Satellite attitude determination method, system and computer readable medium based on star sensor

Abstract

This invention relates to a satellite attitude determination method, system, and computer-readable medium based on a star sensor. The satellite attitude determination method includes: establishing a solar orientation coordinate system OX based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 Where O represents the satellite's center of mass, X i2 The axis is obtained by the cross product of the solar vector and the geocentric vector, Z. i2 The axis points in the negative direction of the solar vector, Y i2 The axis is X i2 Axis and Z i2 The axis is obtained according to the right-hand rule; the installation matrix of the star sensor is constructed according to the sun-oriented coordinate system; based on the installation matrix, the first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector are calculated in the sun-oriented coordinate system; the satellite's attitude is controlled to keep the first angle within a first preset range and the second angle within a second preset range, thereby making the star sensor effective throughout the entire orbital period. This application can prevent the star sensor's field of view from being affected by light.

Description

Technical Field

[0001] This invention relates primarily to the field of aerospace technology, and more specifically to a satellite attitude determination method, system, and computer-readable medium based on a star sensor. Background Technology

[0002] With the development of the space industry, the mission modes of existing satellites are becoming increasingly diverse, generally categorized into Earth-oriented satellites, inertial-oriented satellites, and solar-oriented satellites. Satellites performing different missions employ different coordinate systems. For example, Earth-oriented satellites use an orbital coordinate system, inertial-oriented satellites use an inertial coordinate system, while solar-oriented satellites may use the J2000 coordinate system or a custom coordinate system depending on the specific mission requirements. Satellites often carry multiple mission payloads. For instance, Earth-oriented payloads include cameras for remote sensing imaging of the Earth, sky-survey payloads include ultra-large telescopes for long-exposure observations of the sky in inertial space, and solar-oriented payloads include imagers for long-term observation of solar activity. For solar-oriented payloads that primarily observe the two-dimensional solar plane, they are often insensitive to rotation of the payload's line of sight. During solar observation, it is necessary to ensure the relative stability of the other two coordinate axes perpendicular to the payload's line of sight; therefore, the satellite typically needs to maintain a fixed attitude mode for extended periods.

[0003] For conventional satellites, star sensors are typically installed. These high-precision space attitude measurement devices provide accurate spatial orientation and reference for satellites and other aerospace vehicles. Measurement errors between different star sensors can be reduced through on-orbit calibration, precision temperature control, and information fusion from multiple star sensors. Star sensors are also optical sensors. During long-term observations of space targets, the field of view of star sensors is easily affected by sunlight, atmospheric radiation, and stray light. This can lead to star sensor failure, impacting their attitude determination accuracy and effectiveness. To improve the reliability of satellite attitude determination, backup attitude determination schemes are usually implemented. For example, when all star sensors fail, the satellite switches to other attitude sensors for attitude determination, such as dual-vector attitude determination using a magnetometer and a solar sensor, or gyro-integrated attitude determination. However, these non-star sensor-based attitude determination methods may reduce the satellite's attitude determination accuracy, thus affecting the satellite's on-orbit mission.

[0004] To ensure satellite attitude accuracy, two or more star sensors can be configured on the satellite. If one star sensor's field of view is affected by light and fails, the satellite can switch to another star sensor. However, switching between different star sensors may cause jitter in the satellite's attitude angle and angular velocity. Although this jitter can be mitigated through on-orbit calibration, it cannot be completely eliminated. Therefore, during satellite operation, it is desirable to keep the star sensors constantly available and avoid switching them as much as possible. Currently, the coordinate system used by sun-oriented satellites has a problem where the star sensor's field of view may be affected by light, leading to star sensor failure. Summary of the Invention

[0005] The technical problem to be solved by this application is to provide a satellite attitude determination method, system and computer-readable medium based on a star sensor, which can avoid the star sensor's field of view being affected by sunlight, ground light and stray light, so that the star sensor on the satellite is effective throughout the entire orbital period.

