Deep space satellite on-orbit autonomous inertial sun pointing method, computing device and readable medium

Abstract

The application provides a deep space satellite on-orbit autonomous inertia sun pointing method, a computing device and a readable medium. The method is suitable for a satellite with a sailboard normal line pointing to a non-design direction, and comprises the following steps: obtaining a local system sun target vector, the local system sun target vector being a local system sun vector when a sailboard current value is maximum; calculating an included angle between the local system sun target vector and a local system positive Z axis, and recording the included angle as a first included angle; calculating a projection of the local system positive Z axis in an inertia system according to the first included angle, an inertia system sun vector and an inertia system earth vector, and recording the projection as a first candidate pointing vector; calculating an included angle between the first candidate pointing vector and the inertia system earth vector, and recording the included angle as a second included angle; and determining a target pointing vector and a target deflection angle of satellite rotation according to the second included angle and a range of a measurement and control antenna beam. The application realizes autonomous sun pointing of a satellite local system arbitrary vector in an inertia space and attitude control of taking into account measurement and control requirements.

Description

TECHNICAL FIELD

[0001] The present application mainly relates to the technical field of satellite attitude control, and particularly relates to a deep space satellite on-orbit autonomous inertia sun pointing method, a computing device and a readable medium. BACKGROUND

[0002] The on-orbit autonomous inertia sun pointing method of the deep space satellite refers to a technology that, in a deep space environment far from the earth, a satellite autonomously adjusts an attitude by using on-board sensors (such as a sun sensor, a star sensor, an inertial measurement unit, etc.) and a control system without relying on real-time instructions from the ground, so that a solar array continuously points to the sun while meeting other task constraints (such as earth communication, load pointing, etc.).

[0003] The existing method is to control the rotation of the star body and the rotation of the sailboard so that the sailboard normal line points to the sun. When the sailboard fails to rotate or the sailboard normal line points to a non-design direction, the existing sun pointing method cannot be applied. SUMMARY

[0004] The present application mainly relates to the technical field of satellite attitude control, and particularly relates to a deep space satellite on-orbit autonomous inertia sun pointing method, a computing device and a readable medium.

[0005] To solve the above technical problems, the present application provides an on-orbit autonomous inertia sun pointing method for a deep space satellite, which is applicable to a satellite with a sailboard normal line pointing to a non-design direction, and includes: obtaining a body system sun target vector, which is a body system sun vector when the sailboard current value is maximum; calculating an included angle between the body system sun target vector and a body system positive Z axis, denoted as a first included angle; calculating a projection of the body system positive Z axis in an inertial system according to the first included angle, an inertial system sun vector and an inertial system earth vector, denoted as a first candidate pointing vector; calculating an included angle between the first candidate pointing vector and the inertial system earth vector, denoted as a second included angle; and determining a target pointing vector and a target deflection angle of satellite rotation according to the second included angle and a range of a measurement and control antenna beam.

[0006] Optionally, calculating the projection of the body system positive Z axis in the inertial system according to the first included angle, the inertial system sun vector and the inertial system earth vector includes: establishing a reference coordinate system according to the inertial system sun vector and the inertial system earth vector; obtaining a projection of the first candidate pointing vector in the reference coordinate system by rotating an X axis in the reference coordinate system around a Z axis by the first included angle, denoted as a reference system Z axis vector; and calculating the first candidate pointing vector according to the reference system Z axis vector and a conversion matrix from the inertial system to the reference coordinate system.

[0007] Optionally, establishing the reference coordinate system according to the inertial system sun vector and the inertial system earth vector comprises: taking the inertial system sun vector as an X axis, taking a normal of a plane composed of the inertial system sun vector and the inertial system earth vector as a Z axis, and taking a Y axis perpendicular to the X axis and the Z axis.

[0008] Optionally, the reference system Z axis vector is calculated by the following formula:

[0009]

[0010]

[0011] wherein, is the reference system Z axis vector, is the first included angle, is the inertial system sun vector.

[0012] Optionally, the first candidate pointing vector is calculated by the following formula:

[0013]

[0014] wherein, is the first candidate pointing vector, is a conversion matrix from the inertial system to the reference coordinate system, is the reference system Z axis vector.

[0015] Optionally, determining a target pointing vector and a target declination of satellite rotation according to the second included angle and a range of a TT&C antenna beam comprises:

[0016] Let the second included angle be , the target pointing vector be , the target declination be , and the range of the TT&C antenna beam be and The target pointing vector and the target declination are calculated by the following ways:

[0017] When , a +Z plane TT&C antenna is used for earth communication, and no angle compensation is needed, , .

