Satellite long-term sun-pointing three-axis pointing attitude design method under inter-satellite communication constraints

Abstract

The application discloses a satellite long-term sun-pointing three-axis pointing attitude design method under inter-satellite communication constraints, including the constraint conditions of comprehensively considering sun-pointing, inter-satellite communication pointing and the drastic change of sunlight conditions under low-inclination orbit, proposing a nominal expected attitude design method, a modified expected attitude design method, a transition expected attitude design method, a flip nominal expected attitude design method and a flip modified expected attitude design method, and giving an expected angular velocity calculation method of manipulation aiming at the corresponding expected attitude. The application can meet the basic condition that the satellite solar panel normal direction points to the sun, ensure the communication pointing requirements between space formation satellites at the same time, and avoid the influence of the polarity change of the satellite on the low-inclination orbit, thereby providing a stable and safe long-term sun-pointing attitude for the satellite.

Description

TECHNICAL FIELD

[0001] The application belongs to the field of satellite attitude dynamics and control, and particularly relates to a satellite long-term sun-pointing three-axis pointing attitude design method under inter-satellite communication constraints. BACKGROUND

[0002] Attitude determination and control is a necessary platform function of modern satellites, and is an important prerequisite for ensuring satellite energy supply, thermal control stability and completing on-orbit tasks. Sun-pointing control of a satellite is the most basic control mode, and its main purpose is to maintain the normal direction of the solar panel of the satellite and point to the sun, thereby ensuring that the satellite can obtain sufficient energy. Considering that sun-pointing is the attitude maintained by a satellite in most time on orbit, and modern satellites generally adopt a three-axis stable attitude control mode, therefore, a long-term three-axis stable sun-pointing attitude is an important problem that needs to be solved first in the satellite attitude control design process. Generally, the three-axis stable sun-pointing attitude has only a strong constraint in the direction of one axis, that is, the normal direction of the solar panel needs to be parallel to the sun direction vector. In a multiple-satellite formation flight mission, it is required to maintain a long-term continuous communication link between satellites, which increases the pointing constraint of the satellite during long-term sun-pointing. Meanwhile, considering a satellite running in a low-inclination orbit, the relative relationship between the sun direction vector and the normal direction of the satellite orbit plane has positive and negative polarity changes, which also poses a challenge to the design of a long-term stable sun-pointing attitude. It is necessary to design a satellite long-term stable sun-pointing three-axis pointing attitude under the comprehensive constraints of the above multiple constraints. SUMMARY

[0003] The technical problem to be solved by the application is to provide a satellite long-term sun-pointing three-axis pointing attitude design method under inter-satellite communication constraints, which can meet the basic condition that the normal direction of the solar panel of a satellite points to the sun, while ensuring the communication pointing requirements between space formation satellites, and can avoid the influence of the polarity change of the sunning condition of a satellite in a low-inclination orbit, thereby providing a stable and safe long-term sun-pointing attitude for the satellite.

[0004] To solve the above technical problem, the technical scheme adopted by the application is as follows:

[0005] A satellite long-term sun-pointing three-axis pointing attitude design method under inter-satellite communication constraints comprises the following steps: obtaining a sun direction angle and a satellite motion state; when the sun direction angle is greater than 0 and the unit direction vector of the sun-pointing satellite in the inter-satellite communication star is in the coverage range of the inter-satellite communication beam in the satellite motion state, simultaneously, a nominal desired attitude M1 is designed as the current desired attitude of the satellite, and the function expression of the nominal desired attitude M1 is as follows:

[0006]

[0007] In the above formula, M1 is the nominal desired attitude, x b ,yb and z b is the satellite body coordinate axis, and has:

[0008]

[0009] y b = -r s

[0010]

[0011] In the above formula, r S is a unit vector of the sun direction, and p is a unit direction vector of the satellite pointing to the inter-satellite communication satellite.

