Non-sun-synchronous orbit satellite energy optimal attitude sailboard control method and system

Abstract

The invention provides a non-sun-synchronous orbit satellite energy optimal attitude sailboard control method and system. The method comprises the steps that S1, the illumination angle of each orbit in the satellite life period is calculated according to collected orbit parameters; s2, calculating the illumination area half angle and the illumination duration under the illumination angle of each track; s3, calculating the illumination energy of each track of the sailboard driving shaft under the combination of different swinging positions and different flight attitudes of the satellite; s4, comparing illumination energy of each track, and preferably selecting a combination of a posture with the maximum energy in each period and a sailboard driving position; and S5, obtaining a control strategy for the attitude of all orbital circles of satellites and the position of a sailboard driving shaft in the whole life period of the satellites. By comparing various feasible in-orbit attitude and sailboard position combinations, a sailboard position and attitude control method with optimal energy can be obtained through optimization, the in-orbit receiving energy of a two-dimensional driving solar sailboard satellite is improved, the required sailboard area is reduced under the same power consumption requirement, and the weight of the whole satellite is reduced.

Description

Technical Field

[0001] The present invention relates to the field of satellite attitude control and energy management technology, specifically, to a method for energy-optimal attitude sailboard coupling control of a non-sun-synchronous orbit satellite, and more particularly to an attitude and sailboard swinging strategy for different time periods during the life of a non-sun-synchronous orbit satellite. Background Art

[0002] Because the on-orbit illumination environment for satellites in non-sun-synchronous orbits is more complex than that of satellites in sun-synchronous orbits, an increasing number of satellites are using two-dimensionally driven solar panels to better capture sunlight energy. This requires a sophisticated solar panel actuation strategy to optimize the satellite's on-orbit illumination energy. Due to mission timeliness constraints, a satellite's long-term operational attitude (when not performing a mission) cannot be directed toward the Sun. Optimizing solar panel actuation alone fails to consider the satellite's various possible operational attitudes. Therefore, a combination of solar panel actuation and attitude control is needed to optimize the satellite's on-orbit illumination energy.

[0003] Existing research on the control of satellite on-orbit sailboards mainly focuses on the solar orientation control of satellite one-dimensional driven sailboards, the Xb / Yb swing switching control of two-dimensional driven sailboards, and the uniform solar control of the rotation axis. However, no research has been conducted on the variable speed solar control of two-dimensional driven sailboards, and the problem of combining the satellite's operating attitude and the two-dimensional control of the sailboard to optimize the light energy.

[0004] Patent document CN105035364A discloses a method for driving the solar array oscillation of a low-inclination orbit radar satellite. This method analyzes the illumination conditions when the solar array's rotation axis is located at different inclination positions on the Yb axis of the satellite and when the satellite is in different attitudes. Based on the orbital illumination angle, the radar satellite's solar array drive oscillation state is appropriately set to ensure an illumination angle greater than 50° throughout the satellite's lifespan. However, this method does not comprehensively consider the situation where the solar array's rotation axis swings toward the Xb axis of the satellite, and the designed sailboard control strategy fails to achieve energy optimization.

[0005] Patent document CN102004492A discloses a method for controlling dual-axis solar panels on non-sun-synchronous orbit satellites. The method describes a method for controlling dual-axis solar panels facing the Sun, taking into account potential failure scenarios and designing a complete implementation process. This method can achieve solar-facing solar panels without attitude maneuvers. However, this method requires the satellite's solar panels to be drivable in any two dimensions, which causes periodic changes in the satellite's inertia, reducing attitude control accuracy. Furthermore, the panels obstruct a large area, affecting the field of view layout of various satellite antennas.

[0006] Patent document CN104090612A discloses a method for obtaining energy for spacecraft in inclined orbits based on yaw guidance. This method describes a comprehensive control strategy using yaw attitude guidance and one-dimensional actuation of solar panels. This strategy ensures that the normal of the solar panels points toward the sun, ensuring that the spacecraft receives solar energy while in an inclined orbit. However, at high illumination angles, this method requires a large yaw attitude angle, which is detrimental to the satellite's ability to quickly respond to missions.

[0007] Patent document CN104181941A discloses a bidirectional control method for sailboards used in inclined orbits. This method uses yaw control and sailboard rotation to achieve solar orientation, adapting to both forward and inverted flight scenarios. However, this method only considers one-dimensional sailboard drive and is not optimal for two-dimensional sailboard drive.

