the α-axis is designed to be in a fixed-angle servo control mode, and the fixed angle of rotation thereof is in a total of 3 gears, and the value range is {-∠α0', 0, ∠α0'}, (∠α0' is a pre-set fixed value, which can be obtained by referring to the data table lookup fitting in Table 3).

[0065] The symbols are defined as follows:

[0066] W: space station power requirement, which varies with mission status;

[0067] W 总 : The maximum power generation of the solar wing is a certain value;

[0068] W0: The power generation capacity of the solar wing, which varies with the solar incidence angle;

[0069] ∠α': The rotation angle of the solar wing in the α degree of freedom. When the solar wing α axis is oriented to the sun, ∠α' is defined as 0° when the spacecraft is in an orbital angle of 0°, that is, the horizontal zero state. ∠α' can vary in the range of 0° to 360°;

[0070] ∠β': ∠β' is the rotation angle of the solar wing in the β degree of freedom, that is, the sailboard rotation angle. It is defined that when the normal line of the solar wing is pointing to the zenith direction of the orbital plane along the line connecting the earth and the spacecraft, ∠β' is 0°. ∠β' can vary within the range of -66° to +66°;

[0071] ∠β: solar altitude angle, the angle between the sun-earth line and the orbital plane, which varies between ±(i+γ), where i is the orbital inclination with the equatorial plane as the reference plane, and γ is the ecliptic angle of 23.45°;

[0072] ∠θ: solar incident angle, the incident angle of sunlight on the solar wing, that is, the angle between the normal line of the solar wing surface and the solar incident vector. ∠θ varies in the range of 0° to 180°. When 0°≤∠θ≤90°, it means that the sun is shining on the front of the solar wing; when 90°≤∠θ≤180°, it means that the sun is shining on the back of the solar wing. The solar wing can be tilted to make the sunlight shine directly. The generation of electrical energy of the solar wing is proportional to the component of the light intensity vector in the direction of the normal line of the solar wing, that is, it is proportional to cos∠θ; when ∠θ=0°, the generated electrical energy is the largest; when ∠θ=90°, no electrical energy is generated.

[0073] Δ: gear position difference of ∠β', that is, the angle difference between two adjacent gear positions of ∠β';

[0074] δ: ∠β' gear adjustment threshold. When the difference between ∠β' and ∠β exceeds δ, the β-axis servo is controlled to rotate to the next gear.

[0075] ∠α0': The fixed angle value of ∠α'. When the α-axis adopts the fixed angle servo mode to the sun, the value of ∠α' includes -∠α0', 0°, and ∠α0'.

[0076] The dual-axis coupling fixed angle servo strategy process is as follows Figure 2 As shown, the specific steps include:

[0077] S1: Clarify the solar altitude angle range. According to the definition of solar altitude angle, the solar altitude angle range of the space station is -66° to 66°.

[0078] S2: Preliminarily determine the gear difference Δ and angle adjustment threshold δ of the β-axis to sun fixed angle servo.

[0079] S21: Determine the value of Δ. The fixed angle value range of the β-axis rotation is {-nΔ,…,-3Δ,-2Δ,-Δ,0,Δ,2Δ,3Δ,…,nΔ}. Since the β-axis is used to compensate for the influence of the solar altitude angle on the solar incident angle, and the solar altitude angle range of the space station is -66° to 66°, nΔ≤66°.

[0080] S22: Determine the value of δ. According to the definition of δ, it can be known that δ<Δ. It can be taken that δ=14°, and Δ=20°.

[0081] S23: The β-axis fixed angle servo control strategy of the experimental cabin solar wing is as follows:

[0082] When the solar altitude angle is within the range of (-δ, δ), rotate the β axis so that ∠β'=0°. As the solar altitude angle increases, when ∠β=δ, rotate the β axis so that ∠β'=Δ; when ∠β=Δ+δ, rotate the β axis so that ∠β'=2Δ; and so on, until ∠β'=nΔ. As the solar altitude angle decreases, when ∠β=-δ, rotate the β axis so that ∠β'=-Δ; when ∠β=-Δ-δ, rotate the β axis so that ∠β'=-2Δ; and so on, until ∠β'=-nΔ.

[0083] The corresponding relationship between ∠β and ∠β' is shown in Table 1.