[0012] Optionally, when the sun direction angle is greater than 0, and the unit direction vector of the satellite pointing to the inter-satellite communication satellite is not within the coverage of the inter-satellite communication beam, and the unit direction vector of the satellite pointing to the inter-satellite communication satellite is not approximately parallel to the unit vector of the sun direction in the satellite motion state, a first correction matrix ΔM2 is multiplied by the nominal expected attitude M1 to obtain a corrected expected attitude M2 as the current expected attitude of the satellite, wherein the function expression of the first correction matrix ΔM2 is:

[0013] ΔM2 = R z (ν)

[0014] In the above formula, the function R z (ν) represents a conversion matrix formed by rotating the satellite z b by an angle v, and the calculation formula of the angle v is:

[0015]

[0016] In the above formula, η is the included angle between the unit direction vector of the satellite pointing to the inter-satellite communication satellite and the unit vector of the sun direction, a is the half-cone angle of the inter-satellite communication beam, and a < 90°, and δ is a preset deviation amount.

[0017] Optionally, when the sun direction angle is greater than 0, and the unit direction vector of the satellite pointing to the inter-satellite communication satellite is not within the coverage of the inter-satellite communication beam, and the unit direction vector of the satellite pointing to the inter-satellite communication satellite is approximately parallel to the unit vector of the sun direction, the long-term sun-pointing three-axis pointing attitude of the satellite at the previous moment is obtained as a transition expected attitude, and the transition expected attitude is taken as the current expected attitude of the satellite.

[0018] Optionally, when the sun direction angle is less than or equal to 0, and the unit direction vector of the satellite pointing to the inter-satellite communication satellite is within the coverage of the inter-satellite communication beam, a flipped nominal expected attitude M′1 is obtained by polarity adjustment on the basis of the nominal expected attitude M1 as the current expected attitude of the satellite.

[0019] Optionally, the function expression of the flipped nominal desired attitude M'1 is:

[0020]

[0021] In the above formula, M'1 is the flipped nominal desired attitude, x b , y b and z b are the satellite body coordinate axes, and have:

[0022]

[0023] y b = -r s

[0024]

[0025] In the above formula, r S is the unit vector of the sun direction, and p is the unit direction vector of the home-pointing inter-satellite communication star.

[0026] Optionally, when the sun direction angle is less than or equal to 0, and the unit direction vector of the home-pointing inter-satellite communication star in the satellite motion state is not within the coverage of the inter-satellite communication beam, and the unit direction vector of the home-pointing inter-satellite communication star and the unit vector of the sun direction are not approximately parallel, the modified desired attitude M2 is obtained as the current desired attitude of the satellite on the basis of the flipped nominal desired attitude M'1 by multiplying the second correction matrix AM'2, wherein the function expression of the second correction matrix AM'2 is:

[0027] AM'2 = R z (ν')

[0028] In the above formula, the function R z (ν') represents a conversion matrix formed by rotating the satellite z b axis by an angle v', and the calculation formula of the angle v' is:

[0029]

[0030] In the above formula, η is the included angle between the unit direction vector of the home-pointing inter-satellite communication star and the unit vector of the sun direction, a is the half-cone angle of the inter-satellite communication beam, and a < 90°, and δ is a preset deviation amount.

[0031] Optionally, the method further comprises: when the sun direction angle is less than or equal to 0, and the unit direction vector of the sun in the satellite motion state points to the inter-satellite communication satellite, and the unit direction vector of the sun is approximately parallel to the unit direction vector of the sun, obtaining the satellite long-term three-axis sun-pointing attitude at the previous time as a transition expected attitude, and taking the transition expected attitude as the current expected attitude.

[0032] Optionally, the method further comprises: after obtaining the current expected attitude, calculating an expected angular velocity ω * , and when the sun direction angle is greater than 0, the function expression of the expected angular velocity ω * is:

[0033]

[0034]

[0035]

[0036] When the sun direction angle is less than or equal to 0, the function expression of the expected angular velocity ω * is:

[0037]

[0038]

[0039]

[0040] wherein x b , y b and z b are body coordinate axes of the satellite, r S is a unit vector of the sun direction, p is a unit direction vector of the sun pointing to the inter-satellite communication satellite, is a differential of y b , is a differential of r S , is a differential of z b .