[0018] When , a -Z plane TT&C antenna is used for earth communication, and no angle compensation is needed, , .

[0019] Optionally, it further comprises: when or When the first candidate pointing vector is determined, the target deflection angle is determined by compensating the angle between the earth and the inertial system, the direction of rotation and the angle of rotation of the satellite are determined by a trial rotation algorithm, and the target pointing vector of the satellite rotation is determined according to the direction of rotation and the angle of rotation.

[0020] Optionally, when the first candidate pointing vector is determined, the target deflection angle is determined by compensating the angle between the earth and the inertial system, the direction of rotation and the angle of rotation of the satellite are determined by a trial rotation algorithm, and the target pointing vector of the satellite rotation is determined according to the direction of rotation and the angle of rotation. When the first candidate pointing vector is determined, the target deflection angle is determined by compensating the angle between the earth and the inertial system, the direction of rotation and the angle of rotation of the satellite are determined by a trial rotation algorithm, and the target pointing vector of the satellite rotation is determined according to the direction of rotation and the angle of rotation. When the first candidate pointing vector is determined, the target deflection angle is determined by compensating the angle between the earth and the inertial system, the direction of rotation and the angle of rotation of the satellite are determined by a trial rotation algorithm, and the target pointing vector of the satellite rotation is determined according to the direction of rotation and the angle of rotation. Flag is the direction of rotation of the satellite, and the value of Flag is 1, indicating that the direction of rotation is positive rotation, and the value of Flag is -1, indicating that the direction of rotation is negative rotation.

[0021] Optionally, when the first candidate pointing vector is determined, the target deflection angle is determined by compensating the angle between the earth and the inertial system, the direction of rotation and the angle of rotation of the satellite are determined by a trial rotation algorithm, and the target pointing vector of the satellite rotation is determined according to the direction of rotation and the angle of rotation. When the first candidate pointing vector is determined, the target deflection angle is determined by compensating the angle between the earth and the inertial system, the direction of rotation and the angle of rotation of the satellite are determined by a trial rotation algorithm, and the target pointing vector of the satellite rotation is determined according to the direction of rotation and the angle of rotation. When the first candidate pointing vector is determined, the target deflection angle is determined by compensating the angle between the earth and the inertial system, the direction of rotation and the angle of rotation of the satellite are determined by a trial rotation algorithm, and the target pointing vector of the satellite rotation is determined according to the direction of rotation and the angle of rotation. Flag is the direction of rotation of the satellite, and the value of Flag is 1, indicating that the direction of rotation is positive rotation, and the value of Flag is -1, indicating that the direction of rotation is negative rotation.

[0022] Optionally, the target pointing vector of the satellite rotation is calculated by the following formula:

[0023]

[0024]

[0025] wherein, is the target pointing vector, is the first candidate pointing vector, is the target deflection angle.

[0026] Optionally, further comprising: calculating an included angle between the inertial system sun vector and the inertial system earth vector, denoted as a fifth included angle; judging whether a value of the fifth included angle is less than a first threshold angle or greater than a second threshold angle, and if so, the satellite maintains a current target pointing vector and a target declination angle, wherein a sum of the first threshold angle and the second threshold angle is 180°, and the first threshold angle is less than or equal to 10°.

[0027] To solve the above technical problems, the application provides a computing device, comprising: a memory for storing instructions executed by a processor; and the processor is configured to execute the instructions to implement the on-orbit autonomous inertial sun pointing method of the deep space satellite according to any one of the above.

[0028] To solve the above technical problems, the application provides a computer readable medium storing computer program codes, which, when executed by a processor, implement the on-orbit autonomous inertial sun pointing method of the deep space satellite.

[0029] Compared with the prior art, the application has the following advantages:

[0030] The on-orbit autonomous inertial sun pointing method of the deep space satellite, the computing device and the readable medium of the application, on the one hand, realize that the satellite body system arbitrary vector points to the sun, that is, the specific vector direction of the satellite is consistent with the direction pointing to the sun, and ensure the energy supply of the satellite when the sailboard SADA is not designed to point in the expected direction; on the other hand, the application realizes the compensation control of the consideration of the earth measurement and control angle, and ensures the measurement and control demand of the satellite in the complex deep space orbit. BRIEF DESCRIPTION OF DRAWINGS

[0031] The accompanying drawings are included to provide a further understanding of the application, and they are collected and constitute a part of the application, which shows the embodiments of the application, and together with the specification, plays a role in explaining the principles of the application. In the drawings:

[0032] Figure 1 is the demand-oriented diagram of the autonomous inertial sun pointing algorithm.