[0008] Patent document CN106096148A describes a method for pointing solar panels on high-inclination satellites using simple attitude control. This method uses yaw to switch the panel drive axis between the Xo and Yo axes of the orbital system to achieve a better lighting environment. Compared to control methods where the panel is positioned solely on the Yo axis, this method improves lighting conditions at angles greater than 45°. However, this method does not consider the panel's variable speed solar orientation, resulting in suboptimal energy per orbit. Furthermore, it does not account for illumination issues associated with side-view attitudes and panel tilt offsets. Summary of the Invention

[0009] In view of the defects in the prior art, the purpose of the present invention is to provide a method and system for controlling an energy-optimal attitude sailboard of a non-sun-synchronous orbit satellite.

[0010] According to the present invention, a method for controlling an energy-optimal attitude sailboard of a non-sun-synchronous orbit satellite comprises:

[0011] Step S1: Calculate the illumination angle of each orbit during the satellite life cycle based on the collected orbital parameters;

[0012] Step S2: Calculate the half angle of the illumination area and the illumination duration under the illumination angle of each track;

[0013] Step S3: Calculating the illumination energy per track under different swing positions of the sailboard drive shaft and different satellite flight attitudes according to the illumination area half angle and illumination duration;

[0014] Step S4: comparing the illumination energy of each track, and selecting the posture and sailboard driving position combination with the maximum energy in each cycle;

[0015] Step S5: According to the optimization result, the satellite attitude and sailboard drive axis position control strategy for all orbits during the entire life of the satellite are obtained.

[0016] Preferably, step S1 includes the following sub-steps:

[0017] Step S1.1: Based on the solar position formula and the given sidereal time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system. s (t);

[0018] Step S1.2: Calculate the orbital inclination i and the right ascension Ω of the ascending node at a given stellar time t by orbit recursion in the J2000.0 coordinate system, and obtain the orbital angular momentum coordinate h o (t);

[0019] Step S1.3: According to the sun vector R s (t) and orbital angular momentum h o The illumination angle β(t) value of each orbit of the satellite during its lifespan is obtained through recursive iterative calculation.

[0020] Preferably, step S2 includes the following sub-steps:

[0021] Step S2.1: Calculate the illumination phase half angle θ for each track based on the track semi-major axis a and the illumination angle β(t) S (t):

[0022]

[0023] Among them, R e represents the radius of the Earth;

[0024] Step S2.2: When the illumination angle is β(t), calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period. s (θ):

[0025]

[0026] Here, θ represents the phase angle of the satellite in orbit.

[0027] Preferably, step S3 includes the following sub-steps:

[0028] Step S3.1: Calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ according to the swing position of the sailboard drive shaft and the flight attitude of the satellite. p (θ):

[0029] n p (θ)=n s (θ)-(n s (θ)·a p )a p

[0030] Among them, a p represents the coordinates of the sailboard drive axis vector in the orbital system;

[0031] Step S3.2: Calculate the illumination energy E(t) of each track:

[0032]

[0033] Where a is the semi-major axis of the satellite orbit, μ = 3.9860044 × 10 5 is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per track and the illumination time.

[0034] Preferably, the calculation formula for the illumination angle β(t) value of each track is:

[0035]

[0036] According to the present invention, an energy-optimal attitude sailboard control system for a non-sun-synchronous orbit satellite is provided, comprising:

[0037] Module M1: Calculates the illumination angle of each orbit during the satellite's life cycle based on the collected orbital parameters;

[0038] Module M2: Calculates the half angle of the illumination area and the illumination duration under each track illumination angle;

[0039] Module M3: Calculates the illumination energy per track under different swing positions of the sailboard drive shaft and different satellite flight attitudes based on the illumination area half-angle and illumination duration;

[0040] Module M4: Compare the illumination energy of each track and select the posture and sailboard driving position combination with the maximum energy in each cycle;

[0041] Module M5: Based on the optimization results, obtain the satellite attitude and sailboard drive axis position control strategy for all orbits during the entire life of the satellite.

[0042] Preferably, the module M1 includes the following submodules:

[0043] Module M1.1: Based on the solar position formula and the given sidereal time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system s (t);

[0044] Module M1.2: By orbit recursion in the J2000.0 coordinate system, calculate the orbital inclination i and the right ascension Ω of the ascending node at a given star time t, and obtain the orbital angular momentum coordinate h o (t);

[0045] Module M1.3: According to the solar vector R s (t) and orbital angular momentum h o The illumination angle β(t) value of each orbit of the satellite during its lifespan is obtained through recursive iterative calculation.