[0084] Table 1 Correspondence between solar altitude angle and sailboard rotation angle ∠β'

[0085]

[0086] Secondly, according to the power demand of the space station W and the maximum power generation of the solar wing W 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun.

[0087] S3: Get the maximum power generation W of the solar wing from the engineering developer 总 The numerical value of .

[0088] S4: Calculate the current power demand W of the space station. The power generated by the solar panels of the space station is mainly used to power the platform itself, charge the battery pack, power the application load, and power the docked aircraft. After calculating the power demand, the total is calculated and a margin (5%) is left to obtain W.

[0089] S5: Calculate the fixed angle value ∠α0' of the α-axis.

[0090] S51: When ∠α'=0 and ∠β'=0, if the solar panel power generation meets the power demand of the space station, the α-axis horizontal zeroing state is maintained. In this mission state, the solar panel always adopts the solar orientation control strategy of "α-axis horizontal zeroing, β-axis fixed angle servo mode".

[0091] S52: If ∠α'=0, ∠β'=0, W0<W, and three fixed angles of the α-axis are defined: ∠α0', 0°, and -∠α0′.

[0092] ∠α0' is determined according to the following formula:

[0093] but,

[0094] Again, the maximum threshold of ∠α0' is set to ε. If ∠α0'<ε, the solar panel power generation cannot meet the power demand of the space station, and the α-axis is set to the continuous sun-orientation control mode. When ∠α0'<ε and the solar panel power generation can meet the demand, the α-axis fixed angle servo control strategy is clarified.

[0095] S6: Determine whether ∠α0' is less than ε (80°). If ∠α0'>ε, the α-axis is controlled to continuously orient toward the sun according to the orbital angular velocity; if ∠α0'<ε, the solar wing α-axis fixed angle servo control strategy is executed according to the following steps:

[0096] S61: When ∠β'≠0, the α-axis is kept in a horizontal zero state. When ∠β'=0, the α-axis is controlled to have a fixed angle servo relative to the sun, and its fixed rotation angle ∠α0' is calculated in step S52.

[0097] S62: The α-axis fixed angle servo control strategy for the sun is: when the space station enters the sunlit area, rotate the α-axis so that ∠α'=∠α0'; when the solar incident angle ∠θ=90°-2∠α0', rotate the α-axis so that ∠α'=0°; when the solar incident angle ∠θ=-(90°-∠α0'), rotate the α-axis so that ∠α'=-∠α0'; after entering the shadow area, control the α-axis to return to zero horizontally. The α-axis sun orientation device rotates at a speed of 0.300° / s, and the rotation direction is based on the principle of proximity (360° rotation).

[0098] Finally, if ∠β'=Δ and the α-axis is horizontally zeroed, the solar wing power generation capacity cannot meet the power demand of the space station, then the gear difference Δ and threshold δ of the β-axis are increased until the mission requirements are met.

[0099] S7: When ∠β' is just adjusted to Δ or -Δ, calculate the solar wing power generation capacity when the α axis is horizontally zeroed to evaluate whether it meets the mission requirements. If it can meet the mission requirements, use the current values ​​of Δ and δ and the process ends. If it cannot meet the mission requirements, execute S8.

[0100] S8: Increase the gear difference Δ of the β-axis angle value and the value of the adjustment threshold δ to postpone the timing of the β-axis gear adjustment, delay the time for the α-axis to return to zero horizontally, increase the solar altitude angle to ensure the power output of the solar wing, and execute step S7 until the solar wing's power generation capacity meets the mission requirements when the α-axis returns to zero horizontally.

[0101] The present invention achieves the following effects:

[0102] 1. Reduce the rotation time of the solar panels to reduce the mechanical wear caused by continuous rotation.

[0103] In the α-axis continuous solar orientation mode, the rotation time within one orbital period is equal to the orbital period, about 92 minutes, which can be expressed by the formula t'=4*∠α0' / 0.065 (∠α0'=90°).

[0104] In the present invention, the rotation time of the α-axis in one orbital period is: t=4*∠α0' / 0.3 (∠α0'<90°).

[0105] By comparing the formula, we know that t<t'. That is, the rotation time of the α axis in one orbital cycle in the present invention is significantly shorter than the rotation time in the continuous solar orientation mode, thus effectively reducing the mechanical wear on the solar wing rotation mechanism.