[0033] Figure 2 is the flowchart of the on-orbit autonomous inertial sun pointing method of the deep space satellite according to an embodiment of the application.

[0034] Figure 3 is Figure 2 is the flowchart of step S23 of an embodiment.

[0035] Figure 4 is the schematic diagram of the reference coordinate system according to an embodiment of the application.

[0036] Figure 5 is the satellite sun pointing simulation result diagram according to an embodiment of the application.

[0037] Figure 6 is a simulation diagram of the angle between the Z-axis of the body frame of an embodiment of the present application and the earth's central vector.

[0038] Figure 7 is a system block diagram of a computing device of an embodiment of the present application. DETAILED DESCRIPTION

[0039] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following will briefly introduce the drawings needed to be used in the embodiment description. Obviously, the drawings in the following description are only some examples or embodiments of the present application, and for those skilled in the art, the present application can also be applied to other similar scenarios without creative labor on the basis of these drawings. Unless it is obvious from the language environment or otherwise stated, the same reference numbers in the drawings represent the same structure or operation.

[0040] The present application provides a deep space satellite on-orbit autonomous inertial sun pointing method considering the limited rotation of satellite solar array and the demand for earth measurement and control. Figure 1 is a demand-oriented diagram of the autonomous inertial sun pointing algorithm. As shown in Figure 1 When the solar array drive assembly (SADA) of a complex orbit satellite such as a lunar satellite has an abnormal condition causing the solar array to point in a direction other than the designed expected direction, it is necessary to be able to obtain the best satellite body frame sun vector pointing (Sbm) according to the solar array normal pointing, energy demand and measurement and control demand, and to maintain this pointing in the complex deep space orbit motion of the satellite. Figure 1

[0041] Before describing the deep space satellite on-orbit autonomous inertial sun pointing method of the present application, the common coordinate systems of the satellite need to be introduced:

[0042] Inertial frame, the inertial frame is an absolute reference frame in Newtonian mechanics, which has no acceleration and no rotation, and is used to describe the absolute reference frame of the satellite orbit and attitude.

[0043] Body frame, the coordinate system fixed to the satellite body, the origin is usually at the satellite center of mass or the design reference point, the coordinate axes are defined according to aerospace conventions: the X-axis is along the flight direction of the satellite or the main structure axis, the Y-axis is perpendicular to the X-axis pointing to the side wing of the satellite (such as the solar wing deployment direction), and the Z-axis completes the orthogonal system according to the right-hand rule (usually pointing to the ground or the sun).

[0044] Figure 2 is a flowchart of the deep space satellite on-orbit autonomous inertial sun pointing method of an embodiment of the present application. As shown in Figure 2 The deep space satellite on-orbit autonomous inertial sun pointing method 200 includes: ​

[0045] Step S21: Obtain the solar target vector of this system. The solar target vector of this system is the solar vector of this system when the solar panel current value is at its maximum.

[0046] Step S22: Calculate the angle between the solar target vector of this system and the positive Z-axis of this system, and denote it as the first angle:

[0047] Step S23: Calculate the projection of the positive Z-axis of the local system onto the inertial frame based on the first included angle, the solar vector of the inertial frame, and the Earth vector of the inertial frame, and denote it as the first candidate pointing vector;

[0048] Step S24: Calculate the angle between the first candidate pointing vector and the Earth vector in the inertial frame, and denot it as the second angle;

[0049] Step S25: Determine the target pointing vector and target deflection angle of the satellite rotation based on the second included angle and the beam range of the telemetry and control antenna.

[0050] The following provides a detailed explanation of steps S21 to S25:

[0051] In step S21, the solar target vector Sbm of this system is obtained. The solar target vector Sbm of this system is a calculated quantity or a known quantity.

[0052] In step S22, according to aerospace conventions, the positive Z-axis of this system points towards the solar surface. The angle between the solar target vector Sbm of this system and the positive Z-axis Zb of this system is the first angle. .

