A multi-constraint attitude planning method for a fixed sailboard satellite in cis-lunar space

Abstract

The embodiment of the present application provides a multi-constraint attitude planning method for a fixed sailboard satellite in Earth-Moon space, which can solve the technical problem of poor multi-constraint attitude planning capability of a fixed sailboard satellite in Earth-Moon space. The method includes: determining the satellite state quantity at each moment in a preset time period according to the initial value of the orbit; obtaining the sun vector and the earth vector according to the satellite state quantity, and then obtaining the vector angle at each moment; obtaining the target normal vector and the target charging efficiency according to the normal vector and charging efficiency of the dual solar sailboards, and then obtaining the body angle; confirming the ideal time point based on the first preset condition according to the vector angle and the body angle, and then calculating the ideal satellite attitude angle; obtaining the attitude maneuvering time according to the ideal satellite attitude angle and the preset safety index, and iterating according to the preset functional constraint index to obtain the optimized attitude angle. In this way, by separating the attitude angle and the maneuvering time, the multi-constraint attitude planning of a fixed sailboard satellite in Earth-Moon space is realized.

Description

Technical Field

[0001] The present application relates to the technical field of spacecraft mission design, and in particular to a multi-constraint attitude planning method for a fixed sailboard satellite in cis-lunar space. Background Art

[0002] In the mission of spacecraft in the cis-lunar space, attitude planning is a very critical technology to meet various constraints on the normal operation of the satellite. Among them, energy constraints and measurement and control constraints are the most important safety constraints. Most of the sails of cis-lunar space satellites are equipped with solar array drive assemblies (SADA). In order to save costs, a few satellites such as CubeSats use fixed sails. If the satellite has fixed sails or there is an abnormal rotation of the sails, the sails cannot be adjusted, and the satellite can only rely on attitude adjustment to obtain energy to maintain normal operation.

[0003] At present, multi-constraint attitude planning methods, by modeling complex scenes, study various complex optimization algorithms to optimize attitude angles and maneuvering time simultaneously.

[0004] However, there is an over-reliance on optimization algorithms and a lack of understanding and analysis of the model. Moreover, all of the research is done within the scope of near-Earth space, and there is no research on multi-constraint attitude planning for fixed-sail spacecraft in cis-lunar space. Since there are significant differences in mission constraints between near-Earth space and cis-lunar space, the existing attitude planning methods cannot be applied to the attitude planning tasks of fixed-sail spacecraft in cis-lunar space. Summary of the invention

[0005] The embodiment of the present application provides a multi-constraint attitude planning method for a fixed sailboard satellite in the Earth-Moon space, which can solve the technical problem of poor multi-constraint attitude planning capability of a fixed sailboard satellite in the Earth-Moon space.

[0006] To achieve the above objectives, the embodiments of the present application adopt the following technical solutions:

[0007] In a first aspect, an embodiment of the present application provides a multi-constraint attitude planning method for a fixed-sailboard satellite in Earth-Moon space, and the multi-constraint attitude planning method for a fixed-sailboard satellite in Earth-Moon space includes: determining the satellite state quantity at each moment in a preset time period according to the initial orbit value; the satellite state quantity is used to characterize the three-dimensional position parameters and three-dimensional velocity parameters of the satellite; obtaining the sun vector and the earth vector according to the satellite state quantity; obtaining the vector angle at each moment according to the sun vector and the earth vector; obtaining the target normal vector, the dual measurement and control antenna vector and the target charging efficiency according to the normal vector and the charging efficiency of the dual solar panels; the target normal vector and the target charging efficiency are the values ​​when the dual solar panels are converted to a single solar panel; obtaining the body angle according to the target normal vector and the target charging efficiency; confirming the ideal time point based on the first preset condition according to the vector angle and the body angle; calculating the ideal satellite attitude angle at the ideal time point; obtaining the attitude maneuvering time according to the ideal satellite attitude angle and the preset safety index; optimizing the ideal satellite attitude angle and iterating according to the attitude maneuvering time and the preset functional constraint index to obtain the optimized attitude angle.

[0008] Based on the above description of the multi-constraint attitude planning method for a fixed sailboard satellite in the Earth-Moon space provided in the embodiment of the present application, it can be known that the multi-constraint attitude planning method for a fixed sailboard satellite in the Earth-Moon space includes decomposing the attitude planning task into ideal time determination, ideal attitude angle calculation, optimization of maneuvering time with safety indicators, optimization of attitude angle considering functional constraints and iterating to realize multi-constraint attitude planning of a fixed sailboard satellite in the Earth-Moon space based on the geometric analysis of the satellite orbit in the Sun-Earth rotation system.

[0009] In addition, the attitude angle and maneuvering time are processed separately to reduce the complexity of the problem. Furthermore, this fills the gap in the current multi-constraint attitude planning method for fixed-sail spacecraft in the Earth-Moon space.

[0010] In a feasible implementation of the first aspect, when executing the step of obtaining the sun vector and the earth vector based on the satellite state quantity, the multi-constraint attitude planning method of the Earth-Moon space fixed sailboard satellite also includes: converting the satellite state quantity into the first state information in the geocentric solar-Earth rotation system; and obtaining the sun vector and the earth vector based on the first state information.

[0011] In a feasible implementation of the first aspect, the first preset condition includes that the vector angle is equal to the body angle, then there are four ideal time points; the multi-constraint attitude planning method of the Earth-Moon space fixed sailboard satellite also includes: confirming the ideal satellite attitude angle according to each ideal time point; wherein, one ideal time point corresponds to one ideal satellite attitude angle.