[0106] 2. Reduce the cumulative projected area of ​​the space station's solar panels on the plane orthogonal to the orbital plane (the windward side) to reduce the risk of being hit by space debris in the low-orbit space environment.

[0107] In the continuous sun tracking mode, the cumulative projection area A' of the solar panel on the plane orthogonal to the orbital plane within one orbital cycle can be calculated by the following formula:

[0108]

[0109] but,

[0110] A′=3507.4A0.

[0111] Wherein, A0 is the solar wing area, and v1=0.065(° / s)=0.065π / 180(rad / s) is the orbital angular velocity.

[0112] In the present invention, the cumulative projected area A of the solar wing on the plane orthogonal to the orbital plane within one orbital period can be calculated by the following formula:

[0113]

[0114] Among them, v2 = 0.3 (° / s) = 0.3π / 180 (rad / s) is the rapid rotation speed of the solar wing.

[0115] When ∠α0' takes different values, the resulting values ​​of A are shown in Table 2 below:

[0116] Table 2 Cumulative projection area results

[0117]

[0118] By comparing A and A', it can be seen that when ∠α0'<80°, within one orbital period, the cumulative projection area (3055.4A0) of the solar wing in the present invention is significantly smaller than the cumulative projection area (3507.4A0) in the continuous sun tracking mode, and the smaller the value of ∠α0', the more obvious the advantage.

[0119] Table 3 Minimum power generation percentage of solar panels

[0120]

[0121] As shown in Table 3, when ∠α0'=80°, the control strategy of the present invention is adopted, and the minimum power generation of the solar wing is 98.5% of the maximum power. Assuming that the power margin of the solar wing of the space station is 20%, the value of ∠α0' is between 50° and 60°, and the calculated cumulative projected area is reduced by about 43.5% to 58.6%, and the collision risk is reduced by about 50%. The greater the power margin, the smaller the value of ∠α0', and the lower the risk of the solar wing being collided with space debris under this control strategy.

[0122] Embodiment 2:

[0123] Based on the method provided in the above embodiment 1, this embodiment provides a dual-axis coupled sun-facing fixed angle servo control system for an experimental cabin solar wing for implementing the above method, including the following eight modules, namely:

[0124] The solar altitude angle range determination module is used to determine the solar altitude angle range of the space station;

[0125] The β-axis preliminary setting module is used to preliminarily determine the gear difference Δ and angle adjustment threshold δ of the β-axis fixed angle servo to the sun; set the β-axis of the solar wing of the experimental cabin to the fixed angle servo control mode;

[0126] Power generation acquisition module, used to obtain the maximum power generation W of the solar wing 总 The numerical value of .

[0127] The power demand calculation module of the space station is used to calculate the power demand W of the current space station;

[0128] The α-axis fixed angle calculation module is used to calculate the power demand W of the space station and the maximum power generation power W of the solar wing. 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun;

[0129] The α-axis control mode setting module is used to determine whether ∠α0' is less than ε. If ∠α0'>ε, the α-axis of the solar wing is set to a continuous sun-orientation control mode; if ∠α0'<ε, the α-axis of the solar wing is set to a fixed-angle servo control mode;

[0130] The evaluation module is used to calculate the power generation capacity of the solar wing when the α-axis is horizontally zeroed when ∠β' is just adjusted to Δ or -Δ, and evaluate whether it meets the task requirements. If it meets the task requirements, the current values ​​of Δ and δ are adopted and the process ends; if it does not meet the task requirements, the β-axis adjustment module is executed;

[0131] The β-axis adjustment module is used to increase the gear difference Δ of the β-axis angle value and the value of the adjustment threshold δ to postpone the timing of the β-axis gear adjustment, delay the time for the α-axis to return to zero horizontally, increase the solar altitude angle to ensure the power output of the solar wing, and return to execution module seven until the solar wing's power generation capacity meets the mission requirements when the α-axis returns to zero horizontally.