[0053] In step S23, the inertial frame solar vector Si, the local system solar target vector Sbm, and the local system positive Z-axis Zb are all calculated or known quantities. However, the projection Zi of the local system positive Z-axis onto the inertial frame needs to be calculated in reverse based on the angle between the local system solar target vector Sbm and the local system +Z-axis Zb.

[0054] Figure 3 yes Figure 2 A flowchart of an embodiment of step S23. (See attached flowchart.) Figure 3 As shown, step S23 includes:

[0055] Step S231: Establish a reference coordinate system based on the solar vector and the Earth vector in the inertial frame.

[0056] In one embodiment, the reference coordinate system uses the inertial frame solar vector as the X-axis, the normal of the plane formed by the inertial frame solar vector and the inertial frame Earth vector as the Z-axis, and the Y-axis is perpendicular to the X-axis and Z-axis.

[0057] Figure 4 This is a schematic diagram of a reference coordinate system according to an embodiment of this application. For example... Figure 4As shown, the reference coordinate system takes the inertial system sun vector Si as the X axis, the inertial system sun vector Si and the inertial system earth vector Ei form the SiEi plane, and the cross product term is the normal of the SiEi plane, and the double cross product term is taken as the Z axis, and the triple cross product term is taken as the Y axis. The three coordinate axes establish the orthogonal coordinate system XYZ, and the angle between the inertial system sun vector Si and the projection Zi of the positive Z axis of the satellite body system in the inertial system is . And the conversion matrix from the inertial system to the reference coordinate system designed in the present application is obtained .

[0058] Step S232: Rotate the X axis in the reference coordinate system around the Z axis by the first included angle to obtain the projection of the first candidate pointing vector in the reference coordinate system, denoted as the reference system Z axis vector;

[0059] The angle between the inertial system sun vector Si and the projection Zi of the positive Z axis of the satellite body system in the inertial system is , the angle between the target sun vector Sbm of the satellite body system and the Zb axis of the satellite body system is also , and the angle between the target sun vector Sbm of the satellite body system and the positive Z axis Zb of the satellite body system is also .

[0060] Rotate the projection Xr of the inertial system sun vector Si in the reference system (i.e. the inertial system sun vector Si itself) clockwise around the longitudinal axis by , to obtain the projection Zr of the projection Zi of the Z axis of the satellite body system in the inertial system in the reference coordinate system, and the rotation matrix is ,

[0061]

[0062] Step S233: Calculate the first candidate pointing vector according to the reference system Z axis vector and the conversion matrix from the inertial system to the reference coordinate system.

[0063] Further, the temporary value of the projection Zi of the Z axis of the satellite body system in the inertial system is obtained, denoted as the first candidate pointing vector.

[0064]

[0065] wherein, is the first candidate pointing vector, is the conversion matrix from the inertial system to the reference coordinate system, is the reference system Z axis vector.

[0066] In steps S24-S25, as Figure 1As shown, the angle γ between the axis of the satellite telemetry and control antenna and the geocentric vector is set to not exceed 70°. Therefore, the antenna beam range is [0, 70°] and [110°, 180°]. To meet the requirements of Earth telemetry and control, when the satellite's pointing direction deviates from this angle, compensation angle and compensation polarity calculations are required. The calculation of the compensation angle polarity is problematic because there are many complex geometric relationships between the satellite pointing axis vector, the geocentric vector, the satellite rotation direction, and the satellite rotation angle value. Therefore, the conventional method of obtaining calculation results through logical judgment via traversal is difficult to reliably express and implement. This application achieves accurate determination of compensation polarity through a rotational trial-and-error algorithm.

[0067] Find the first candidate pointing vector and the Earth's vector in the inertial frame The included angle is , .

[0068] if Then, communication with the ground should be conducted via the telemetry and control antenna on the +Z plane, according to... The logic for further determining whether compensation is needed is as follows:

[0069] 1) It uses a +Z plane telemetry and control antenna for ground communication, which does not require angle compensation and reduces target deflection. The projection vector of the Z-axis in the inertial frame is: ;

[0070] 2) For ground communication using a +Z-plane telemetry and control antenna, angle compensation is required. The compensation angle and polarity are determined through a rotational trial-and-error algorithm as follows:

[0071] Forward Then, calculate the new temporary inertial frame vector. :

[0072] Rotation matrix ,

[0073]

[0074] According to the projection vector of the inertial frame and the Earth's vector in the inertial frame Calculate the included angle ,

[0075] like , , indicates that the rotation direction is reversed, and Flag indicates the rotation direction of the satellite;

[0076] like , , indicates the rotation direction is positive rotation;

[0077] the target deflection angle .

