Deep space satellite in-orbit autonomous inertia sun alignment method, calculation device and readable medium

Abstract

The invention provides a deep space satellite in-orbit autonomous inertia sun alignment method, a calculation device and a readable medium. The method is suitable for a satellite with a sailboard normal pointing to a non-design direction, and comprises the following steps: obtaining a sun target vector of a system, the sun target vector of the system being a sun vector of the system when a sailboard current value is maximum; calculating an included angle between the sun target vector of the system and the positive Z axis of the system, and recording the included angle as a first included angle; calculating a projection of the positive Z axis of the system in the inertial system according to the first included angle, the sun vector of the inertial system and the earth vector of the inertial system, and recording the projection as a first candidate pointing vector; calculating an included angle between the first candidate pointing vector and the inertial 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 the measurement and control antenna beam range. According to the invention, attitude control which autonomously faces the sun by any vector of the satellite body system and gives consideration to measurement and control requirements in an inertial space is realized.

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

[0011] Optionally, the first candidate pointing vector is calculated by the following formula:

[0012] 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.

[0013] Optionally, determining a target pointing vector and a target declination angle of satellite rotation according to the second included angle and a range of a TT&C antenna beam comprises: Let the second included angle be , the target pointing vector be , the target declination angle be , and the range of the TT&C antenna beam be and The target pointing vector and the target declination angle are calculated by the following ways: When , a +Z plane TT&C antenna is used for earth communication, and no angle compensation is needed, , ; When , a -Z plane TT&C antenna is used for earth communication, and no angle compensation is needed, , .

[0014] 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 and 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 and angle of rotation.

[0015] 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 and 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 and 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 and 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 and 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 and 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 and 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.

[0016] 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 and 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 and 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 and 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 and 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 and 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 and 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.

[0017] Optionally, the target pointing vector of the satellite rotation is calculated by the following formula:

[0018]

[0019] wherein, is the target pointing vector, is the first candidate pointing vector, is the target deflection angle.

[0020] Optionally, the method further comprises: 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°.

[0021] 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.

[0022] 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.

[0023] Compared with the prior art, the application has the following advantages: The on-orbit autonomous inertial sun pointing method of the deep space satellite, the computing device and the computer readable medium provided by the application can point any vector of the satellite body to the sun, i.e., make the specific vector direction of the satellite point to the sun, so as to ensure the energy supply of the satellite when the sailboard SADA is not designed to point in the expected direction; and the application can also compensate the control of the angle of the earth observation and control, so as to ensure the observation and control requirements of the satellite in the complex deep space orbit. BRIEF DESCRIPTION OF DRAWINGS

[0024] 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 show embodiments of the application, and together with the specification, serve to explain the principles of the application. In the drawings: Figure 1 is a demand-oriented diagram of the autonomous inertial sun pointing algorithm.

[0025] Figure 2 is a flowchart of the on-orbit autonomous inertial sun pointing method of the deep space satellite according to an embodiment of the application.

[0026] Figure 3 is Figure 2 is a flowchart of step S23 in the method of claim 1.

[0027] Figure 4 is a schematic diagram of the reference coordinate system according to an embodiment of the application.

[0028] Figure 5 is a satellite sun pointing simulation result diagram according to an embodiment of the application.

[0029] Figure 6is a simulation diagram of an angle between a Z-axis of a body frame of an embodiment of the present application and an earth center vector.

[0030] Figure 7 is a system block diagram of a computing device of an embodiment of the present application. DETAILED DESCRIPTION

[0031] 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 numerals in the drawings represent the same structure or operation.

[0032] The present application provides a deep space satellite on-orbit autonomous inertia sun pointing method considering satellite solar array rotation restriction and earth measurement and control requirements. Figure 1 is a demand-oriented diagram of autonomous inertia sun pointing algorithm. As shown in Figure 1 When the solar array drive mechanism (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 design expected direction, it is necessary to obtain the best satellite body frame sun vector pointing (Sbm) according to the solar array normal pointing, energy requirement and measurement and control requirement, and to maintain this pointing in the complex deep space orbit motion of the satellite. Figure 1

[0033] Before describing the deep space satellite on-orbit autonomous inertia sun pointing method of the present application, the common coordinate systems of the satellite need to be introduced: Inertial frame, the inertial frame is an absolute reference frame in Newtonian mechanics, with no acceleration and no rotation, used to describe the absolute reference frame of satellite orbit and attitude.