[0041] In addition, the application further provides a satellite long-term three-axis sun-pointing attitude design system under inter-satellite communication constraints, comprising a microprocessor and a memory connected to each other, the microprocessor is programmed or configured to execute the steps of the satellite long-term three-axis sun-pointing attitude design method under inter-satellite communication constraints, or the memory stores a computer program programmed or configured to execute the satellite long-term three-axis sun-pointing attitude design method under inter-satellite communication constraints.

[0042] In addition, the application also provides a computer readable storage medium, which stores a computer program programmed or configured to perform the satellite long-term sun-pointing three-axis pointing attitude design method under the inter-satellite communication constraint.

[0043] Compared with the prior art, the technical scheme of the application has the following beneficial effects: the application can meet the basic condition that the satellite solar panel normal direction points to the sun, ensure the communication pointing requirement between the space formation satellites, avoid the influence of the polarity change of the sun exposure condition on the satellite in the low-inclination orbit, provide a stable and safe long-term sun-pointing attitude for the satellite, ensure the attitude pointing condition of the satellite energy safety, meet the continuity requirement of the formation communication, and ensure the stable sun exposure condition of the satellite. BRIEF DESCRIPTION OF DRAWINGS

[0044] In order to more clearly illustrate the technical solutions in the embodiments of the application or the prior art, the following will briefly introduce the drawings needed to be used in the embodiments or the prior art description. Obviously, the drawings in the following description only show some embodiments of the application, and for those skilled in the art, other drawings can also be obtained from the structures shown in the drawings without creative labor.

[0045] Figure 1 The expected attitude design branch flow under different conditions in the embodiments of the application.

[0046] Figure 2 The relative geometric relationship between the sun vector direction and the communication pointing in the embodiments of the application.

[0047] Figure 3 The geometric relative relationship between the antenna pointing and the sun vector direction when the sun direction angle is less than 0 in the embodiments of the application.

[0048] Figure 4 The satellite sun-pointing deviation simulation result in the embodiments of the application.

[0049] Figure 5 The satellite communication pointing deviation simulation result in the embodiments of the application. DETAILED DESCRIPTION

[0050] In order to make the technical problems, technical solutions and beneficial effects of the application more clearly understood, the following will further describe the application in combination with the drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the application, and are not used to limit the application.

[0051] Referring to Figure 1The long-term solar-oriented three-axis pointing attitude design method for satellites under inter-satellite communication constraints in this embodiment includes obtaining the solar azimuth angle and the satellite's motion state. When the solar azimuth angle is greater than 0 and the unit direction vector of the satellite pointing to the inter-satellite communication satellite in the satellite's motion state is simultaneously within the coverage area of ​​the inter-satellite communication beam (i.e., η∈[90°-α,90°+α]), the nominal desired attitude M1 is designed as the current desired satellite attitude, and the functional expression of the nominal desired attitude M1 is:

[0052]

[0053] In the above formula, M1 is the nominal desired attitude, x b ,y b and z b Let be the coordinate axes of the satellite body, and we have:

[0054]

[0055] y b =-r s

[0056]

[0057] In the above formula, r S ρ is the unit vector in the direction of the Sun, and ρ is the unit direction vector from the local star to the inter-satellite communication star.

[0058] like Figure 2 As shown in the figure, the coordinate axes of the satellite body are denoted as x and x. b y b and z b The half-cone angle of the inter-satellite communication beam is α, and α < 90°; the unit direction vector from the local satellite to the inter-satellite communication satellite is ρ, and the unit vector in the direction of the Sun is r. S Let η be the angle between the two vectors. Then we have:

[0059] η = arccos(ρ·r) S )

[0060] Without loss of generality, adopt -y b The axis points towards the sun, and the direction of the inter-satellite communication center beam is along the x-axis. b The axial direction, i.e., the x-axis direction, is required. b The axis points to the inter-satellite communication satellite. At a certain moment, -y b If the expected direction of the axis is towards the sun, then x b The axis can only lie in a plane perpendicular to the sun. When x is required... b When the angle between the axis and the direction of the primary star is at its minimum, x will be constrained. b The axis should be located in the plane formed by the direction of the sun and the direction pointing to the inter-satellite communication star.