[0078] If , the communication with the ground should be carried out by the -Z surface tracking and control antenna, and according to the value, further judgment is made as to whether compensation is needed or not, and the logic is as follows:

[0079] 3) , the -Z surface tracking and control antenna is used for the communication with the ground, and no angle compensation is needed, and the target deflection angle , the Z-axis projection vector in the inertial system is: ;

[0080] 4) , the -Z surface tracking and control antenna is used for the communication with the ground, and angle compensation is needed. The compensation angle and the compensation polarity mode are determined through a rotation trial algorithm, and the logic is as follows:

[0081] positive rotation , a new temporary inertial system vector

[0082] rotation matrix ,

[0083]

[0084] According to the inertial system projection vector and the inertial system earth vector , the included angle ,

[0085] If , , indicates the rotation direction is reverse rotation, and Flag indicates the rotation direction of the satellite;

[0086] If , , indicates the rotation direction is positive rotation;

[0087] the target deflection angle .

[0088] If compensation to the ground is needed (corresponding to the above-mentioned cases 2) and 4) ), the Z-axis projection vector in the inertial system is the vector after rotating a compensation angle:

[0089]

[0090]

[0091] wherein, is the target pointing vector, is a first candidate pointing vector, is a target offset angle. The target pointing vector in the present application is the final determined Z-axis projection vector in the inertial system.

[0092] The present application determines the satellite attitude based on the inertial system sun vector and the inertial system earth vector. The double vector attitude determination needs to pay attention to the calculation problem caused by the small angle between the two vectors in engineering practice. The deep space satellite on-orbit autonomous inertial sun pointing method of the present application also includes:

[0093] The angle between the inertial system sun vector and the inertial system earth vector is calculated, denoted as a fifth angle. It is judged whether the value of the fifth angle is less than a first threshold angle or greater than a second threshold angle. If so, the satellite maintains the current target pointing vector and target offset angle, wherein the sum of the first threshold angle and the second threshold angle is 180°, and the first threshold angle is less than or equal to 10°.

[0094] In an embodiment, the first threshold angle is set to 10°, and the second threshold angle is 170°. The angle between the inertial system sun vector and the inertial system earth vector is When <10° or >170°, it is determined that the angle between the two vectors is small, and the current inertial pointing and offset angle are maintained.

[0095] On the one hand, the present application realizes that the satellite body system arbitrary vector points to the sun, that is, the specific vector direction on the satellite is consistent with the direction pointing to the sun, and ensures the energy supply of the satellite when the sailboard SADA is not designed to point in the expected direction. Figure 5 is a satellite sun pointing simulation result diagram of an embodiment of the present application. As shown in Figure 5 , the body system sun target vector is [1, -1, 1]. After 1000 seconds, the X-axis of the satellite body system specified vector tends to 0.6, the Y-axis tends to -0.6, and the Z-axis tends to 0.6. Thus, the direction of the satellite body system specified vector is consistent with that of the body system sun target vector, and the satellite body system arbitrary vector points to the sun is realized.

[0096] On the other hand, the present application realizes the compensation control considering the earth measurement and control angle, and ensures the measurement and control demand of the satellite in the complex deep space orbit. Figure 6 is a simulation diagram of the angle between the body system positive Z-axis and the earth center vector of an embodiment of the present application. As shown in Figure 6 , between 200 seconds and 250 seconds, the angle between the body system positive Z-axis and the earth center vector is greater than 70°, and thus the measurement and control cannot realize communication. Under the compensation mechanism of the present application, after 250 seconds, the angle between the body system positive Z-axis and the earth center vector starts to decrease, and after 400 seconds, the angle between the body system positive Z-axis and the earth center vector is maintained at about 70°, and the measurement and control communication can be realized.

[0097] This application proposes an attitude control method in inertial space that enables autonomous solar orientation with arbitrary vectors of the satellite's own system while also taking into account telemetry and control requirements. The method includes algorithms for inverse calculation of the Z-axis vector of the inertial system, compensation angle for telemetry and control angles, and polarity determination. It achieves energy and telemetry and control coordination and support under complex deep space orbit conditions where the solar panel normal does not point as designed.