[0034] Body frame, a coordinate system fixed to the satellite body, with the origin usually at the satellite center of mass or a design reference point, and the coordinate axes defined according to aerospace conventions: the X-axis along the satellite flight direction or the main structure axis, the Y-axis perpendicular to the X-axis pointing to the satellite wing (such as the solar wing deployment direction), and the Z-axis completing the orthogonal system according to the right-hand rule (usually pointing to the ground or the sun).

[0035] Figure 2 is a flowchart of a deep space satellite on-orbit autonomous inertia sun pointing method of an embodiment of the present application. As shown in Figure 2 The deep space satellite on-orbit autonomous inertia sun pointing method 200 includes: Step S21: obtaining a body frame sun target vector, the body frame sun target vector being a body frame sun vector when the solar array current value is maximum;​ Step S22: Calculate the angle between the solar target vector of the system and the positive Z axis of the system, which is recorded as the first angle: Step S23: Calculate the projection of the positive Z axis of the system in the inertial system based on the first included angle, the inertial system sun vector, and the inertial system earth vector, and record it as the first candidate pointing vector; Step S24: Calculate the angle between the first candidate pointing vector and the inertial system earth vector, which is recorded as the second angle; Step S25: Determine the target pointing vector and target deflection angle of the satellite rotation according to the second angle and the beam range of the measurement and control antenna.

[0036] The following describes steps S21 to S25 in detail: In step S21, the solar target vector Sbm of the system is obtained, and the solar target vector Sbm of the system is a calculated quantity or a known quantity.

[0037] In step S22, according to the definition of aerospace convention, the positive Z axis of the system points to the solar side. The angle between the solar target vector Sbm of the system and the positive Z axis Zb of the system is the first angle .

[0038] In step S23, the inertial system solar vector Si, the system solar target vector Sbm, and the system positive Z axis Zb are all calculated or known quantities, but the projection Zi of the system positive Z axis in the inertial system needs to be reversely positioned based on the angle between the system solar target vector Sbm and the system +Z axis Zb.

[0039] Figure 3 yes Figure 2 Flowchart of step S23 in the embodiment. Figure 3 As shown, step S23 includes: Step S231: Establish a reference coordinate system according to the inertial system sun vector and the inertial system earth vector.

[0040] In one embodiment, the reference coordinate system uses the inertial frame sun vector as the X axis, the normal of the plane formed by the inertial frame sun vector and the inertial frame earth vector as the Z axis, and the Y axis is perpendicular to the X axis and the Z axis.

[0041] Figure 4 Schematic diagram of a reference coordinate system according to an embodiment of the present application. Figure 4 As 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. The cross product term is the normal of the SiEi plane, and the cross product As the Z axis, the double cross product The three coordinate axes establish an orthogonal coordinate system XYZ, and the angle between the inertial system sun vector Si and the projection Zi of the satellite body system positive Z axis in the inertial system is The conversion matrix from the inertial system to the reference coordinate system of the application is obtained .

[0042] 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; The angle between the inertial system sun vector Si and the projection Zi of the satellite body system positive Z axis in the inertial system is The angle between the target body system sun vector Sbm and the satellite body system Zb axis is also The angle between the target body system sun vector Sbm and the satellite body system positive Z axis Zb is also .

[0043] 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 Obtain the projection Zr of the projection Zi of the satellite body system Z axis in the inertial system in the reference coordinate system, and the rotation matrix is ,

[0044] 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.

[0045] Further obtain the temporary value of the projection Zi of the satellite body system Z axis in the inertial system, denoted as the first candidate pointing vector.

[0046]

[0047] 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.

[0048] In steps S24-S25, as Figure 1As shown, the angle γ between the axis of the satellite tracking and control antenna and the geocentric vector is set to no more than 70°, then the tracking and control antenna beam range is [0, 70°] and [110°, 180°]. In order to meet the ground tracking and control requirements, when the satellite pointing deviates from this angle, it is necessary to calculate the compensation angle and compensation polarity. The calculation problem of the compensation angle polarity is difficult to reliably express and implement by performing logical judgment through traversal and then obtaining the calculation result due to the complex geometric relationships between the satellite pointing axis vector, the geocentric vector, the satellite rotation direction and the satellite rotation angle value. This application uses a rotation heuristic algorithm to achieve accurate judgment of the compensation polarity.