[0012] In a feasible implementation of the first aspect, when executing the step of calculating the ideal satellite attitude angle at an ideal time point, the multi-constraint attitude planning method for a fixed sailboard satellite in Earth-Moon space also includes: converting the sun vector and the earth vector into second state information in the satellite inertial system; obtaining a coordinate change matrix based on the ideal time point and the second state information; and obtaining the ideal satellite attitude angle based on the coordinate change matrix.

[0013] In a feasible implementation of the first aspect, when executing the step of obtaining the attitude maneuvering time according to the satellite attitude angle and the preset safety index, the preset safety index includes the preset energy condition and the preset measurement and control condition, and the multi-constraint attitude planning method of the Earth-Moon space fixed sailboard satellite also includes: establishing a target optimization model according to the preset energy condition and the preset measurement and control condition; inputting the ideal satellite attitude angle into the target optimization model to obtain the maneuvering time range; calculating the Pareto front based on the maneuvering time range, the energy optimization target and the measurement and control angle optimization target; and obtaining the attitude maneuvering time according to the Pareto front.

[0014] In a feasible implementation of the first aspect, when executing the step of optimizing the ideal satellite attitude angle and iterating according to the attitude maneuvering time and the preset functional constraint index to obtain the optimized attitude angle, the multi-constraint attitude planning method of the Earth-Moon space fixed sailboard satellite also includes: establishing a functional constraint model according to the functional constraint index; inputting the attitude maneuvering time into the functional constraint model to solve and obtain the first attitude angle; if the first attitude angle and the attitude adjustment time cannot simultaneously meet the preset safety index and the preset functional constraint index, optimizing the first attitude angle according to the iterative method; obtaining the attitude adjustment time according to the Pareto frontier; if the first attitude angle and the attitude adjustment time simultaneously meet the preset safety index and the preset functional constraint index, the first attitude angle is the optimized attitude angle.

[0015] In a feasible implementation of the first aspect, when executing the step of establishing a target optimization model according to preset energy conditions and preset measurement and control conditions, the multi-constraint attitude planning method for a fixed sailboard satellite in Earth-Moon space also includes: establishing an energy optimization target and a first constraint condition according to the preset energy condition; establishing a measurement and control angle optimization target and a second constraint condition according to the preset measurement and control condition; inputting the ideal satellite attitude angle into the first constraint condition to obtain a first time interval; inputting the ideal satellite attitude angle into the second constraint condition to obtain a second time interval; and confirming the intersection of the first time interval and the second time interval to obtain an attitude maneuvering time range.

[0016] In a feasible implementation method of the first aspect, the maneuvering time range is substituted into the energy optimization target to confirm the average energy; the maneuvering time range is substituted into the measurement and control angle optimization target to confirm the average measurement and control angle; the Pareto front is calculated based on the maneuvering time range, average energy and average measurement and control angle; and the attitude maneuvering time is obtained based on the Pareto front.

[0017] In a second aspect, an embodiment of the present application provides a multi-constraint attitude planning system for a fixed sailboard satellite in Earth-Moon space, the multi-constraint attitude planning system for a fixed sailboard satellite in Earth-Moon space comprising: at least one processor; a memory communicatively connected to the at least one processor; wherein the memory stores instructions executable by the at least one processor, the instructions being executed by the at least one processor so that the at least one processor can execute the method provided in the first aspect.

[0018] The multi-constraint attitude planning system for a fixed sailboard satellite in the Earth-Moon space implements the method provided in the first aspect, and based on the geometric analysis of the satellite orbit in the Earth-Sun rotation system, decomposes the attitude planning task into the determination of the ideal time, the calculation of the ideal attitude angle, the optimization of the maneuvering time with the safety index, the optimization of the attitude angle considering the functional constraints and iterates, so as to realize the multi-constraint attitude planning of the fixed sailboard satellite in the Earth-Moon space. In addition, the attitude angle and the maneuvering time are processed separately to reduce the complexity of the problem. In addition, it fills the gap in the current multi-constraint attitude planning method for fixed sail spacecraft in the Earth-Moon space.

[0019] In a third aspect, an embodiment of the present application provides a computer-readable medium having computer program instructions stored thereon, and the computer program instructions can be executed by a processor to implement the method provided in the first aspect.

[0020] The computer program instructions in the computer-readable medium implement the method provided in the first aspect, based on the geometric analysis of the satellite orbit in the solar-terrestrial rotation system, and decompose the attitude planning task into the determination of the ideal time, the calculation of the ideal attitude angle, and the optimization of the maneuvering time with the safety index, so as to realize the multi-constraint attitude planning of the Earth-Moon space fixed sailboard satellite. In addition, the attitude angle and the maneuvering time are processed separately to reduce the complexity of the problem. In addition, the gap in the current multi-constraint attitude planning method of the Earth-Moon space fixed sail spacecraft is filled. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 A schematic diagram of the structure of a multi-constraint attitude planning system for a cis-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0022] Figure 2 A schematic flow chart of a multi-constraint attitude planning method for a fixed sailboard satellite in cis-lunar space provided in an embodiment of the present application;

[0023] Figure 3 A flowchart of an implementation method of a multi-constraint attitude planning method for a fixed sailboard satellite in cis-lunar space provided in an embodiment of the present application;