[0061] Referring to Figure 1 , the embodiment further comprises multiplying the nominal desired attitude M1 by a first correction matrix ΔM2 to obtain a corrected desired attitude M2 (M2=ΔM2·M1) as the current desired attitude of the satellite, when the sun direction angle is greater than 0, and the unit direction vector of the satellite to the inter-satellite communication satellite is not within the coverage of the inter-satellite communication beam and the unit direction vector of the satellite to the inter-satellite communication satellite is not approximately parallel to the unit vector of the sun direction (i.e., η∈(δ,90°-α)∪(90°+α,180°-δ)) in the motion state of the satellite, wherein the function expression of the first correction matrix ΔM2 is:

[0062] ΔM2=R z (ν)

[0063] In the above formula, the function R z (ν) represents a conversion matrix formed by rotating the satellite z b axis by an angle v, and the calculation formula of the angle v is:

[0064]

[0065] In the above formula, η is the included angle between the unit direction vector of the satellite to the inter-satellite communication satellite and the unit vector of the sun direction, α is the half-cone angle of the inter-satellite communication beam, and α<90°, and δ is a preset deviation.

[0066] Referring to Figure 1 , the embodiment further comprises, when the sun direction angle is greater than 0, and the unit direction vector of the satellite to the inter-satellite communication satellite is not within the coverage of the inter-satellite communication beam and the unit direction vector of the satellite to the inter-satellite communication satellite is approximately parallel to the unit vector of the sun direction (i.e., η∈[0,δ]∪[180°-δ,180°]) in the motion state of the satellite, multiplying the nominal desired attitude M1 by a second correction matrix ΔM3 to obtain a corrected desired attitude M3 (M3=ΔM3·M1) as the current desired attitude of the satellite, wherein the function expression of the second correction matrix ΔM3 is: b

[0067] Referring to Figure 1 , the embodiment further comprises, when the sun direction angle is less than or equal to 0 (such as shown in Figure 3 ), and the unit direction vector of the satellite to the inter-satellite communication satellite is within the coverage of the inter-satellite communication beam in the motion state of the satellite, obtaining a flipped nominal desired attitude M′1 by polarity adjustment (to avoid the satellite z b axis pointing to the center of the Earth and pointing to space) as the current desired attitude of the satellite on the basis of the nominal desired attitude M1.

[0068] Referring to Figure 1 , the function expression of the flipped nominal desired attitude M′1 of the embodiment is: ​

[0069]

[0070] In the above formula, M'1 is the flipped nominal desired attitude, x b ,y b and z b are the satellite body coordinate axes, and have:

[0071]

[0072] y b = -r s

[0073]

[0074] In the above formula, r S is the unit vector of the sun direction, and p is the unit direction vector of the home star pointing to the inter-satellite communication star.

[0075] With reference to Figure 1 , the embodiment further includes multiplying the flipped nominal desired attitude M'1 by a second correction matrix AM'2 to obtain a corrected desired attitude M2 (M'2 = AM'2 · M'1) as the current desired attitude of the satellite, when the sun direction angle is less than or equal to 0, and the unit direction vector of the home star pointing to the inter-satellite communication star is not within the coverage of the inter-satellite communication beam, and the unit direction vector of the home star pointing to the inter-satellite communication star is not approximately parallel to the unit vector of the sun direction, simultaneously, wherein the function expression of the second correction matrix AM'2 is:

[0076] AM'2 = R z (ν')

[0077] In the above formula, the function R z (ν') represents a conversion matrix formed by rotating the satellite z b axis by an angle v', and the calculation formula of the angle v' is:

[0078]

[0079] In the above formula, η is the included angle between the unit direction vector of the home star pointing to the inter-satellite communication star and the unit vector of the sun direction, a is the half-cone angle of the inter-satellite communication beam, and a < 90°, and δ is a preset deviation amount.