[0049] Obtain the first candidate pointing vector and the inertial Earth vector The angle is , .

[0050] if , then the +Z plane TT&C antenna should be used for ground communication, according to The logic for further determining whether compensation is required is as follows: 1) , using the +Z plane measurement and control antenna for ground communication, no angle compensation is required, the target deflection , the projection vector of the Z axis in the inertial system is: ; 2) , using the +Z plane tracking antenna for ground communication, angle compensation is required. The compensation angle and compensation polarity are determined by the rotation heuristic algorithm as follows: Forward Then, calculate the new temporary inertial system vector : Rotation Matrix ,

[0051] According to the projection vector of the inertial system and the inertial Earth vector Calculate the angle , like , , indicating that the rotation direction is reverse, Flag indicates the rotation direction of the satellite; like , , indicating that the direction of rotation is forward; The target deflection angle .

[0052] if If so, the measurement and control antenna of the -Z plane should be used for the communication with the earth, and the angle compensation is not needed according to The logic for further determining whether compensation is needed is as follows: 3) The measurement and control antenna of the -Z plane is used for the communication with the earth, and the angle compensation is not needed. The projection vector of the Z axis in the inertial system is: ; 4) The measurement and control antenna of the -Z plane is used for the communication with the earth, and the angle compensation is needed. The compensation angle and the compensation polarity mode are determined by the rotation trial algorithm as follows: Positive rotation Then, the new temporary inertial system vector is calculated

[0053] Rotation matrix ,

[0054] The included angle is calculated according to the projection vector of the Z axis in the inertial system and the earth vector in the inertial system , , If , , the rotation direction is reverse, and Flag represents the rotation direction of the satellite. If , , the rotation direction is positive rotation. The target deflection angle .

[0055] If the compensation to the earth is needed (corresponding to the cases 2 and 4 in the foregoing), the projection vector of the Z axis in the inertial system is the vector after rotating a compensation angle:

[0056]

[0057] Among them, is the target pointing vector, is the first candidate pointing vector, is the target deflection angle. In the present application, the target pointing vector is the final determined projection vector of the Z axis in the inertial system.

[0058] The attitude of the satellite is determined based on the inertial system sun vector and the inertial system earth vector in the present application. The double vector attitude determination needs to pay attention to the calculation problem caused by the small included angle in the engineering practice. The deep space satellite on-orbit autonomous inertial sun-seeking method in the present application further includes: The angle between the inertial frame sun vector and the inertial frame earth vector is calculated and recorded as the fifth angle. A determination is made as to 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 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°.

[0059] In one 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 When it is greater than 170°, the double vector angle is determined to be small, and the current inertial direction and deflection angle are maintained.

[0060] On the one hand, the present application realizes that any vector of the satellite system can point to the sun, that is, the specific vector direction on the satellite is consistent with the direction pointing to the sun, ensuring the energy supply of the satellite when the sailboard SADA is not pointed in the expected direction. Figure 5 This is a diagram showing the satellite-sun simulation results of an embodiment of the present application. Figure 5 As shown, the system's solar target vector is [1, -1, 1]. After 1000 seconds, the satellite's system's designated vector approaches 0.6 on the X axis, -0.6 on the Y axis, and 0.6 on the Z axis. Thus, the satellite's system's designated vector and the system's solar target vector are aligned, allowing any vector in the satellite system to point toward the sun.

[0061] On the other hand, this application realizes compensation control that takes into account the ground measurement and control angle, ensuring the measurement and control needs of the satellite in complex deep space orbits. Figure 6 : This is a schematic diagram of the angle simulation between the positive Z axis and the geocentric vector of the system in one embodiment of the present application. Figure 6 As shown, between 200 and 250 seconds, the angle between the system's positive Z axis and the Earth's center vector is greater than 70°, making measurement and control impossible. With the compensation mechanism of this application, after 250 seconds, the angle between the system's positive Z axis and the Earth's center vector begins to decrease. After 400 seconds, the angle between the system's positive Z axis and the Earth's center vector remains around 70°, allowing measurement and control communication.