[0098] This application also provides a computing device. Figure 7 This is a system block diagram of a computing device according to an embodiment of this application. (Reference) Figure 7 As shown, the computing device 700 may include an internal communication bus 701, a processor 702, a read-only memory (ROM) 703, a random access memory (RAM) 704, and a communication port 705. When applied to a personal computer, the computing device 700 may also include a hard disk 706. The internal communication bus 701 enables data communication between components of the computing device 700. The processor 702 can make judgments and issue prompts. In some embodiments, the processor 702 may consist of one or more processors. The communication port 705 enables data communication between the computing device 700 and external devices. In some embodiments, the computing device 700 can send and receive information and data from a network through the communication port 705. The computing device 700 may also include different forms of program storage units and data storage units, such as the hard disk 706, the read-only memory (ROM) 703, and the random access memory (RAM) 704, capable of storing various data files used for computer processing and / or communication, as well as possible program instructions executed by the processor 702. The processor executes these instructions to implement the main part of the method. The results processed by the processor are transmitted to the user device through the communication port and displayed on the user interface.

[0099] The above-described operation method can be implemented as a computer program, stored in hard disk 706, and loaded into processor 702 for execution, in order to implement the deep space satellite on-orbit autonomous inertial sun alignment method of this application.

[0100] This application also includes a computer-readable medium storing computer program code that, when executed by a processor, implements the aforementioned deep-space satellite on-orbit autonomous inertial sun alignment method.

[0101] The deep space satellite on-orbit autonomous inertial sun pointing method, when implemented as a computer program, can also be stored in a computer-readable storage medium as an article of manufacture. For example, the computer-readable storage medium can include, but is not limited to, magnetic storage devices (e.g., hard disk; floppy disk); optical disk devices (e.g., compact disk (CD), digital versatile disk (DVD)); smart cards; and flash memory devices (e.g., electrically erasable programmable read only memory (EPROM), card, stick, key drive). Additionally, the various storage mediums described herein can represent one or more devices and / or other machine-readable media for storing information. The term "machine-readable medium" can include, without being limited to, wireless channels and various other media (and / or storage media) that are capable of storing, containing, and / or carrying code and / or instructions and / or data.

[0102] Flow diagrams have been used herein to illustrate the operations performed by systems in accordance with embodiments of the present application. It should be understood that the operations previously or hereafter described do not necessarily have to be performed in the precise order described. Rather, various steps can be handled in reverse order, or simultaneously. Also, other operations can be added to, or removed from, these processes, or one or more steps can be omitted.

[0103] Having described above the basic concepts, it is obvious that the above-described disclosure of the invention is merely meant to be exemplary and not restrictive of the present application. Various modifications, improvements, and alterations of the present application can be made by those skilled in the art, although not explicitly described herein. Such modifications, improvements, and alterations are suggested by the present application and are still within the spirit and scope of the exemplary embodiments of the present application.

[0104] Also, specific terminology used herein is to be taken only in a descriptive sense, and not as a limiting factor. As used herein, the expression "one embodiment", "an embodiment", and / or "some embodiments" means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the application. Therefore, it is expressly intended that the descriptions of "one embodiment", "an embodiment", and / or "some embodiments" should not be interpreted as applying only to a single embodiment or to a single aspect of an embodiment. Furthermore, various features of the one or more embodiments of the application can be combined with each other, as appropriate, and can be removed therefrom where appropriate.

[0105] Aspects of the application can be implemented in hardware, firmware, software, or any combination thereof. The hardware and software components can be implemented within one or more application specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DSPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, micro-controllers, microprocessors, other electronic units designed to perform the functions described herein, or a combination thereof. The processor can be implemented as one or more general processors from any processor family or as processor(s) implementing more than one family of processors. Also, the various features of the application can be embodied as or performed by one or more computer-readable media containing computer program code, which can be executed by a processor to perform the techniques described herein. As used herein, the term "computer- readable medium" encompasses only a computer-readable medium that can contain, store, communicate, propagate, or transport the program for use by or in connection with the processor.

[0106] Computer readable media can include a propagated data signal with computer- readable program code embodied therein, for example, in baseband or as part of a carrier wave. Such a propagated signal can take any of a variety of forms, including, but not limited to, electro-magnetic, optical, or any suitable combination thereof. Computer readable media can be any media that can be accessed by a computer. Such computer-readable media can include computer storage media and communication media. Computer storage media can include volatile and non-volatile, removable and non-removable media implemented in any technology for the purpose of storage of information such as computer readable instructions, data structures, program modules or other data. Communication media typically embody computer readable instructions, data structures, program modules or other data in a modulated data signal such as a carrier wave or other transport mechanism and include any information delivery media. As an example, the foregoing can be employed taking the entity of a communications network, such as the Internet, to communicate computer-readable program code.