[0024] Figure 4 A schematic diagram of satellite orbits and related vector angles in a geocentric solar-terrestrial rotation system in a multi-constraint attitude planning method for a terrestrial-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0025] Figure 5 A diagram showing changes in the angle between the satellite's sun vector and the earth vector in a multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0026] Figure 6 A schematic diagram of a feasible maneuvering time range that satisfies energy safety constraints under a theoretical attitude angle configuration in a multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0027] Figure 7 A schematic diagram of a feasible maneuvering time range that satisfies measurement and control safety constraints under a theoretical attitude angle configuration in a multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0028] Figure 8 The Pareto frontier of energy and measurement and control safety indicators under the optimization of attitude maneuvering time in a multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0029] Fig. 9 A satellite energy variation diagram under a multi-constraint attitude maneuver sequence in a multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0030] Fig.10 A satellite measurement and control angle change diagram under a multi-constraint attitude maneuver sequence in a multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite provided in an embodiment of the present application;

[0031] Fig.11 A diagram of the change in the satellite's lunar field of view angle under a multi-constraint attitude maneuver sequence in a multi-constraint attitude planning method for a cis-lunar space fixed-sailboard satellite provided in an embodiment of the present application. DETAILED DESCRIPTION

[0032] The technical solutions in the embodiments of the present invention will be described below in conjunction with the accompanying drawings in the embodiments of the present invention. In the description of the embodiments of the present invention, unless otherwise specified, "multiple" refers to two or more than two. "At least one of the following" or similar expressions refers to any combination of these items, including any combination of single items or plural items. For example, at least one of a, b, or c can represent: a, b, c, ab, ac, bc, or abc, where a, b, and c can be single or multiple.

[0033] In addition, in order to facilitate the clear description of the technical solutions of the embodiments of the present invention, in the embodiments of the present invention, the words "first", "second" and the like are used to distinguish the same items or similar items with substantially the same functions and effects. Those skilled in the art will understand that the words "first", "second" and the like do not limit the quantity and execution order, and the words "first", "second" and the like do not necessarily limit the difference. At the same time, in the embodiments of the present invention, the words "exemplary" or "for example" are used to indicate examples, illustrations or explanations. Any embodiment or design described as "exemplary" or "for example" in the embodiments of the present invention should not be interpreted as being more preferred or more advantageous than other embodiments or design. Specifically, the use of words such as "exemplary" or "for example" is intended to present related concepts in a concrete way for easy understanding.

[0034] The principles and features of the present application are described below. The examples given are only used to explain the present application and are not used to limit the scope of the present application.

[0035] The embodiment of the present application provides a multi-constraint attitude planning method for a fixed sailboard satellite in the Earth-Moon space, which is applicable to various scenarios of satellite operation attitude planning. For example, in view of the need for the Earth-Moon space fixed sail satellite to meet the platform safety constraints (ground measurement and control visibility and energy balance) and other payload functional constraints (such as ground orientation, moon imaging, etc.) during operation, a multi-constraint attitude planning method is proposed. Based on the geometric analysis of the satellite orbit in the Sun-Earth rotation system, the attitude planning task is decomposed into four parts: ideal time determination, ideal attitude angle calculation, optimization of maneuvering time with safety indicators, and local optimization of attitude angle with functional constraints. The attitude angle and maneuvering time are separated and processed to reduce the complexity of the problem. Furthermore, it fills the gap in the current multi-constraint attitude planning method for fixed sail spacecraft in the Earth-Moon space.

[0036] An embodiment of the present application provides a multi-constraint attitude planning system for a fixed sailboard satellite in an Earth-Moon space, which can execute a multi-constraint attitude planning method for a fixed sailboard satellite in an Earth-Moon space provided by an embodiment of the present application. Figure 1 A schematic structural diagram of a multi-constraint attitude planning system for a fixed sailboard satellite in cis-lunar space provided in an embodiment of the present application.

[0037] like Figure 1 As shown, the multi-constraint attitude planning system 001 for a fixed sailboard satellite in Earth-Moon space includes at least one processor 011 and a memory 012 communicatively connected to the at least one processor; wherein the memory 012 stores instructions executable by the at least one processor 011, and the instructions are executed by the at least one processor 011 so that the at least one processor 011 can execute the multi-constraint attitude planning method for a fixed sailboard satellite in Earth-Moon space provided in an embodiment of the present application.

[0038] Figure 2 A schematic diagram of a multi-constraint attitude planning method for a fixed sailboard satellite in cis-lunar space provided in an embodiment of the present application. Figure 2 As shown, in some embodiments, the multi-constraint attitude planning method for the cis-lunar space fixed sailboard satellite includes the following steps:

[0039] S1, determine the satellite status at each moment in the preset time period based on the initial orbit value.

[0040] In some embodiments, at the epoch time t0 of the Earth-Moon space satellite, in the Earth-Moon rotation system, the orbit initial value is The coordinate origin of the Earth-Moon rotation system is the common center of mass of the Earth and the Moon. The X-axis points from the center of the Earth to the center of the Moon. The Z-axis is along the angular momentum direction of the Moon relative to the Earth's orbit. The Y-axis, X-axis, and Z-axis form a right-handed rectangular coordinate system.

[0041] In some embodiments, multiple moments are determined within a preset time period T with a fixed step length as the time interval. In one implementation, the preset time period is a lunar month, which may be 29.53 days. In this way, the preset time period is set based on the period of the periodic retrograde rotation of the sun around the earth-moon system. With a step length of 10 minutes, multiple moments within 29.53 days are selected.