[0132] The β-axis preliminary setting module specifically performs the following steps:

[0133] S21: Determine the value of Δ, where the fixed angle value range of the β-axis rotation is {-nΔ, …, -3Δ, -2Δ, -Δ, 0, Δ, 2Δ, 3Δ, …, nΔ}. Since the β-axis is used to compensate for the influence of the solar altitude angle on the solar incident angle, and the solar altitude angle range of the space station is -66° to 66°, nΔ≤66°;

[0134] S22: Determine the value of δ, δ<Δ;

[0135] S23: The servo control strategy for the fixed angle of the β axis of the solar wing of the experimental cabin is as follows: when the solar altitude angle is within the range of (-δ, δ), the β axis is rotated so that ∠β'=0°; as the solar altitude angle increases, when ∠β=δ, the β axis is rotated so that ∠β'=Δ; when ∠β=Δ+δ, the β axis is rotated so that ∠β'=2Δ; and so on, until ∠β'=nΔ. As the solar altitude angle decreases, when ∠β=-δ, the β axis is rotated so that ∠β'=-Δ; when ∠β=-Δ-δ, the β axis is rotated so that ∠β'=-2Δ; and so on, until ∠β'=-nΔ.

[0136] In the power demand calculation module of the space station, the current power demand W of the space station is calculated. Specifically, the power generated by the solar panels of the space station is used to power its own platform, charge the battery pack, consume electricity for application payloads, and power the docked aircraft. After calculating each power demand, the total is calculated and a margin is left to obtain W.

[0137] The α-axis fixed angle calculation module specifically performs the following process:

[0138] When ∠α'=0 and ∠β'=0, if the solar panel power generation meets the power demand of the space station, the α-axis is kept horizontally zeroed; in this mission state, the solar panel always adopts the solar orientation control strategy of "α-axis horizontally zeroed, β-axis fixed angle servo mode";

[0139] If ∠α'=0, ∠β'=0, W0<W, and the α-axis is defined as three fixed angles: ∠α0', 0°, -∠α0'; ∠α0' is:

[0140]

[0141] In the α-axis control mode setting module, if ∠α0'<ε, the α-axis of the solar wing is set to a fixed angle servo control mode, and the following steps are performed:

[0142] When ∠β'≠0, the α-axis keeps the horizontal zero state. When ∠β'=0, the α-axis fixed angle servo to the sun is controlled, and its fixed rotation angle ∠α0' is calculated. The control strategy of the α-axis fixed angle servo to the sun is as follows: when the space station enters the sunlit area, the α-axis is rotated so that ∠α'=∠α0'; when the solar incident angle ∠θ=90°-2∠α0', the α-axis is rotated so that ∠α'=0°; when the solar incident angle ∠θ=-(90°-∠α0'), the α-axis is rotated so that ∠α'=-∠α0'; after entering the shadow area, the α-axis is controlled to return to the horizontal zero state; the α-axis sun orientation device rotates at a speed of 0.300° / s, and the rotation direction is based on the principle of proximity.

[0143] In summary, the above are only preferred embodiments of the present invention and are not intended to limit the protection scope of the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. The experimental cabin solar wing dual-axis coupling fixed angle servo control method is characterized by: The steps include: S1: Determine the solar altitude angle range of the space station; S2: Preliminarily determine the gear difference Δ and angle adjustment threshold δ of the β-axis fixed angle servo to the sun; set the β-axis of the solar wing in the experimental cabin to the fixed angle servo control mode; S3: Get the maximum power generation W of the solar wing 总 The value of S4: Calculate the power demand W of the current space station; S5: Based on the power demand of the space station W and the maximum power generation of the solar wing W 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun; S6: Determine whether ∠α0' is less than ε. If ∠α0'>ε, the α-axis of the solar wing is set to a continuous sun-orientation control mode; if ∠α0'<ε, the α-axis of the solar wing is set to a fixed-angle servo control mode; S7: When ∠β' is just adjusted to Δ or -Δ, the power generation capacity of the solar wing is calculated when the α axis is horizontally zeroed, and whether it meets the mission requirements is evaluated. If it meets the mission requirements, the current values ​​of Δ and δ are adopted, and the process ends; If the task requirements cannot be met, execute S8; S8: Increase the gear difference Δ of the β-axis angle value and the value of the adjustment threshold δ, so as to postpone the timing of the β-axis gear adjustment, delay the time for the α-axis to return to zero horizontally, increase the solar altitude angle to ensure the power output of the solar wing, and return to execute step S7 until the solar wing's power generation capacity meets the mission requirements when the α-axis returns to zero horizontally.