[0046] Preferably, the module M2 includes the following submodules:

[0047] Module M2.1: Calculate the illumination phase half angle θ for each track based on the track semi-major axis a and the illumination angle β(t) S (t):

[0048]

[0049] Among them, R e represents the radius of the Earth;

[0050] Module M2.2: Calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period when the illumination angle is β(t). s (θ):

[0051]

[0052] Here, θ represents the phase angle of the satellite in orbit.

[0053] Preferably, the module M3 includes the following submodules:

[0054] Module M3.1: Calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ based on the swing position of the sailboard drive shaft and the flight attitude of the satellite. p (θ):

[0055] n p (θ)=n s (θ)-(n s (θ)·a p )a p

[0056] Among them, a p represents the coordinates of the sailboard drive axis vector in the orbital system;

[0057] Module M3.2: Calculate the illumination energy E(t) per track:

[0058]

[0059] Where a is the semi-major axis of the satellite orbit, μ = 3.9860044 × 10 5 is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per track and the illumination time.

[0060] Preferably, the calculation formula for the illumination angle β(t) value of each track is:

[0061]

[0062] Compared with the prior art, the present invention has the following beneficial effects:

[0063] 1. By comparing multiple feasible on-orbit attitudes and panel position combinations, the present invention can optimize the panel position and attitude control method with the best energy, improve the on-orbit received energy of satellites using two-dimensional driven solar panels, reduce the required panel area under the same power consumption requirements, and reduce the weight of the entire satellite.

[0064] 2. The present invention can meet the needs of level flight, side-view flight and sailboard tilt and offset satellites through the side-view and offset angle settings; by comparing multiple side-view postures and sailboard rotation axis swing position combinations, the energy-optimal control strategy can be obtained, reducing the required sailboard area under the same power consumption requirements.

[0065] 3. The present invention is simple in calculation, has universality, can overcome the defects of the prior art, and has good practicality.

[0066] Other beneficial effects of the present invention will be explained through the introduction of specific technical features and technical solutions in the specific implementation methods. Those skilled in the art should be able to understand the beneficial technical effects brought about by the introduction of these technical features and technical solutions. BRIEF DESCRIPTION OF THE DRAWINGS

[0067] Other features, objects and advantages of the present invention will become more apparent upon reading the detailed description of non-limiting embodiments with reference to the following drawings:

[0068] Figure 1 This is a specific flow chart of the invention for designing an optimal control strategy for satellite on-orbit illumination energy.

[0069] Figure 2 This is a schematic diagram used by the present invention to describe the satellite orbit coordinate system and this system.

[0070] Figure 3 This is a curve diagram of the illumination angle of the satellite in one year in an embodiment of the present invention.

[0071] Figure 4 This is a curve diagram of the satellite's illumination time in one year in an embodiment of the present invention.

[0072] Figure 5 Graphs of illumination energy of a satellite when its sailboard swings to different positions in an embodiment of the present invention.

[0073] Figure 6 Schematic diagram of the swing position and operating posture of the sailboard in one year under the condition of optimal satellite energy in an embodiment of the present invention. DETAILED DESCRIPTION

[0074] The present invention will be described in detail below with reference to specific embodiments. The following examples will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those skilled in the art, several changes and improvements can be made without departing from the scope of the present invention. These all fall within the scope of protection of the present invention.

[0075] Reference Figure 1 and Figure 2 As shown, a method for controlling the energy-optimal attitude sailboard coupling of a non-sun-synchronous orbit satellite comprises:

[0076] Step S1: According to the solar position formula and the given stellar time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system s (t);

[0077] Step S2: Calculate the orbital inclination i and the right ascension Ω of the ascending node at a given stellar time t by orbit recursion in the J2000.0 coordinate system, and obtain the orbital angular momentum coordinate h o (t);

[0078] Step S3: According to the sun vector R s (t) and orbital angular momentum h o (t) value, the orbital illumination angle β(t) value of each orbit of the satellite during its life cycle is obtained by recursive iterative calculation, and its value satisfies:

[0079]

[0080] Step S4: Calculate the illumination phase half angle θ of each track based on the track semi-major axis a and the illumination angle β(t) S (t), which is half the phase angle corresponding to the illumination area of each track, is as follows:

[0081]

[0082] Among them, R e = represents the radius of the Earth. Since the illumination angle β(t) changes only slightly within one orbital period, it is assumed to be constant within one orbital period.