[0042] In some embodiments, the orbital initialization The satellite state quantity is obtained by integrating the time [t0, t0+T], that is, x em (t) = [r em (t),v em (t)]([t0,t0+T]).

[0043] Satellite state quantity, used to characterize the three-dimensional position parameters and three-dimensional velocity parameters of the satellite. Wherein, r represents the position and v represents the velocity. In one implementation, at the epoch time of the Earth-Moon space satellite, in the Earth-Moon rotation system, the three-dimensional position parameters are the values ​​of the Y-axis, X-axis and Z-axis directions, and the three-dimensional velocity parameters are the values ​​of the velocity in the Y-axis, X-axis and Z-axis directions.

[0044] S2, based on the satellite state quantity, obtain the sun vector and the earth vector.

[0045] To simplify the posture analysis, Figure 3 As shown, in some embodiments, when executing the step of obtaining the sun vector and the earth vector according to the satellite state quantity, the multi-constraint attitude planning method for the cis-lunar space fixed sailboard satellite also includes:

[0046] S211, converting the satellite state quantity into the first state information in the geocentric solar-terrestrial rotation system.

[0047] As the satellite state quantity is converted into the first state information, the Earth-Moon space orbit is converted from the Earth-Moon rotation system to the Earth-Center Sun-Earth Rotation System (ECSER), and the first state information x under ECSER is obtained. se (t) = [r se (t),v se (t)]. The origin of the geocentric sun-earth rotation system is the center of the earth, the X-axis points from the center of the sun to the center of the earth, and the Z-axis is along the angular momentum direction of the earth's orbit relative to the sun. The X-axis, Z-axis and Y-axis form a right-handed rectangular coordinate system.

[0048] like Figure 4 As shown, in the geocentric sun-earth rotation system, the Earth-Moon space orbit approximately circles the Earth in one lunar month, the Sun vector approximately remains unchanged along the negative direction of the X-axis, and the Earth vector is the inverse vector of the satellite position vector.

[0049] S212, obtaining a sun vector and an earth vector according to the first state information.

[0050] Based on x se (t) ， Under ECSER, the Earth vector r e The opposite of the satellite position vector, i.e. r e (t) = -r se (t). The sun vector is r s (t) = r se (t)-r s2e , where r s2e =[1,0,0], which indicates the vector from the sun to the earth under ECSER. In the sun-earth rotation system, the sun vector is approximately constant along the negative direction of the X axis.

[0051] like Figure 4 As shown, the sun vector and the earth vector at several moments under ECSER.

[0052] In some embodiments, the sun vector and the earth vector at each of the multiple moments in the preset time period are determined. For example, the sun vector and the earth vector at each moment in the preset time period of a lunar month with a fixed time interval of 10 minutes are calculated.

[0053] It can be understood that according to the sun vector r s and Earth vector e , the satellite current and the tracking angle at a certain moment can be obtained. Through the tracking angle, the maximum angle between the antenna and the ground is limited, that is, the angle between the antenna and the ground is less than or equal to the tracking angle, to meet the tracking requirements. Among them, the antenna can be a dual antenna, which is used to meet the satellite's tracking requirements by receiving ground station signals.

[0054] S3, based on the sun vector and the earth vector, obtain the vector angle at each moment.

[0055] The vector angle is α(t)=cos -1 (r s *r e ),t∈[t0,t0+T].

[0056] like Figure 5 As shown, the vector angle first decreases from nearly 180 degrees to nearly 0 degrees, and then increases to nearly 180 degrees.

[0057] S4, obtaining a target normal vector and a target charging efficiency according to the normal vectors and the charging efficiency of the dual solar panels.

[0058] In some embodiments, the dual solar sail panels are configured as fixed sail panels, which generate current to provide energy for the satellite by receiving sunlight. The solar sail current is proportional to the cosine value of the angle between the sun vector and the sail panel normal vector. The smaller the angle, the greater the current. If it is lower than the minimum energy requirement, the satellite cannot operate normally.

[0059] The target normal vector and target charging efficiency are the values ​​when the dual solar panels are converted to a single solar panel.

[0060] In some embodiments, the normal vectors of the dual solar sail panels are n1 and n2, and the charging efficiencies are η1 and η2, respectively. The target normal vector n is calculated as n=(η1n1+η2n2) / |η1n1+η2n2|, and the target charging efficiency is calculated as η=|η1n1+η2n2|. That is, for the satellite, the maximum current that can be generated is η, and at this time, the sun vector and the solar sail normal vector are collinear, and the angle is 0.

[0061] S5, obtaining the body angle according to the target normal vector, the dual measurement and control antenna vector and the target charging efficiency.

[0062] Body angle is the body angle between the satellite solar sail normal vector and the antenna vector, and the calculation formula is:

[0063]

[0064] Among them, z1 and z2 represent the dual measurement and control antenna vectors, and n represents the target normal vector.

[0065] S6, confirming the ideal time point according to the vector angle and the body angle based on the first preset condition.

[0066] Calculate the ideal time point when the angle between the sun vector and the earth vector is equal to the body angle

[0067]

[0068] Based on the above formula, find The nearest time point as the ideal time point

[0069] like Figure 4 As shown, in some embodiments, the first preset condition includes that the vector angle is equal to the body angle, and there are four ideal time points. It can be understood that, from a geometric point of view, under ECSER, there must be two ideal moments in the Earth-Moon space orbit of a lunar month to satisfy that the angle between the sun vector and the earth vector is equal to the body solar sail normal vector and the antenna angle. At this moment, the sun vector and the sailboard normal vector are collinear at the same time, and the antenna is facing the center of the earth, that is, the energy and measurement and control conditions are optimized at the same time. For a dual-antenna satellite, there are four ideal moments.