[0107] It should also be noted that, as used in the specification and the claims, the singular form "a", "an" and "the" include plural references unless the context clearly dictates otherwise. As used herein, the expression "and / or" includes combinations of all of the associated terms. As used herein, the expression "at least one of A and B" is intended to mean "A or B or both A and B." As used herein, the expression "at least one of A, B and C" is intended to mean "A or B or C or any combination of these, including A and B and C." As used herein, the expression "at least one of A, B, C and D" is intended to mean "A or B or C or D or any combination of these, including A and B and C and D."

[0108] As used in this application, the terms "comprises", "comprising", "includes", "including", or the like are used in the sense of "including but not limited to", and not in the sense of "consisting only of", unless specifically indicated otherwise. As used in this application, the term "coupled" means a direct or indirect connection between components, and can be electrical, mechanical, logical, or any combination thereof.

[0109] The foregoing is considered as illustrative only of the principles of the application. Further, since numerous modifications and changes will readily occur to those skilled in the art, it is not desired to limit the application to the exact construction and practice described. Accordingly, all suitable modifications and equivalents can be resorted to falling within the scope of the application.

[0110] In addition, it should be understood that any numerical value or range range specified in this application will include all values falling within such numerical value and range, as if such values were expressly written herein. A range includes each individual number falling within that range, inclusive of the endpoints of the range. Furthermore, it should be understood that the apparatus specifically described herein includes all possible combinations of two or more of the described components.

[0111] In some embodiments, numerical descriptors of ingredients, amounts of attributes are used. It should be understood that such numerical descriptors used in the description of embodiments, in some examples, are modified by the modifier "about," "approximately," or "generally." Unless otherwise stated, "about," "approximately," or "generally" indicate that the number can vary ±20%. Accordingly, numerical parameters in the description are approximations only, and in some embodiments, the numerical parameters are approximations that can vary depending on the desired properties sought to be obtained by the individual embodiment. In some embodiments, numerical parameters are determined by the use of standard techniques. Although the numerical ranges and parameters setting forth the broad scope of the application are approximations, the numerical values set forth in the specific examples are reported as precisely as possible. Any numerical value, however, can contain certain errors necessarily resulting from the standard deviation found in their respective testing measurements.

[0112] While the application has been described with reference to the currently preferred embodiments and examples, the application is not to be limited by such description. Various modifications and changes can be made thereto by those skilled in the art without departing from the spirit of the application, and it is understood that all such modifications and changes are intended to be included within the scope of the application.

Claims

1. A deep space satellite on-orbit autonomous inertial sun pointing method, applicable to a satellite whose sail panel normal points to a non-design direction, characterized in that, include: Obtain the solar target vector of this system, which is the solar target vector of this system when the solar panel current value is at its maximum; Calculate the angle between the solar target vector of this system and the positive Z-axis of this system, and denote it as the first angle: The projection of the positive Z-axis of the system onto the inertial frame is calculated based on the first included angle, the solar vector of the inertial frame, and the Earth vector of the inertial frame, and is denoted as the first candidate pointing vector; Calculate the angle between the first candidate pointing vector and the Earth vector of the inertial frame, and denote it as the second angle; The target pointing vector and target deflection angle of the satellite rotation are determined based on the second included angle and the beam range of the telemetry and control antenna.

2. The deep space satellite on-orbit autonomous inertia sun pointing method according to claim 1, characterized in that, The projection of the positive Z-axis of the system onto the inertial frame, calculated based on the first included angle, the solar vector of the inertial frame, and the Earth vector of the inertial frame, includes: A reference coordinate system is established based on the solar vector of the inertial frame and the Earth vector of the inertial frame; The first candidate pointing vector is projected onto the reference coordinate system by rotating the X-axis around the Z-axis by the first included angle, and is denoted as the Z-axis vector of the reference system. The first candidate pointing vector is calculated based on the Z-axis vector of the reference system and the transformation matrix from the inertial frame to the reference coordinate system.

3. The deep space satellite on-orbit autonomous inertia sun pointing method according to claim 2, characterized in that, Establishing a reference coordinate system based on the solar vector and the Earth vector of the inertial frame includes: The inertial frame solar vector is taken as the X-axis, the normal of the plane formed by the inertial frame solar vector and the inertial frame Earth vector is taken as the Z-axis, and the Y-axis is perpendicular to the X-axis and the Z-axis.