2. The method for controlling the fixed angle of the dual-axis coupling solar wing of the experimental cabin according to claim 1 is characterized in that: S2: preliminarily determining the gear difference Δ and angle adjustment threshold δ of the β-axis fixed angle servo to the sun, and setting the β-axis of the solar wing in the experimental cabin to a fixed angle servo control mode, specifically includes the following steps: S21: Determine the value of Δ, where the fixed angle value range of the β-axis rotation is {-nΔ, …, -3Δ, -2Δ, -Δ, 0, Δ, 2Δ, 3Δ, …, nΔ}. Since the β-axis is used to compensate for the influence of the solar altitude angle on the solar incident angle, and the solar altitude angle range of the space station is -66° to 66°, nΔ≤66°; S22: Determine the value of δ, δ<Δ; S23: The fixed-angle servo control strategy of the β-axis of the solar wing of the experimental cabin is as follows: when the solar altitude angle is within the range of (-δ,δ), rotate the β-axis so that ∠β'=0°; as the solar altitude angle increases, when ∠β=δ, rotate the β-axis so that ∠β'=Δ; when ∠β=Δ+δ, rotate the β-axis so that ∠β'=2Δ, until ∠β'=nΔ; as the solar altitude angle decreases, when ∠β=-δ, rotate the β-axis so that ∠β'=-Δ; when ∠β=-Δ-δ, rotate the β-axis so that ∠β'=-2Δ; until ∠β'=-nΔ.

3. The method for controlling the fixed angle of the dual-axis coupling solar wing of the experimental cabin according to claim 1 is characterized in that: S4: Calculate the power demand W of the current space station, specifically: the power generated by the solar panels of the space station is used to power the platform itself, charge the battery pack, power the application load, and power the docked aircraft. After calculating the various power demands, the total is calculated and a margin is left to obtain W.

4. The method for controlling the fixed angle of the dual-axis coupling solar wing of the experimental cabin according to claim 1 is characterized in that: S5: According to the power demand W of the space station and the maximum power generation power W of the solar wing 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun, specifically: When ∠α'=0 and ∠β'=0, if the solar panel power generation meets the power demand of the space station, the α-axis will be kept in the horizontal zero state; in this mission state, the solar panel always adopts the solar orientation control strategy of "α-axis horizontal zeroing, β-axis fixed angle servo mode"; If ∠α'=0, ∠β'=0, W0<W, and the α-axis is defined as three fixed angles: ∠α0', 0°, -∠α0'; ∠α0' is:

5. The method for controlling the fixed angle of the dual-axis coupling solar wing of the experimental cabin according to claim 1 is characterized in that: If ∠α0'<ε, the α axis of the solar wing is set to a fixed angle servo control mode, and the following steps are performed: When ∠β'≠0, the α-axis keeps the horizontal zero state. When ∠β'=0, the α-axis fixed angle servo to the sun is controlled, and its fixed rotation angle ∠α0' is calculated. The α-axis fixed angle servo control strategy to the sun is as follows: when the space station enters the sunlit area, the α-axis is rotated so that ∠α'=∠α0'; when the solar incident angle ∠θ=90°-2∠α0', the α-axis is rotated so that ∠α'=0°; when the solar incident angle ∠θ=-(90°-∠α0'), the α-axis is rotated so that ∠α'=-∠α0'; After entering the shadow area, the α axis is controlled to return to zero horizontally; the α solar orientation device rotates at a speed of 0.300° / s, and the rotation direction is based on the principle of proximity.