[0083] Step S5: When the illumination angle is β(t), calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period. s (θ), satisfying:

[0084]

[0085] Where θ represents the phase angle of the satellite in orbit. The moment when phase θ = 0 is taken as the center moment of the orbit illumination area, then the corresponding phase range of the illumination area is: (-θ S (t),θS (t)).

[0086] Step S6: Calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ according to the swing position of the sailboard drive shaft (Xb axis or Yb axis in this system) and the satellite's optional long-term flight attitude (level flight, left view, right view, etc.). p (θ), satisfying:

[0087] n p (θ)=n s (θ)-(n s (θ)·a p )a p

[0088] Among them, a p Represents the coordinates of the sailboard drive axis vector in the orbital system.

[0089] Step S7: Calculate the illumination energy E(t) of each track:

[0090]

[0091] Where a is the semi-major axis of the satellite orbit, μ = 3.9860044 × 10 5 is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per track and the illumination time.

[0092] Step S8: Comprehensively compare the illumination energy under the combination of the two swing positions of the sailboard drive shaft and different flight attitudes of the satellite in each orbital period (such as: X-axis + left view, X-axis + right view, Y-axis + left view, Y-axis + right view, etc.), and select the attitude and sailboard drive position combination with the maximum energy in each period.

[0093] Step S9: The optimal attitude and sailboard drive combination in each cycle is synthesized to obtain the satellite attitude and sailboard drive axis position control strategy for all orbits during the entire life of the satellite.

[0094] By comparing multiple feasible on-orbit attitudes and panel position combinations, the present invention can optimize the panel position and attitude control method with the best energy, improve the on-orbit received energy of satellites using two-dimensional driven solar panels, reduce the required panel area under the same power consumption requirements, and reduce the weight of the entire satellite.

[0095] For satellites in non-sun-synchronous orbits, the illumination angle of the orbital plane varies within a large range of positive and negative angles. How to control the satellite attitude and two-dimensionally driven solar panels to optimize satellite energy is a problem. The present invention discloses a combination of two swing positions of two-dimensionally driven solar panels (the Xb axis and Yb axis of this system) and multiple satellite operating attitudes. The variable speed exposure to sunlight energy under different combinations is calculated according to the derivation formula, and the optimal panel and attitude drive strategy within the service life is optimized by comparing multiple combination schemes. The method is simple to calculate and practical, and is applicable to various types of satellites operating in non-sun-synchronous orbits. For the control strategy of the panel at a uniform speed facing the sun, the same optimization idea can be used to obtain the optimal control scheme.

[0096] In order to achieve the above object, the present invention is implemented by the following technical solution, which specifically includes the following steps:

[0097] Step 1: Based on the solar position formula and the given stellar time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system. s (t);

[0098] Step 2: Calculate the orbital inclination i and the right ascension Ω of the ascending node at a given stellar time t by orbit recursion in the J2000.0 coordinate system, and obtain the orbital angular momentum coordinate h o (t).

[0099] Step 3: According to the sun vector R s (t) and orbital angular momentum h o (t) value, calculate the orbit illumination angle β(t), whose value satisfies:

[0100]

[0101] Through recursive iterative calculation, the orbital illumination angle β(t) value of each orbit of the satellite during its life cycle is obtained.

[0102] Step 4: Based on the track semi-major axis a and the illumination angle β(t), the illumination phase half angle θ of each track can be calculated. S (t), which is half the phase angle corresponding to the illumination area of each track, is as follows:

[0103]

[0104] Among them, R e = represents the radius of the Earth. Since the illumination angle β(t) changes only slightly within one orbital period, it is assumed to be constant within one orbital period.

[0105] Step 5: When the illumination angle is β(t), calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period. s (θ), satisfying:

[0106]

[0107] Where θ represents the phase angle of the satellite in orbit. The moment when phase θ = 0 is taken as the center moment of the orbit illumination area, then the corresponding phase range of the illumination area is: (-θ S (t),θ S (t)).

[0108] Step 6: Calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ based on the swing position of the sailboard drive shaft (Xb axis or Yb axis in this system) and the satellite's optional long-term flight attitude (level flight, left view, right view, etc.). p (θ), satisfying:

[0109] n p (θ)=n s (θ)-(n s (θ)·a p )a p

[0110] Among them, a p Represents the coordinates of the sailboard drive axis vector in the orbital system.