[0062] This application proposes a method for realizing autonomous attitude control of the satellite system towards the sun with arbitrary vectors in inertial space while taking into account the measurement and control needs. It includes algorithms such as reverse calculation of the Z-axis vector of the inertial system, compensation angle and polarity judgment for the measurement and control angle, and realizes the coordinated support of energy and measurement and control for the non-designed direction of the sailboard normal under complex deep space orbit conditions.

[0063] The present application also provides a computing device. Figure 7 is a system block diagram of a computing device according to an embodiment of the present application. Figure 7As shown, the computing device 700 can 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 on a personal computer, the computing device 700 can also include a hard disk 706. The internal communication bus 701 can enable data communication among the components of the computing device 700. The processor 702 can make decisions and issue prompts. In some embodiments, the processor 702 can be composed of one or more processors. The communication port 705 can enable data communication of the computing device 700 with the outside. 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 can 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, which can store various data files used by the computer processing and / or communication, and possible program instructions executed by the processor 702. The processor executes these instructions to implement the main part of the method. The results of the processor processing are transmitted to the user equipment through the communication port, and displayed on the user interface.

[0064] The above-mentioned operation method can be implemented as a computer program, saved in the hard disk 706, and loaded into the processor 702 for execution, to implement the on-orbit autonomous inertial sun pointing method of the deep space satellite of the present application.

[0065] The present application also includes 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 described above.

[0066] When the on-orbit autonomous inertial sun pointing method of the deep space satellite is implemented as a computer program, it 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, a magnetic storage device (e.g., a hard disk, a floppy disk, a magnetic strip), an optical disk (e.g., a compact disk (CD), a digital versatile disk (DVD)), a smart card, and a flash memory device (e.g., an electrically erasable programmable read only memory (EPROM), a card, a stick, a key drive). In addition, the various storage media described herein can represent one or more devices and / or other machine readable media for storing information. The term "machine readable medium" can include, but is not limited to, a wireless channel and various other media (and / or storage media) that can store, contain, and / or carry code and / or instructions and / or data.

[0067] Flow diagrams have been used herein to illustrate the operation of systems in accordance with embodiments of the present application. It will be appreciated that the acts need not be performed in the precise order shown. Rather, various steps can be handled in different order or simultaneously. Moreover, or alternatively, other acts can be added or removed.

[0068] Having described 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 changes can be made by those skilled in the art to the present application, although not explicitly described herein. Such modifications, improvements and changes are suggested by the present application and are still within the spirit and scope of the exemplary embodiments of the present application.

[0069] Also, specific terminology used when describing the embodiments of the present application should be interpreted only as descriptive, and not as a limitation on the present application. As used herein, the expression "one embodiment", "an embodiment", and / or "some embodiments" means a certain feature, structure, or characteristic that is associated with at least one embodiment of the present application. Therefore, it is to be understood that the phraseology "in one embodiment" or "in an embodiment" or "in some embodiments" appearing in various places throughout this specification does not necessarily refer to the same embodiment. Furthermore, some of the features, structures, or characteristics of the one or more embodiments of the present application can be combined in any suitable manner.

[0070] Some aspects of the present application can be performed entirely in hardware, entirely in software (including firmware, resident software, microcode, etc.), or a combination of hardware and software. The hardware or software can be referred to as a "block", "module", "engine", "unit", "component", or "system". The processor can be one or more application specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DAPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors, or combinations thereof. Furthermore, aspects of the present application can be manifested as a computer product including a computer readable medium storing computer program code. The computer readable medium can include, but is not limited to, magnetic storage devices (e.g., hard disk; floppy disk; magnetic strips...), optical disks (e.g., compact disk (CD); digital versatile disk (DVD)...), smart cards, and flash devices (e.g., card, stick, key drive...).

[0071] Computer readable media can include a propagated data signal with a 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 combination thereof. Computer readable media can be any media that can be accessed by a computer. By way of example, and not limitation, such computer readable media can comprise RAM, ROM, EEPROM, CD-ROM or any combination thereof. The computer readable media can be transitory, such as a modulated data signal, including, but not limited to, carrier waves, or other transmission media and / or storage media. Alternatively, the computer readable media can be non-transitory, such as a semiconductor manufacturing mask, other hardware implementation, and / or any program code on a computer readable medium that is devoid of transitory signals per se.