[0070] In some embodiments, the antennas in this system are anti-vectors to each other and are approximately symmetrical between two of them at four ideal moments in the orbit under ECSER.

[0071] S7, calculate the ideal satellite attitude angle at the ideal time point.

[0072] In some embodiments, when executing the step of calculating the ideal satellite attitude angle at the ideal time point, the multi-constraint attitude planning method for a terrestrial-lunar space fixed sailboard satellite further includes:

[0073] S711, converting the sun vector and the earth vector into second state information in the satellite inertial system.

[0074] The four ideal points The sun vector at r s and Earth vector e , converted to the solar vector in the satellite inertial system and earth vector and It is the second state information.

[0075] S712, obtaining a coordinate change matrix according to the ideal time point and the second state information.

[0076] Based on the geometric relationship that the sun vector and the normal vector of the sailboard are collinear and the earth vector and the antenna vector are collinear at the ideal time point, the ideal inertial attitude angle E is calculated. k , that is, solving the system of equations:

[0077]

[0078] Among them, R k is the attitude angle E kThe coordinate transformation matrix from the satellite local system to the satellite inertial system is as follows:

[0079]

[0080] S713, obtaining the ideal satellite attitude angle according to the coordinate change matrix.

[0081] Four ideal time points The optimal attitude angle is Rotate in the order of Z axis, Y axis, and X axis.

[0082] In some embodiments, when executing step S7, the multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite further includes:

[0083] S721, confirming the ideal satellite attitude angle according to each ideal time point.

[0084] Among them, an ideal time point corresponds to an ideal satellite attitude angle.

[0085] In this way, at the ideal moment, the angle between the sun vector and the earth vector is equal to the angle between the sailboard normal vector and the antenna. By configuring the attitude angle, the sun vector is facing the sailboard normal vector and the antenna is facing the center of the earth, achieving the best energy and measurement and control conditions at the same time. This attitude angle is the ideal satellite attitude angle, also known as the optimal attitude angle. By determining the ideal satellite attitude angle through geometric analysis, it is only necessary to plan the attitude time period, avoiding complex optimization calculations of the attitude angle.

[0086] S8: Obtain attitude maneuvering time based on the ideal satellite attitude angle and preset safety indicators.

[0087] In this way, the satellite attitude angle is fixed and the attitude maneuvering time is optimized with safety as the indicator.

[0088] The ideal satellite attitude angle in step S7 is used as the attitude angle in the attitude maneuver sequence. The attitude angle in the attitude maneuver sequence is fixed, and the attitude maneuver time is globally optimized with safety as the optimization goal. Set the optimization variable As the attitude maneuver time, Only the time of the next three attitude maneuvers needs to be optimized. The attitude angle after the kth attitude maneuver is E k , k=0,1,2,3.

[0089] In some embodiments, when executing the step of fixing the satellite attitude angle and optimizing the attitude maneuvering time with safety as an indicator, the multi-constraint attitude planning method for a fixed sailboard satellite in the cis-lunar space further includes:

[0090] A target optimization model is established based on preset energy conditions and preset measurement and control conditions.

[0091] In this way, the target optimization model is a multi-objective optimization model for safety.

[0092] Exemplarily, the multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite also includes:

[0093] S811, establishing an energy optimization target and a first constraint condition according to a preset energy condition.

[0094] In this way, according to the preset energy conditions, the energy optimization target and the first constraint condition are established.

[0095] The first constraint is the energy security constraint, which requires that the current at all times in the planning section is not less than the minimum energy requirement I min ,Right now:

[0096]

[0097] Indicated in Time planning and E k T in attitude configuration j The satellite current at the moment is proportional to the cosine of the angle between the normal vector of the sailboard and the sun vector, that is:

[0098]

[0099] The energy optimization target is the energy security index, which is the negative of the average energy over the entire period:

[0100]

[0101] S812, establishing a measurement and control angle optimization target and a second constraint condition according to preset measurement and control conditions.

[0102] In this way, based on the preset measurement and control conditions, the measurement and control angle optimization target and the second constraint condition are established.

[0103] The second constraint is the measurement and control safety constraint, which requires that the antenna measurement and control angle at all times in the planning section is not greater than the maximum antenna measurement and control angle φ max ,Right now:

[0104]

[0105] Indicated in Maneuvering time and E k The antenna control angle under attitude configuration is:

[0106]

[0107] The optimization target of the measurement and control angle is the measurement and control safety index, that is, the average measurement and control angle for the entire period:

[0108]

[0109] S813, establish a complete target optimization model.

[0110] Take all the time t of the planning period j (t0≤t j The average energy and control angle of the time interval ≤t0+T) are used as safety indicators, and a multi-objective optimization model is established (including energy, control angle optimization objectives and the first constraint and the second constraint):

[0111]

[0112] In some embodiments, when executing the step of fixing the satellite attitude angle and optimizing the attitude maneuvering time with safety as an indicator, the multi-constraint attitude planning method for a fixed sailboard satellite in the cis-lunar space further includes:

[0113] The ideal satellite attitude angle is input into the constraint condition to obtain the maneuvering time range. The multi-constraint attitude planning method for the Earth-Moon space fixed sailboard satellite also includes:

[0114] S821, input the ideal satellite attitude angle into the first constraint condition to obtain a first time interval.