[0111] Step 7. Calculate the illumination energy E(t) of each track:

[0112]

[0113] Where a is the semi-major axis of the satellite orbit, μ = 3.9860044 × 10 5 is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per track and the illumination time.

[0114] Step 8: Comprehensively compare the illumination energy under the combinations of two swing positions of the sailboard drive shaft and different flight attitudes of the satellite in each orbital period, and select the attitude and sailboard drive position combination with the maximum energy.

[0115] Step 9: Combining the optimal attitude and sailboard drive combination in each cycle, the satellite attitude and sailboard drive position curves for all orbits in the entire life of the satellite are obtained.

[0116] Specifically, in one embodiment, a method for energy-optimal attitude sailboard coupling control for a non-sun-synchronous orbit satellite is further described in detail. This method includes orbit illumination angle calculation, illumination time calculation, illumination energy calculation, and optimal control strategy design. For example, a radar satellite operating at an altitude of 600 km, in a 50° circular orbit with a side-view attitude angle of 30° is used as an example, with a time period of one year. The specific steps are as follows:

[0117] 1. Based on the given star time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system s (t);

[0118] 2. Calculate the orbital inclination i and the right ascension Ω of the ascending node at a given star time in the J2000.0 coordinate system through orbital recursion, and calculate the orbital angular momentum coordinate h in the J2000.0 geocentric equatorial inertial coordinate system. o (t);

[0119] 3. According to the sun vector R s (t) and orbital angular momentum h o (t) value, calculate the orbit illumination angle β(t), whose value satisfies:

[0120]

[0121] like Figure 3 As shown, through recursive iterative calculation, the orbital illumination angle β(t) value of each orbit of the satellite during its life cycle is obtained.

[0122] 4. According to the track semi-major axis a and the illumination angle β(t), the illumination phase half angle θ of each track can be calculated S (t), which is half the phase angle corresponding to the illumination area of each track, is as follows:

[0123]

[0124] Among them, R e = represents the radius of the Earth. Since the illumination angle β(t) changes only slightly within one orbital period, it is assumed to be constant within one orbital period.

[0125] like Figure 4 As shown, the illumination time curve of the selected case satellite in one year can be obtained by calculation.

[0126] 5. When the illumination angle is β(t), calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period s (θ), satisfying:

[0127]

[0128] Where θ represents the phase angle of the satellite in orbit. The moment when phase θ = 0 is taken as the center moment of the orbit illumination area, then the corresponding phase range of the illumination area is: (-θ S (t),θ S (t)).

[0129] 6. According to the swing position of the sailboard drive axis (Xb axis or Yb axis of this system) and the optional long-term flight attitude of the satellite (level flight, left view, right view, etc.), calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ p (θ), satisfying:

[0130] n p (θ)=n s (θ)-(n s (θ)·a p )a p

[0131] Among them, a p Represents the coordinates of the sailboard drive axis vector in the orbital system.

[0132] 7. Calculate the illumination energy E(t) of each track:

[0133]

[0134] Where a is the semi-major axis of the satellite orbit, μ = 3.9860044 × 10 5 is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per track and the illumination time.

[0135] like Figure 5 As shown, the illumination energy of the selected case satellite can be calculated in three cases: the sailboard swings to the Xb axis, the sailboard swings to the left side of the Yb axis, and the sailboard swings to the right side of the Yb axis.

[0136] 8. Comprehensively compare the light energy under the two swing positions of the sailboard drive shaft and the different flight attitudes of the satellite in each orbital period, and select the attitude and sailboard drive position combination with the maximum energy.

[0137] 9. Combining the optimal attitude and sailboard drive combination in each cycle, the satellite attitude and sailboard drive position curves for all orbits during the entire life of the satellite are obtained.

[0138] like Figure 6 As shown, the swing position and running attitude of the sailboard within one year under the optimal satellite energy condition of the selected case can be calculated; in the figure, the curve value "1" represents that the sailboard is located on the Xb axis, the curve value "2" represents that the sailboard is located on the Yb axis and the attitude is kept on the left side, and the curve value "3" represents that the sailboard is located on the Yb axis and the attitude is kept on the right side.