[0080] With reference to Figure 1 , the embodiment further includes, when the sun direction angle is less than or equal to 0, and the unit direction vector of the home star pointing to the inter-satellite communication star is not within the coverage of the inter-satellite communication beam, and the unit direction vector of the home star pointing to the inter-satellite communication star is approximately parallel to the unit vector of the sun direction, simultaneously, t is avoided due to calculation errors, and the z bThe axis pointing is flipped multiple times in a short period of time to obtain the satellite's long-term sun-facing three-axis pointing attitude at the previous moment as the transitional expected attitude, and this transitional expected attitude is used as the current satellite expected attitude. Based on the x in the nominal expected attitude... b ,y b and z b The function expression, in the inertial frame, z b The axis points to r s ×ρ direction, if ρ and r s If the order of rotation in the plane changes, then z b The axis orientation is reversed. Therefore, when ρ and r s When they are approximately parallel, there is a possibility that the rotation order of the two components may change multiple times within a short period of time. To avoid errors in calculation leading to changes in z... b If the axis direction flips multiple times in a short period of time, the desired posture at the previous moment is taken as the current desired posture.

[0081] In addition, to ensure the stability of the attitude adjustment process, this embodiment also includes calculating the desired angular velocity ω after obtaining the current desired satellite attitude. * The steps are as follows, and when the solar direction angle is greater than 0, the desired angular velocity ω is calculated. * The function expression is:

[0082]

[0083]

[0084]

[0085] Calculate the desired angular velocity ω when the solar direction angle is less than or equal to 0. * The function expression is:

[0086]

[0087]

[0088]

[0089] Where, x b ,y b and z b Let r be the satellite body coordinate axis. S ρ is the unit vector in the direction of the Sun, and ρ is the unit direction vector from the local star to the inter-satellite communication star. For y b The differential, For r S The differential, For z b The differential.

[0090] Based on the above implementation process, the following will illustrate the design simulation examples. In the simulation, the local satellite operates in a circular orbit with an altitude of 500 km and an inclination of 35°; the formation satellites operate in the same orbit; the formation distance between the two satellites is 300 km. The half-cone angle of the inter-satellite communication beam is 70°. The simulation results are as follows: Figure 4 and Figure 5 As shown, by Figure 4 and Figure 5 It can be seen that the designed attitude can meet the requirements for sun-pointing and inter-satellite communication pointing. The three-axis sun-pointing attitude design method for long-term satellite sun-pointing under inter-satellite communication constraints in this embodiment provides a three-axis stable sun-pointing attitude design method that not only ensures the satellite's energy security attitude pointing conditions, but also meets the continuity requirements of formation communication, while ensuring stable sun exposure conditions for the satellite.

[0091] It should be noted that the satellite long-term solar-orientation three-axis pointing attitude design method under inter-satellite communication constraints in this embodiment, while having multiple constraints, results in various current desired satellite attitudes that are coupled in some way, but are independent of each other.

[0092] Furthermore, this embodiment also provides a satellite long-term solar orientation three-axis pointing attitude design system under inter-satellite communication constraints, including a microprocessor and a memory interconnected thereto. The microprocessor is programmed or configured to execute the steps of the aforementioned satellite long-term solar orientation three-axis pointing attitude design method under inter-satellite communication constraints, or the memory stores a computer program programmed or configured to execute the aforementioned satellite long-term solar orientation three-axis pointing attitude design method under inter-satellite communication constraints.

[0093] In addition, this embodiment also provides a computer-readable storage medium storing a computer program that is programmed or configured to execute the aforementioned satellite long-term solar three-axis pointing attitude design method under inter-satellite communication constraints.