[0115] S822, input the ideal satellite attitude angle into the second constraint condition to obtain a second time interval.

[0116] S823, confirming the intersection of the first time interval and the second time interval to obtain an attitude maneuvering time that satisfies the constraint.

[0117] Combined Figure 6 and Figure 7 , represents the feasible attitude maneuvering time range when only energy security or measurement and control security is considered. The intersection of the two feasible ranges is taken to obtain the maneuvering time range.

[0118] In some embodiments, the multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite further includes:

[0119] The Pareto frontier of energy and measurement and control angle optimization objectives is calculated to obtain the optimal maneuvering time. The multi-constraint attitude planning method for the Earth-Moon space fixed sailboard satellite also includes:

[0120] S831, substituting the attitude maneuvering time range into the energy optimization target to determine the average energy;

[0121] S832, substituting the attitude maneuvering time range into the measurement and control angle optimization target to determine the average measurement and control angle;

[0122] S833, calculate the Pareto frontier based on the attitude maneuver time range, average energy and average measurement and control angle;

[0123] The Pareto front is the Pareto front of energy optimization and measurement and control angle optimization. Figure 8 , representing the Pareto frontier of energy and measurement and control.

[0124] S834, according to the Pareto front, obtain the attitude maneuvering time.

[0125] In this way, the optimization model is simplified, and the average energy and measurement and control angle within the maneuvering time range are used as safety indicators. The attitude adjustment time is obtained as the updated attitude maneuvering time.

[0126] S9, optimizing the ideal satellite attitude angle and performing iterations according to the attitude maneuvering time and the preset functional constraint index to obtain an optimized attitude angle.

[0127] In this way, the attitude maneuvering time is fixed, the attitude angle is optimized considering the functional constraints and iterated.

[0128] In this embodiment, two functional constraints are given.

[0129] Functional constraint 1 requires that the planning section must have 3 days of geo-orientation time. Geo-orientation requires that the satellite antenna remain collinear with the satellite earth vector for 3 consecutive days. In order to complete the geo-orientation task, the attitude angle needs to be set under the satellite VVLH system. Since there are only 3 days, and the energy measurement and control safety constraints must be met as much as possible, the attitude angle is locally optimized around the ideal time point.

[0130] Functional constraint 2 requires that lunar imaging can be performed on the last day of the planning period. Lunar imaging requires that the angle between the normal direction of the camera installed on the satellite and the lunar vector (lunar field of view angle) be smaller than the camera field of view angle.

[0131] In some embodiments, when executing the steps of fixing the attitude maneuvering time, optimizing the attitude angle by considering the functional constraints and iterating, the multi-constraint attitude planning method for the Earth-Moon space fixed sailboard satellite further includes:

[0132] S911: Establish a functional constraint model according to the functional constraint index.

[0133] In this embodiment, considering functional constraint 1, the local posture maneuvering time is fixed as the ideal time point One and a half days before and after, set the initial value of the local attitude angle of the VVLH system to

[0134] The ground orientation requirement is θ seq =0,

[0135] The local attitude angle is optimized with the local average energy, and the optimization model is established as follows:

[0136]

[0137] Considering functional constraint 2, the local attitude maneuvering time is fixed to the last day of the planning period, and the initial value of the local attitude angle is set to E seq =[0,0,0], satisfying the constraint of the lunar field of view angle β seq (t j ) is no greater than the camera field of view angle β max , optimize the local attitude angle with the local average energy, and establish the optimization model as follows:

[0138]

[0139] in, is the moon vector in the inertial system, and c is the normal direction of the camera in the satellite's own system.

[0140] S912, inputting the attitude maneuvering time into the functional constraint model, and solving to obtain the first attitude angle.

[0141] The optimal maneuvering time is fixed, and the optimization model is solved to obtain the first attitude angle, so as to further obtain the optimal attitude angle that meets the functional constraints.

[0142] In this embodiment, the above two optimization models are solved separately to obtain the optimal posture angle E that satisfies each functional constraint. seq .

[0143] S913, determining whether the current optimal attitude angle and attitude adjustment time can satisfy both safety and functional constraints. If not, performing iteration.

[0144] That is to say, if the first posture angle and the posture adjustment time cannot simultaneously meet the preset safety indicators and the preset functional constraint indicators, the first posture angle is optimized according to the iterative method; the posture adjustment time is obtained according to the Pareto frontier; if the first posture angle and the posture adjustment time simultaneously meet the preset safety indicators and the preset functional constraint indicators, the first posture angle is the optimized posture angle.

[0145] The specific iterative method is to fix the local attitude angle, optimize the local maneuvering time considering the safety constraint, and iterate until the attitude maneuvering sequence meets the constraint. In this embodiment, the local attitude planning sequence is verified to not meet the safety constraint, so iterative calculation is performed to finally obtain an attitude maneuvering sequence that meets both the functional constraint and the safety constraint.

[0146] Fig. 9 , Fig.10 and Fig.11 The diagrams showing the changes in satellite energy, antenna tracking angle, and lunar field of view angle during this attitude maneuver sequence are given in Figure 2. Fig. 9As shown in Figure 1, the satellite energy under the current attitude planning is always kept close to the optimal value. The local current drop in the middle is due to the functional constraint of setting the three-day orientation to the ground, but it still meets the minimum energy constraint. Fig.10 As shown in the figure, the satellite antenna tracking angle under the current attitude planning always meets the tracking angle constraint. At the same time, the tracking angle is 0 for three consecutive days, achieving the functional constraint 1 of the three-day orientation to the earth. Fig.11 As shown, the lunar field of view angle under the current attitude planning meets the functional constraint 2 on the last day, and the lunar field of view angle is smaller than the camera field of view angle.