[0006] The technical solution adopted in this application to solve the above-mentioned technical problems is a satellite attitude determination method based on a star sensor, including: establishing a solar orientation coordinate system OX based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 Where O represents the satellite's center of mass, X i2 The axis is obtained by the cross product of the solar vector and the geocentric vector, Z. i2 The axis points in the negative direction of the solar vector, Y i2 The axis is X i2 Axis and Z i2 The axes are obtained according to the right-hand rule; based on the sun-oriented coordinate system OX. i2 Y i2 Z i2 Construct the installation matrix of the star sensor; based on the installation matrix, in the sun-oriented coordinate system OX... i2 Y i2 Z i2 The system calculates the first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector; it controls the satellite's attitude to keep the first angle within a first preset range and the second angle within a second preset range, thereby making the star sensor effective throughout the entire orbital period.

[0007] In one embodiment of this application, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 The steps also include: normalizing the solar vector S in the inertial coordinate system using the following formula. iObtain the normalized solar vector S z :

[0008]

[0009] According to the formula: Z i2 =-S z The normalized solar vector S z The negative direction is used as the sun-oriented coordinate system OX. i2 Y i2 Z i2 Z i2 axis.

[0010] In one embodiment of this application, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 The steps also include: based on the position vector R in the inertial coordinate system i Use the formula: Establish a unit vector R pointing towards the Earth's center. z .

[0011] In one embodiment of this application, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 The steps also include: establishing the sun-oriented coordinate system OX using the following formula. i2 Y i2 Z i2 X i2 axis:

[0012]

[0013] X i =S z ×R z

[0014] Where S z R represents the normalized solar vector. z represents a unit vector pointing towards the Earth's center, and × represents the vector cross product operation.

[0015] In one embodiment of this application, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 The steps also include: establishing the sun-oriented coordinate system OX using the following formula. i2 Y i2 Z i2 Yi2 axis:

[0016]

[0017] Y i =Z i2 ×X i2

[0018] Z i2 Represents the sun-oriented coordinate system OX i2 Y i2 Z i2 Z i2 Axis, X i2 Represents the sun-oriented coordinate system OX i2 Y i2 Z i2 X i2 axis.

[0019] In one embodiment of this application, the satellite attitude determination method further includes: based on the sun-oriented coordinate system OX i2 Y i2 Z i2 Constructing the satellite's attitude matrix The steps for calculating the satellite's attitude angles based on the attitude matrix and controlling the satellite's attitude include: controlling the satellite's attitude angles.

[0020] In one embodiment of this application, the inertial coordinate system includes the J2000 flat equatorial geocentric coordinate system.

[0021] In one embodiment of this application, the star sensor includes a first star sensor and a second star sensor, and the mounting matrix A of the first star sensor is... bmASTR include:

[0022]

[0023] And the installation matrix A of the second star sensor bmBSTR include:

[0024]

[0025] Where cosd(·) represents the cosine of the angle, and the numbers in cosd(·) represent the angle values.

[0026] To address the aforementioned technical problems, this application also proposes a satellite attitude determination system based on a star sensor, which is installed within the satellite's onboard computer. The satellite attitude determination system includes: a memory for storing instructions executable by a processor; and a processor for executing the instructions to implement the satellite attitude determination method described above.

[0027] To address the aforementioned technical problems, this application also proposes a computer-readable medium storing computer program code, which, when executed by a processor, implements the satellite attitude determination method described above.

[0028] The technical solution of this application is based on the satellite's mission requirements and establishes a solar orientation coordinate system OX according to the solar vector and position vector in the inertial coordinate system. i2 Y i2 Z i2 , where X i2 Axis, Y i2 Axis and Z i2 The axes form an orthogonal coordinate system, which is used as the basic coordinate system for satellite attitude control; based on the OX coordinate system... i2 Y i2 Z i2 The system constructs the installation matrix of the star sensor and calculates the first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector. Throughout the satellite's orbital period, the first and second angles are maintained within preset ranges by controlling the satellite's attitude. This ensures the star sensor is unaffected by sunlight, atmospheric radiation, and stray light from the sun and Earth, maintaining its continuous effectiveness without switching. This improves the reliability of the satellite's attitude determination algorithm and attitude measurement system, guarantees on-orbit attitude accuracy, and enhances operational safety. The satellite computer reduces the time spent determining the communication and arbitration status of the star sensor, simplifying the star sensor selection logic in the satellite computer code, improving code robustness, and increasing the speed at which the satellite computer processes satellite attitude data. Attached Figure Description