[0072] It should also be noted that, as used in the specification and the appended claims, the singular forms "a," "an" and "the" include plural referents unless the context clearly dictates otherwise. Thus, for example, reference to "a component" or "the component" can include a plurality of such components, and so forth.

[0073] As indicated, unless otherwise stated, the terms "one" and "the" include plural referents. Generally, the terms "including," "includes," "containing," "contains," or "containing," when used in this disclosure, mean "including, but not limited to." Unless otherwise noted, the use of the term "about" with respect to a given value or a range of values means that the value or range of values will vary to some extent. When the term "about" is used in reference to a particular value, the exact value is also considered to be within the scope of the disclosure.

[0074] Unless otherwise noted, the relative arrangement of components and steps, the numerical expressions, and numerical values set forth in the various embodiments presented herein are not limiting. Rather, these embodiments are presented for illustrative purposes and the scope of the disclosure is not limited to the embodiments presented herein. Moreover, the dimensions and other physical characteristics relating to the embodiments disclosed herein are not necessarily drawn to scale. It is to be understood that other embodiments can be utilized, and structural, logical, and / or electrical changes can be made without departing from the scope of the present disclosure. It is intended, therefore, to be limited only by the scope of the appended claims. In the drawings, like reference numerals refer to like elements throughout the various figures and the specification. In the drawings, the terms "first," "second," "third," etc., are used to identify various elements, but these elements should not be construed as limited to these three elements. Rather, these terms are used merely as labels to facilitate explanation of the illustrated embodiments. In the drawings, the term "coupled" is used to indicate that two or more elements, modules, components, or the like cooperate or interact with each other to also have a functional relationship.

[0075] In addition, it should be noted that the use of "first", "second", and / or the like herein is merely to distinguish one element from another, but is not otherwise intended to limit the scope of the application unless otherwise stated. Moreover, although the terms "first", "second", etc. can be used herein to describe various elements or components, these elements should not be limited by these terms since such elements can be either first, second, third or otherwise according to some embodiments. The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting. It is also possible in the present disclosure that terms could be used whose meanings are consistent with terms used by those skilled in the art and thus, if not explicitly stated as having an unusual meaning, no unusual meaning should be conveyed by such terms. Moreover, although the terms "include", "have", "contain", and / or the like are used herein to describe and reference certain features, elements and / or components, these terms are not intended to be limiting. Rather, these terms are used to simply mean that something comprises at least the recited feature, element or component. As such, these terms are used in their broadest context, and are intended to be open-ended, permitting the description of something that has the recited feature, element or component, but also something that has additional features, elements or components.

[0076] Some embodiments use numerals to describe components, quantities of attributes. It should be understood that such numerals used in the description of embodiments are, in some examples, modified by the words "about", "approximately", or "generally". Unless otherwise stated, "about", "approximately", or "generally" indicates that the described numeral can vary ±20%. Accordingly, numerical parameters in the description are approximations, and as such, the approximations can vary depending upon the desired properties sought to be obtained by the individual embodiments. 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. The numerical values set forth in the specific examples are provided to be as precise as reasonably possible. However, some variations may

[0077] Although the present application has been described with reference to the current embodiments, persons having ordinary skill in the art will recognize that changes can be made in form and detail without departing from the spirit and the scope of the application. Therefore, the disclosed embodiments should be considered in all respects as illustrative and not restrictive, and all changes coming within the meaning and equivalency range of the appended claims are intended to be embraced therein.

Claims

1. A method for autonomous inertial tracking of a deep space satellite on orbit, suitable for satellites whose sailboard normal points in a non-designed direction, characterized by: include: Obtaining a solar target vector of the system, where the solar target vector of the system is the solar vector of the system when the current value of the sailboard is maximum; Calculate the angle between the solar target vector of the system and the positive Z axis of the system, which is recorded as the first angle: Calculate the projection of the positive Z axis of the system in the inertial system based on the first included angle, the inertial system sun vector, and the inertial system earth vector, and record it as a first candidate pointing vector; Calculating the angle between the first candidate pointing vector and the inertial system earth vector, which is recorded as a second angle; The target pointing vector and target deflection angle of the satellite rotation are determined according to the second angle and the beam range of the measurement and control antenna.