[0139] The present invention also provides a non-sun-synchronous orbit satellite energy-optimal attitude sailboard control system, which can be implemented by executing the process steps of the non-sun-synchronous orbit satellite energy-optimal attitude sailboard control method, that is, those skilled in the art can understand the non-sun-synchronous orbit satellite energy-optimal attitude sailboard control method as a preferred implementation of the non-sun-synchronous orbit satellite energy-optimal attitude sailboard control system.

[0140] Specifically, an energy-optimal attitude sailboard control system for a non-sun-synchronous orbit satellite includes:

[0141] Module M1: Calculates the illumination angle of each orbit during the satellite's life cycle based on the collected orbital parameters;

[0142] Module M2: Calculates the half angle of the illumination area and the illumination duration under each track illumination angle;

[0143] Module M3: Calculates the illumination energy per track under different swing positions of the sailboard drive shaft and different satellite flight attitudes based on the illumination area half-angle and illumination duration;

[0144] Module M4: Compare the illumination energy of each track and select the posture and sailboard driving position combination with the maximum energy in each cycle;

[0145] Module M5: Based on the optimization results, obtain the satellite attitude and sailboard drive axis position control strategy for all orbits during the entire life of the satellite.

[0146] The module M1 includes the following submodules:

[0147] Module M1.1: Based on the solar position formula and the given sidereal time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system s (t);

[0148] Module M1.2: By orbit recursion in the J2000.0 coordinate system, calculate the orbital inclination i and the right ascension Ω of the ascending node at a given star time t, and obtain the orbital angular momentum coordinate h o (t);

[0149] Module M1.3: According to the solar vector R s (t) and orbital angular momentum h o The illumination angle β(t) value of each orbit of the satellite during its lifespan is obtained through recursive iterative calculation.

[0150] The module M2 includes the following submodules:

[0151] Module M2.1: Calculate the illumination phase half angle θ for each track based on the track semi-major axis a and the illumination angle β(t) S (t):

[0152]

[0153] Among them, R e represents the radius of the Earth;

[0154] Module M2.2: Calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period when the illumination angle is β(t). s (θ):

[0155]

[0156] Here, θ represents the phase angle of the satellite in orbit.

[0157] The module M3 includes the following submodules:

[0158] Module M3.1: Calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ based on the swing position of the sailboard drive shaft and the flight attitude of the satellite. p (θ):

[0159] n p (θ)=n s (θ)-(n s (θ)·a p )a p

[0160] Among them, a p represents the coordinates of the sailboard drive axis vector in the orbital system;

[0161] Module M3.2: Calculate the illumination energy E(t) per track:

[0162]

[0163] Where a is the semi-major axis of the satellite orbit, μ = 3.9860044 × 10 5 is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per track and the illumination time.

[0164] The calculation formula for the illumination angle β(t) of each track is:

[0165]

[0166] Those skilled in the art will appreciate that, in addition to implementing the system and its various devices, modules, and units provided by the present invention in purely computer-readable program code, it is entirely possible to implement the same functions of the system and its various devices, modules, and units provided by the present invention in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system and its various devices, modules, and units provided by the present invention can be considered a hardware component, and the devices, modules, and units included therein for implementing various functions can also be considered as structures within the hardware component; the devices, modules, and units for implementing various functions can also be considered as both software modules implementing the method and structures within the hardware component.

[0167] The above describes specific embodiments of the present invention. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art may make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. The embodiments of this application and the features in the embodiments may be combined with each other in any manner unless there is a conflict.

Claims

1. A method for controlling an energy-optimal attitude sailboard of a non-sun-synchronous orbit satellite, characterized in that: include: Step S1: Calculate the illumination angle of each orbit during the satellite life cycle based on the collected orbital parameters; Step S2: Calculate the half angle of the illumination area and the illumination duration under the illumination angle of each track; Step S3: Calculating the illumination energy per track under different swing positions of the sailboard drive shaft and different satellite flight attitudes according to the illumination area half angle and illumination duration; Step S4: comparing the illumination energy of each track, and selecting the posture and sailboard driving position combination with the maximum energy in each cycle; Step S5: According to the optimization result, the satellite attitude and sailboard drive axis position control strategy for all orbits during the entire life of the satellite are obtained.