[0147] Based on the same application concept, an embodiment of the present application also provides a multi-constraint attitude planning system for a fixed sailboard satellite in Earth-Moon space, and the method corresponding to the multi-constraint attitude planning system for a fixed sailboard satellite in Earth-Moon space can be the multi-constraint attitude planning method for a fixed sailboard satellite in Earth-Moon space in the aforementioned embodiment, and its principle of solving the problem is similar to that of the method. The multi-constraint attitude planning system for a fixed sailboard satellite in Earth-Moon space provided in the embodiment of the present application includes: at least one processor; and a memory communicatively connected to the at least one processor; wherein the memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor so that the at least one processor can execute the methods and / or technical solutions of the aforementioned multiple embodiments of the present application.

[0148] Another embodiment of the present application further provides a computer-readable storage medium having computer program instructions stored thereon, wherein the computer program instructions can be executed by a processor to implement the methods and / or technical solutions of any one or more embodiments of the present application described above.

[0149] Specifically, the present embodiment may adopt any combination of one or more computer-readable media. The computer-readable medium may be a computer-readable signal medium or a computer-readable storage medium. The computer-readable storage medium may be, for example, but not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device or device, or any combination thereof. More specific examples (non-exhaustive list) of computer-readable storage media include: an electrical connection with one or more wires, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination thereof. In this document, a computer-readable storage medium may be any tangible medium containing or storing a program that may be used by or in combination with an instruction execution system, device, or device.

[0150] Computer readable signal media may include a data signal propagated in baseband or as part of a carrier wave, which carries a computer readable program code. Such propagated data signals may take a variety of forms, including, but not limited to, electromagnetic signals, optical signals, or any suitable combination of the above. Computer readable signal media may also be any computer readable medium other than a computer readable storage medium, which may send, propagate, or transmit a program for use by or in conjunction with an instruction execution system, apparatus, or device.

[0151] Program code embodied on a computer readable medium may be transmitted using any appropriate medium, including but not limited to wireless, wireline, optical fiber cable, RF, etc., or any suitable combination of the foregoing.

[0152] Computer program code for performing the operations of the present application may be written in one or more programming languages ​​or a combination thereof, including object-oriented programming languages, such as Java, Smalltalk, C++, and conventional procedural programming languages, such as "C" or similar programming languages. The program code may be executed entirely on the user's computer, partially on the user's computer, as a separate software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In the case of a remote computer, the remote computer may be connected to the user's computer through any type of network, including a local area network (LAN) or a wide area network (WAN), or may be connected to an external computer (e.g., via the Internet using an Internet service provider).

[0153] The flow chart or block diagram in the accompanying drawings shows the possible architecture, function and operation of the equipment, method and computer program product according to various embodiments of the present application. In this regard, each square box in the flow chart or block diagram can represent a module, a program segment or a part of a code, and the module, the program segment or a part of the code contains one or more executable instructions for realizing the specified logical function. It should also be noted that in some implementations as replacements, the functions marked in the square box can also occur in a sequence different from that marked in the accompanying drawings. For example, two square boxes represented in succession can actually be executed substantially in parallel, and they can sometimes be executed in the opposite order, depending on the functions involved. It should also be noted that each square box in the block diagram and / or flow chart, and the combination of the square boxes in the block diagram and / or flow chart can be implemented with a dedicated system for hardware that performs a specified function or operation, or can be implemented with a combination of dedicated hardware and computer instructions.

[0154] Those skilled in the art can clearly understand that, for the convenience and brevity of description, the specific working processes of the systems, devices and units described above can refer to the corresponding processes in the aforementioned method embodiments and will not be repeated here.

[0155] In the several embodiments provided in the present application, it should be understood that the disclosed systems, devices and methods can be implemented in other ways. For example, the device embodiments described above are only schematic. For example, the division of the units is only a logical function division. There may be other division methods in actual implementation. For example, multiple units or page components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed can be an indirect coupling or communication connection through some interfaces, devices or units, which can be electrical, mechanical or other forms.

[0156] The units described as separate components may or may not be physically separated, and the components shown as units may or may not be physical units, that is, they may be located in one place or distributed on multiple network units. Some or all of the units may be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0157] In addition, each functional unit in each embodiment of the present application may be integrated into one processing unit, or each unit may exist physically separately, or two or more units may be integrated into one unit. The above-mentioned integrated unit may be implemented in the form of hardware or in the form of hardware plus software functional units.

[0158] The above-mentioned integrated unit implemented in the form of a software functional unit can be stored in a computer-readable storage medium. The above-mentioned software functional unit is stored in a storage medium, including a number of instructions for a computer device (which can be a personal computer, a server, or a network device, etc.) or a processor to perform some steps of the method described in each embodiment of the present application. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (Read-Only Memory, ROM), random access memory (Random Access Memory, RAM), disk or optical disk and other media that can store program codes.

[0159] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present application, rather than to limit it. Although the present application has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present application.

[0160] In addition, it is clear that the word "comprising" does not exclude other units or steps, and the singular does not exclude the plural. Multiple units or devices stated in a device claim can also be implemented by one unit or device through software or hardware. The words first, second, etc. are used to indicate names, and do not indicate any particular order.