[0094] In summary, to meet the requirements of the satellite in-orbit three-axis stabilization sun pointing and formation inter-satellite communication pointing tasks, the satellite long-term sun pointing three-axis pointing attitude design method under the inter-satellite communication constraint of the embodiment includes establishing a satellite three-axis stabilization sun pointing control condition, and giving an expected attitude to ensure that the solar sail points to the solar vector; On this basis, the pointing condition of satisfying the inter-satellite communication beam coverage is added; Secondly, for low-inclination orbit satellites, due to the existence of the process of positive and negative changes of the sunlight angle, an attitude adjustment strategy is designed to avoid the dramatic change of the load illumination condition; In order to ensure the stability of the attitude adjustment process, an expected angular velocity design method is given. The satellite long-term sun pointing three-axis pointing attitude design method under the inter-satellite communication constraint of the embodiment not only can ensure the safe attitude pointing condition of the satellite energy, but also can meet the continuity requirement of the formation communication, and can also ensure the stable sunning condition of the satellite.

[0095] The above only describes the preferred embodiments of the present application, and it should be noted that for ordinary skilled persons in the art, some improvements and refinements can be made without departing from the principles of the present application, and these improvements and refinements should be considered as the protection scope of the present application.

Claims

1. A method for designing a satellite's long-term solar-orientation three-axis pointing attitude under inter-satellite communication constraints, characterized in that, This includes acquiring the solar azimuth angle and satellite motion state. When the solar azimuth angle is greater than 0, and the unit direction vector from the satellite to the inter-satellite communication satellite is simultaneously within the coverage area of ​​the inter-satellite communication beam, a nominal desired attitude M1 is designed as the current desired satellite attitude. The functional expression for the nominal desired attitude M1 is: In the above formula, M1 is the nominal desired attitude, x b ,y b and z b Let be the coordinate representation of the three direction vectors of the satellite system coordinate axes in the inertial coordinate system, and we have: y b ＝-r s In the above formula, r S ρ is the unit vector in the direction of the Sun, and ρ is the unit direction vector from the local star to the inter-satellite communication star.

2. The method for designing a satellite's long-term solar-orientation three-axis pointing attitude under inter-satellite communication constraints according to claim 1, characterized in that, This also includes situations where, simultaneously, the unit direction vector from the satellite to the inter-satellite communication satellite is greater than 0, the unit direction vector from the satellite to the inter-satellite communication satellite is not within the coverage area of ​​the inter-satellite communication beam, and the unit direction vector from the satellite to the inter-satellite communication satellite is not parallel to the unit vector of the solar direction, then, the corrected desired attitude M2 ​​is obtained by multiplying the nominal desired attitude M1 by the first correction matrix ΔM2 as the current desired satellite attitude. The functional expression of the first correction matrix ΔM2 is as follows: ΔM2=R z (n) In the above formula, the function R z (ν) represents orbiting the satellite z b The transformation matrix formed by the rotation angle ν, and the formula for calculating the angle ν is: In the above formula, η is the angle between the unit direction vector of the local satellite pointing to the inter-satellite communication satellite and the unit vector of the solar direction, α is the half-cone angle of the inter-satellite communication beam, and α < 90°, and δ is the preset deviation.

3. The method for designing a satellite's long-term solar-orientation three-axis pointing attitude under inter-satellite communication constraints according to claim 2, characterized in that, It also includes obtaining the satellite's long-term solar orientation attitude at the previous moment as a transitional expected attitude when the solar orientation angle is greater than 0, the unit direction vector of the satellite pointing to the inter-satellite communication satellite is not within the coverage area of ​​the inter-satellite communication beam, and the unit direction vector of the satellite pointing to the inter-satellite communication satellite is parallel to the unit vector of the solar orientation. The transitional expected attitude is then used as the current expected attitude of the satellite.