4. The deep space satellite on-orbit autonomous inertia sun pointing method according to claim 3, characterized in that, The Z-axis vector of the reference frame is calculated using the following formula: wherein, is the reference system Z axis vector, is the first included angle, is the inertial system sun vector.

5. The deep space satellite on-orbit autonomous inertia sun pointing method according to claim 2, characterized in that, The first candidate pointing vector is calculated using the following formula: wherein, is the first candidate pointing vector, is a transformation matrix from the inertial frame to the reference frame, is the reference frame Z-axis vector.

6. The method for autonomous inertial sun alignment of a deep-space satellite in orbit as described in claim 5, characterized in that, The target pointing vector and target deflection angle of the satellite rotation are determined based on the second included angle and the beam range of the telemetry and control antenna, including: Let the second included angle be denoted as The target pointing vector is denoted as The target deflection angle is denoted as The beam range of the telemetry and control antenna is and The target pointing vector and target deflection angle are calculated in the following manner: when It uses a +Z-plane telemetry and control antenna for ground communication, eliminating the need for angle compensation. , ; when It uses a Z-plane telemetry and control antenna for ground communication, eliminating the need for angle compensation. , .

7. The method for autonomous inertial sun alignment of a deep-space satellite in orbit as described in claim 6, characterized in that, Also includes: when or At this time, it is necessary to compensate for the geocentric angle, determine the satellite rotation direction and rotation angle through a rotational trial algorithm, and determine the target deflection angle based on the rotation direction and rotation angle; The target pointing vector for satellite rotation is calculated based on the target deflection angle and the first candidate pointing vector.

8. The method for autonomous inertial sun alignment of a deep-space satellite in orbit as described in claim 7, characterized in that, when The target deflection angle is determined as follows: Control the satellite to rotate forward by a preset angle, calculate the projection of the positive Z-axis of the system onto the inertial frame after rotation, and denote it as the second candidate pointing vector; Calculate the angle between the second candidate pointing vector and the Earth vector of the inertial frame, and denote it as the third angle; When the third included angle is less than the second included angle, the rotation direction is clockwise; when the third included angle is greater than the second included angle, the rotation direction is counterclockwise, and the rotation angle is... The target deflection angle Flag indicates the direction of satellite rotation. A value of 1 indicates forward rotation, and a value of -1 indicates reverse rotation.

9. The method for autonomous inertial sun alignment of a deep-space satellite in orbit as described in claim 7, characterized in that, when The target deflection angle is determined as follows: Control the satellite to rotate forward by a preset angle, calculate the projection of the positive Z-axis of the system onto the inertial frame after rotation, and denote it as the third candidate pointing vector; Calculate the angle between the third candidate pointing vector and the Earth vector of the inertial frame, and denote it as the fourth angle; When the fourth included angle is less than the second included angle, the rotation direction is clockwise; when the fourth included angle is greater than the second included angle, the rotation direction is counterclockwise, and the rotation angle is... The target deflection angle Flag indicates the direction of satellite rotation. A value of 1 indicates forward rotation, and a value of -1 indicates reverse rotation.

10. The method for autonomous inertial sun alignment of a deep-space satellite in orbit as described in claim 7, characterized in that, The target pointing vector of the satellite rotation is calculated using the following formula: in, The target pointing vector, For the first candidate pointing vector, The target deflection angle is denoted as .

11. The method for autonomous inertial sun alignment of a deep-space satellite in orbit as described in claim 1, characterized in that, Also includes: Calculate the angle between the solar vector of the inertial frame and the Earth vector of the inertial frame, and denote it as the fifth angle; Determine whether the value of the fifth included angle is less than the first threshold angle or greater than the second threshold angle. If so, the satellite maintains the current target pointing vector and target deflection angle, wherein the sum of the first threshold angle and the second threshold angle is 180°, and the first threshold angle is less than or equal to 10°.

12. A computing device, characterized in that, include: Memory is used to store instructions executed by the processor; A processor for executing the instructions to implement the deep space satellite on-orbit autonomous inertial sun alignment method as described in any one of claims 1 to 11.

13. A computer-readable medium storing computer program code that, when executed by a processor, implements the deep space satellite on-orbit autonomous inertial sun alignment method as described in any one of claims 1 to 11.

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