Claims

1. A multi-constraint attitude planning method for a fixed sailboard satellite in cis-lunar space, characterized in that: include: According to the initial orbit value, determine the satellite state quantity at each moment in the preset time period; The satellite state quantity is used to characterize the three-dimensional position parameters and three-dimensional velocity parameters of the satellite; According to the satellite state quantity, a sun vector and an earth vector are obtained; According to the sun vector and the earth vector, obtaining the vector angle at each time; According to the normal vector and charging efficiency of the dual solar panels, a target normal vector and a target charging efficiency are obtained; the target normal vector and the target charging efficiency are values ​​when the dual solar panels are converted into a single solar panel; Obtaining a body angle according to the target normal vector, the dual measurement and control antenna vector and the target charging efficiency; According to the vector angle and the body angle, based on a first preset condition, determining an ideal time point; Calculate the ideal satellite attitude angle at the ideal time point; Obtaining attitude maneuvering time according to the ideal satellite attitude angle and a preset safety indicator; According to the attitude maneuvering time and the preset functional constraint index, the ideal satellite attitude angle is optimized and iterated to obtain the optimized attitude angle.

2. The multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite according to claim 1, characterized in that: When executing the step of obtaining the sun vector and the earth vector according to the satellite state quantity, the multi-constraint attitude planning method for the cis-lunar space fixed sailboard satellite also includes: Converting the satellite state quantity into first state information in the geocentric solar-terrestrial rotation system; The sun vector and the earth vector are obtained according to the first state information.

3. The multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite according to claim 1 or 2, characterized in that: The first preset condition includes that the vector angle is equal to the body angle, and then there are four ideal time points; the multi-constraint attitude planning method for the Earth-Moon space fixed sailboard satellite also includes: The ideal satellite attitude angle is confirmed according to each of the ideal time points; wherein one of the ideal time points corresponds to one of the ideal satellite attitude angles.

4. The multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite according to claim 1 or 2, characterized in that: When executing the step of calculating the ideal satellite attitude angle at the ideal time point, the multi-constraint attitude planning method for the Earth-Moon space fixed sailboard satellite also includes: Converting the sun vector and the earth vector into second state information in a satellite inertial system; Obtaining a coordinate change matrix according to the ideal time point and the second state information; The ideal satellite attitude angle is obtained according to the coordinate change matrix.

5. The multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite according to claim 1 or 2, characterized in that: When executing the step of obtaining the attitude maneuvering time according to the ideal satellite attitude angle and the preset safety index, the preset safety index includes a preset energy condition and a preset measurement and control condition, and the multi-constraint attitude planning method for the Earth-Moon space fixed sailboard satellite also includes: Establish a target optimization model based on preset energy conditions and preset measurement and control conditions; Inputting the ideal satellite attitude angle into the target optimization model to obtain a maneuvering time range; Calculating the Pareto frontier based on the maneuvering time range, energy optimization target and measurement and control angle optimization target; According to the Pareto front, the attitude maneuvering time is obtained.

6. The multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite according to claim 5, characterized in that: When executing the step of optimizing the ideal satellite attitude angle and iterating according to the attitude maneuvering time and the preset functional constraint index to obtain the optimized attitude angle, the multi-constraint attitude planning method for the Earth-Moon space fixed sailboard satellite also includes: Establishing a functional constraint model according to the functional constraint index; Inputting the attitude maneuvering time into the functional constraint model to solve for a first attitude angle; If the first posture angle and the posture adjustment time cannot simultaneously meet the preset safety index and the preset functional constraint index, optimizing the first posture angle according to an iterative method; According to the Pareto frontier, the posture adjustment time is obtained; If the first posture angle and the posture adjustment time simultaneously meet the preset safety index and the preset functional constraint index, the first posture angle is the optimized posture angle.

7. The multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite according to claim 5, characterized in that: When executing the step of establishing a target optimization model according to the preset energy conditions and the preset measurement and control conditions, the multi-constraint attitude planning method for a fixed sailboard satellite in the terrestrial and lunar space further includes: According to the preset energy conditions, establishing the energy optimization target and the first constraint condition; According to the preset measurement and control conditions, the measurement and control angle optimization target and the second constraint condition are established; Inputting the ideal satellite attitude angle into a first constraint condition to obtain a first time interval; Inputting the ideal satellite attitude angle into a second constraint condition to obtain a second time interval; An intersection of the first time interval and the second time interval is determined to obtain the maneuvering time range.

8. The multi-constraint attitude planning method for a cis-lunar space fixed sailboard satellite according to claim 7, characterized in that: The multi-constraint attitude planning method for a fixed sailboard satellite in the cis-lunar space also includes: Substituting the maneuvering time range into the energy optimization target to determine the average energy; Substituting the maneuvering time range into the measurement and control angle optimization target to determine the average measurement and control angle; Calculating a Pareto front according to the maneuvering time range, the average energy and the average measurement and control angle; According to the Pareto front, the attitude maneuvering time is obtained.

9. A multi-constraint attitude planning system for a fixed sailboard satellite in cis-lunar space, characterized in that: include: at least one processor; a memory communicatively coupled to the at least one processor; The memory stores instructions that can be executed by the at least one processor, and the instructions are executed by the at least one processor so that the at least one processor can execute the method according to any one of claims 1 to 8.

10. A computer readable medium having computer program instructions stored thereon, wherein the computer program instructions can be executed by a processor to implement the method according to any one of claims 1 to 9.

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