4. The method for designing a satellite's long-term solar-orbiting three-axis pointing attitude under inter-satellite communication constraints according to claim 3, characterized in that, It also includes the following: when the solar orientation angle is less than or equal to 0 and the unit direction vector of the satellite pointing to the inter-satellite communication satellite is simultaneously established within the coverage area of ​​the inter-satellite communication beam, the nominal desired attitude M′1 is obtained by polarity adjustment based on the nominal desired attitude M1 and is used as the current satellite desired attitude.

5. The method for designing a satellite's long-term solar-orientation three-axis pointing attitude under inter-satellite communication constraints according to claim 4, characterized in that, The functional expression for flipping the nominal desired attitude M′1 is: In the above formula, M′1 represents the nominal desired orientation for flipping, and x b ,y b and z b Let be the coordinate axes of the satellite body, and we have: y b ＝-r s In the above formula, r S ρ is the unit vector in the direction of the Sun, and ρ is the unit direction vector from the local star to the inter-satellite communication star.

6. The method for designing a satellite's long-term solar-orientation three-axis pointing attitude under inter-satellite communication constraints according to claim 5, characterized in that, This also includes situations where, simultaneously, the solar orientation angle is less than or equal to 0, and the unit orientation vector from the satellite to the inter-satellite communication satellite is not within the coverage area of ​​the inter-satellite communication beam, and the unit orientation vector from the satellite to the inter-satellite communication satellite is not parallel to the unit orientation vector of the solar direction, then, by multiplying the left side of the flipped nominal desired attitude M′1 by the second correction matrix ΔM′2, the corrected desired attitude M′2 is obtained as the current desired satellite attitude. The functional expression of the second correction matrix ΔM′2 is: ΔM′2=R z (n′) In the above formula, the function R z (ν′) represents orbiting the satellite z b The transformation matrix formed by the rotation angle ν′, and the formula for calculating the angle ν′ is: In the above formula, η is the angle between the unit direction vector of the local satellite pointing to the inter-satellite communication satellite and the unit vector of the solar direction, α is the half-cone angle of the inter-satellite communication beam, and α < 90°, and δ is the preset deviation.

7. The method for designing a satellite's long-term solar-orientation three-axis pointing attitude under inter-satellite communication constraints according to claim 6, characterized in that, It also includes obtaining the satellite's long-term solar orientation attitude at the previous moment as a transitional expected attitude when the solar orientation angle is less than or equal to 0, the unit orientation vector of the satellite pointing to the inter-satellite communication satellite in the satellite's motion state is not within the coverage of the inter-satellite communication beam, and the unit orientation vector of the satellite pointing to the inter-satellite communication satellite is parallel to the unit orientation vector of the sun. The transitional expected attitude is then used as the current expected attitude of the satellite.

8. The method for designing a satellite's long-term solar-orbiting three-axis pointing attitude under inter-satellite communication constraints according to any one of claims 1 to 7, characterized in that, After obtaining the current desired satellite attitude, the calculation of the desired angular velocity ω is also included. * The steps are as follows, and when the solar direction angle is greater than 0, the desired angular velocity ω is calculated. * The function expression is: Calculate the desired angular velocity ω when the solar direction angle is less than or equal to 0. * The function expression is: Where, x b ,y b and z b Let r be the satellite body coordinate axis. S ρ is the unit vector in the direction of the Sun, and ρ is the unit direction vector from the local star to the inter-satellite communication star. For y b The differential, For r S The differential, For z b The differential.

9. A satellite long-term solar-orbiting three-axis pointing attitude design system under inter-satellite communication constraints, comprising a microprocessor and a memory interconnected, characterized in that, The microprocessor is programmed or configured to execute the long-term solar orientation three-axis pointing attitude design method for a satellite under inter-satellite communication constraints as described in any one of claims 1 to 8, or the memory stores a computer program programmed or configured to execute the long-term solar orientation three-axis pointing attitude design method for a satellite under inter-satellite communication constraints as described in any one of claims 1 to 8.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that is programmed or configured to execute the long-term solar-orbiting three-axis pointing attitude design method for a satellite under inter-satellite communication constraints as described in any one of claims 1 to 8.

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