2. The deep space satellite on-orbit autonomous inertial solar tracking method according to claim 1, characterized in that: Calculating the projection of the positive Z axis of the system in the inertial system according to the first 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; Rotate the X-axis in the reference coordinate system around the Z-axis by the first angle to obtain a projection of the first candidate pointing vector in the reference coordinate system, which is recorded as the reference system Z-axis vector; The first candidate pointing vector is obtained by calculation based on the Z-axis vector of the reference system and the transformation matrix from the inertial system to the reference coordinate system.

3. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 2, characterized in that: Establishing a reference coordinate system according to the inertial system sun vector and the inertial system earth vector includes: The inertial system sun vector is used as the X-axis, the normal of the plane formed by the inertial system sun vector and the inertial system earth vector is used as the Z-axis, and the Y-axis is perpendicular to the X-axis and the Z-axis.

4. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 3, wherein: The Z-axis vector of the reference system is calculated by the following formula: in, is the reference system Z-axis vector, is the first angle, is the sun vector of the inertial system.

5. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 2, wherein: The first candidate pointing vector is calculated by the following formula: in, is the first candidate pointing vector, is the transformation matrix from the inertial system to the reference coordinate system, is the Z-axis vector of the reference system.

6. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 5, characterized in that: Determining the target pointing vector and target deflection angle of the satellite rotation according to the second angle and the beam range of the measurement and control antenna includes: The second angle is recorded as , the target pointing vector is recorded as , the target deflection angle is recorded as , the measurement and control antenna beam range is and , the target pointing vector and target deflection angle are calculated in the following way: when , using the +Z plane measurement and control antenna for ground communication, no angle compensation is required, , ; when , using the -Z plane measurement and control antenna for ground communication, no angle compensation is required, , .

7. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 6, characterized in that: Also includes: when or When the angle between the Earth and the Earth is centered is compensated, the satellite rotation direction and rotation angle are determined by a rotation heuristic algorithm, and the target deflection angle is determined according to the rotation direction and rotation angle; A target pointing vector of the satellite rotation is calculated according to the target deflection angle and the first candidate pointing vector.

8. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 7, wherein: when When , the target deflection angle is determined as follows: Control the satellite to rotate forward to a preset angle, calculate the projection of the positive Z axis of the system in the inertial system after rotation, and record it as the second candidate pointing vector; Calculate the angle between the second candidate pointing vector and the inertial system earth vector, and record it as a third angle; When the third angle is smaller than the second angle, the rotation direction is forward; when the third angle is larger than the second angle, the rotation direction is reverse, and the rotation angle is , the target deflection angle , Flag is the rotation direction of the satellite. A Flag value of 1 indicates that the rotation direction is forward, and a Flag value of -1 indicates that the rotation direction is reverse.

9. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 7, wherein: when When , the target deflection angle is determined as follows: Control the satellite to rotate forward to a preset angle, calculate the projection of the positive Z axis of the system in the inertial system after rotation, and record it as the third candidate pointing vector; Calculating the angle between the third candidate pointing vector and the inertial system earth vector, which is recorded as a fourth angle; When the fourth angle is smaller than the second angle, the rotation direction is forward; when the fourth angle is larger than the second angle, the rotation direction is reverse, and the rotation angle is , the target deflection angle , Flag is the rotation direction of the satellite. A Flag value of 1 indicates that the rotation direction is forward, and a Flag value of -1 indicates that the rotation direction is reverse.

10. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 7, wherein: The target pointing vector of the satellite rotation is calculated using the following formula: in, is the target pointing vector, is the first candidate pointing vector, is the target deflection angle.

11. The method for autonomous inertial tracking of a deep space satellite on orbit as claimed in claim 1, wherein: Also includes: Calculate the angle between the inertial system sun vector and the inertial system earth vector, and record it as the fifth angle; Determine 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 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: a memory for storing instructions to be executed by the processor; A processor, configured to execute the instructions to implement the on-orbit autonomous inertial solar tracking method for a deep space satellite as described in any one of claims 1 to 11.

13. A computer-readable medium storing computer program code, wherein when executed by a processor, the computer program code implements the on-orbit autonomous inertial solar tracking method for a deep space satellite according to any one of claims 1 to 11.

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