2. The energy-optimal attitude sailboard control method for a non-sun-synchronous orbit satellite according to claim 1, characterized in that: The step S1 includes the following sub-steps: Step S1.1: Based on the solar position formula and the given sidereal time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system. s (t); Step S1.2: Calculate the orbital inclination i and the right ascension Ω of the ascending node at a given stellar time t by orbit recursion in the J2000.0 coordinate system, and obtain the orbital angular momentum coordinate h o (t); Step S1.3: According to the sun vector R s (t) and orbital angular momentum h o The illumination angle β(t) value of each orbit of the satellite during its lifespan is obtained through recursive iterative calculation.

3. The energy-optimal attitude sailboard control method for a non-sun-synchronous orbit satellite according to claim 2, characterized in that: The step S2 includes the following sub-steps: Step S2.1: Calculate the illumination phase half angle θ for each track based on the track semi-major axis a and the illumination angle β(t) S (t): Among them, R e represents the radius of the Earth; Step S2.2: When the illumination angle is β(t), calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period. s (θ): Here, θ represents the phase angle of the satellite in orbit.

4. The energy-optimal attitude sailboard control method for a non-sun-synchronous orbit satellite according to claim 3, characterized in that: The step S3 includes the following sub-steps: Step S3.1: Calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ according to the swing position of the sailboard drive shaft and the flight attitude of the satellite. p (θ): n p (θ)=n s (θ)-(n s (i)·a p )a p Among them, a p represents the coordinates of the sailboard drive axis vector in the orbital system; Step S3.2: Calculate the illumination energy E(t) of each track: Where a is the semi-major axis of the satellite orbit, μ is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per orbit and the illumination time.

5. The energy-optimal attitude sailboard control method for a non-sun-synchronous orbit satellite according to claim 2, characterized in that: The calculation formula for the illumination angle β(t) of each track is:

6. An energy-optimal attitude sailboard control system for a non-sun-synchronous orbit satellite, characterized in that: include: Module M1: Calculates the illumination angle of each orbit during the satellite's life cycle based on the collected orbital parameters; Module M2: Calculates the half angle of the illumination area and the illumination duration under each track illumination angle; Module M3: Calculates the illumination energy per track under different swing positions of the sailboard drive shaft and different satellite flight attitudes based on the illumination area half-angle and illumination duration; Module M4: Compare the illumination energy of each track and select the posture and sailboard driving position combination with the maximum energy in each cycle; Module M5: Based on the optimization results, obtain the satellite attitude and sailboard drive axis position control strategy for all orbits during the entire life of the satellite.

7. The energy-optimal attitude sailboard control system for a non-sun-synchronous orbit satellite according to claim 6, characterized in that: The module M1 Includes the following submodules: Module M1.1: Based on the solar position formula and the given sidereal time t, recursively calculate the coordinate R of the solar vector in the J2000.0 coordinate system s (t); Module M1.2: By orbit recursion in the J2000.0 coordinate system, calculate the orbital inclination i and the right ascension Ω of the ascending node at a given star time t, and obtain the orbital angular momentum coordinate h o (t); Module M1.3: According to the solar vector R s (t) and orbital angular momentum h o The illumination angle β(t) value of each orbit of the satellite during its lifespan is obtained through recursive iterative calculation.

8. The energy-optimal attitude sailboard control system for a non-sun-synchronous orbit satellite according to claim 7, characterized in that: The module M2 Includes the following submodules: Module M2.1: Calculate the illumination phase half angle θ for each track based on the track semi-major axis a and the illumination angle β(t) S (t): Among them, R e represents the radius of the Earth; Module M2.2: Calculate the coordinate n of the sun direction vector in the orbital coordinate system within one orbital period when the illumination angle is β(t). s (θ): Here, θ represents the phase angle of the satellite in orbit.

9. The energy-optimal attitude sailboard control system for a non-sun-synchronous orbit satellite according to claim 8, characterized in that: The module M3 includes the following submodules: Module M3.1: Calculate the coordinate n of the sailboard normal vector in the orbital coordinate system at different phases θ based on the swing position of the sailboard drive shaft and the flight attitude of the satellite. p (θ): n p (θ)=n s (θ)-(n s (i)·a p )a p Among them, a p represents the coordinates of the sailboard drive axis vector in the orbital system; Module M3.2: Calculate the illumination energy E(t) per track: Where a is the semi-major axis of the satellite orbit, μ is the gravitational constant, and the illumination energy E(t) is the product of the average illumination efficiency per orbit and the illumination time.

10. The energy-optimal attitude sailboard control system for a non-sun-synchronous orbit satellite according to claim 7, characterized in that: The calculation formula for the illumination angle β(t) of each track is:

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