6. The experimental cabin solar wing dual-axis coupling fixed angle servo control system is characterized by: It includes the following eight modules: The solar altitude angle range determination module is used to determine the solar altitude angle range of the space station; The β-axis preliminary setting module is used to preliminarily determine the gear difference Δ and angle adjustment threshold δ of the β-axis to the sun fixed angle servo; Set the β-axis of the solar wing in the experimental cabin to a fixed-angle servo control mode; Power generation acquisition module, used to obtain the maximum power generation W of the solar wing 总 The value of The power demand calculation module of the space station is used to calculate the power demand W of the current space station; The α-axis fixed angle calculation module is used to calculate the power demand W of the space station and the maximum power generation power W of the solar wing. 总 , calculate the fixed angle value ∠α0' of the α-axis fixed angle servo to the sun; The α-axis control mode setting module is used to determine whether ∠α0' is less than ε. If ∠α0'>ε, the α-axis of the solar wing is set to a continuous sun-orientation control mode; if ∠α0'<ε, the α-axis of the solar wing is set to a fixed-angle servo control mode; The evaluation module is used to calculate the power generation capacity of the solar wing when the α-axis is horizontally zeroed when ∠β' is just adjusted to Δ or -Δ, and evaluate whether it meets the task requirements. If it meets the task requirements, the current values ​​of Δ and δ are adopted and the process ends; if it does not meet the task requirements, the β-axis adjustment module is executed; The β-axis adjustment module is used to increase the gear difference Δ of the β-axis angle value and the value of the adjustment threshold δ to postpone the timing of the β-axis gear adjustment, delay the time for the α-axis to return to zero horizontally, increase the solar altitude angle to ensure the power output of the solar wing, and return to execution module seven until the solar wing's power generation capacity meets the mission requirements when the α-axis returns to zero horizontally.

7. The experimental cabin solar wing dual-axis coupling fixed-angle servo control system according to claim 6 is characterized in that: The β-axis preliminary setting module specifically performs the following steps: S21: Determine the value of Δ, where the fixed angle value range of the β-axis rotation is {-nΔ, …, -3Δ, -2Δ, -Δ, 0, Δ, 2Δ, 3Δ, …, nΔ}. Since the β-axis is used to compensate for the influence of the solar altitude angle on the solar incident angle, and the solar altitude angle range of the space station is -66° to 66°, nΔ≤66°; S22: Determine the value of δ, δ<Δ; S23: The fixed-angle servo control strategy of the β-axis of the solar wing of the experimental cabin is as follows: when the solar altitude angle is within the range of (-δ,δ), rotate the β-axis so that ∠β'=0°; as the solar altitude angle increases, when ∠β=δ, rotate the β-axis so that ∠β'=Δ; when ∠β=Δ+δ, rotate the β-axis so that ∠β'=2Δ; until ∠β'=nΔ; as the solar altitude angle decreases, when ∠β=-δ, rotate the β-axis so that ∠β'=-Δ; when ∠β=-Δ-δ, rotate the β-axis so that ∠β'=-2Δ; until ∠β'=-nΔ.

8. The experimental cabin solar wing dual-axis coupling fixed-angle servo control system according to claim 6 is characterized in that: In the power demand calculation module of the space station, the power demand W of the current space station is calculated. Specifically, the power generated by the solar panels of the space station is used to power its own platform, charge the battery pack, consume electricity for application loads, and power the docked aircraft and connect to the grid. After calculating each power demand, the total is calculated and a margin is left to obtain W.

9. The experimental cabin solar wing dual-axis coupling fixed-angle servo control system according to claim 6, characterized in that: The α-axis fixed angle calculation module specifically performs the following process: When ∠α'=0 and ∠β'=0, if the solar panel power generation meets the power demand of the space station, the α-axis will be kept in the horizontal zero state; in this mission state, the solar panel always adopts the solar orientation control strategy of "α-axis horizontal zeroing, β-axis fixed angle servo mode"; If ∠α'=0, ∠β'=0, W0<W, and the α-axis is defined as three fixed angles: ∠α0', 0°, -∠α0'; ∠α0' is:

10. The experimental cabin solar wing dual-axis coupling fixed-angle servo control system according to claim 6, characterized in that: In the α-axis control mode setting module, if ∠α0'<ε, the α-axis of the solar wing is set to a fixed angle servo control mode, and the following steps are performed: When ∠β'≠0, the α-axis keeps the horizontal zero state. When ∠β'=0, the α-axis fixed angle servo to the sun is controlled, and its fixed rotation angle ∠α0' is calculated. The α-axis fixed angle servo control strategy to the sun is as follows: when the space station enters the sunlit area, the α-axis is rotated so that ∠α'=∠α0'; when the solar incident angle ∠θ=90°-2∠α0', the α-axis is rotated so that ∠α'=0°; when the solar incident angle ∠θ=-(90°-∠α0'), the α-axis is rotated so that ∠α'=-∠α0'; After entering the shadow area, the α axis is controlled to return to zero horizontally; the α solar orientation device rotates at a speed of 0.300° / s, and the rotation direction is based on the principle of proximity.

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