[0029] To make the above-mentioned objectives, features and advantages of this application more apparent and understandable, the specific embodiments of this application will be described in detail below with reference to the accompanying drawings, wherein:

[0030] Figure 1 This is an exemplary flowchart of a satellite attitude determination method based on a star sensor according to an embodiment of this application;

[0031] Figure 2 This application describes an embodiment of a sun-oriented coordinate system OX. i2 Y i2 Z i2 Exemplary schematic diagram;

[0032] Figure 3 This is a schematic diagram of the structure of a satellite equipped with a first star sensor and a second star sensor in one embodiment of this application;

[0033] Figure 4This is a schematic diagram of the simulation results of the angles between the optical axis of the first star sensor in one embodiment of this application and the geocentric vector and the solar vector, respectively.

[0034] Figure 5 This is a schematic diagram of the simulation results of the angles between the optical axis of the second star sensor and the geocentric vector and the solar vector in one embodiment of this application;

[0035] Figure 6 This is a system block diagram of a satellite attitude determination system based on a star sensor according to an embodiment of this application. Detailed Implementation

[0036] To make the above-mentioned objectives, features and advantages of this application more apparent and understandable, the specific embodiments of this application will be described in detail below with reference to the accompanying drawings.

[0037] Many specific details are set forth in the following description in order to provide a full understanding of this application. However, this application may also be implemented in other ways different from those described herein, and therefore this application is not limited to the specific embodiments disclosed below.

[0038] As indicated in this application and claims, unless the context clearly indicates otherwise, the words "a," "an," "an," and / or "the" are not specifically singular and may include plural forms. Generally speaking, the terms "comprising" and "including" only indicate the inclusion of explicitly identified steps and elements, which do not constitute an exclusive list, and the method or apparatus may also include other steps or elements.

[0039] Flowcharts are used in this application to illustrate the operations performed by the system according to embodiments of this application. It should be understood that the preceding or following operations are not necessarily performed in exact order. Instead, various steps can be processed in reverse order or simultaneously. Furthermore, other operations may be added to these processes, or one or more steps may be removed from these processes.

[0040] This application proposes a satellite attitude determination method based on a star sensor, which can be applied to satellites performing solar observation and solar exploration missions to ensure that the star sensor on the satellite is effective at all times throughout the entire orbital period.

[0041] Figure 1 This is an exemplary flowchart of a satellite attitude determination method based on a star sensor according to an embodiment of this application, with reference to... Figure 1 As shown, the satellite attitude determination method based on a star sensor in this embodiment includes the following steps:

[0042] Step S110: Establish the solar orientation coordinate system OX based on the solar vector and the position vector in the inertial coordinate system. i2 Y i2 Zi2 Where O represents the satellite's center of mass, X i2 The axis is obtained by the cross product of the solar vector and the geocentric vector, Z. i2 The axis points in the negative direction of the solar vector, Y i2 The axis is X i2 Axis and Z i2 The axis is obtained according to the right-hand rule.

[0043] Step S120: Based on the sun-oriented coordinate system OX i2 Y i2 Z i2 Construct the installation matrix for the star sensor.

[0044] Step S130: Based on the installation matrix in the sun-oriented coordinate system OX i2 Y i2 Z i2 The calculations include the first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector.

[0045] Step S140: Control the attitude of the satellite to keep the first included angle within a first preset range and the second included angle within a second preset range, so that the star sensor is effective throughout the entire orbital period.

[0046] First, let me briefly introduce the sun-oriented coordinate system OX in one embodiment of this application. i2 Y i2 Z i2 The meaning of each coordinate axis. Figure 2 This application describes an embodiment of a sun-oriented coordinate system OX. i2 Y i2 Z i2 An exemplary schematic diagram, with reference to Figure 2 As shown, O represents the satellite's center of mass, X i2 Axis, Y i2 Axis and Z i2 The axes form an orthogonal coordinate system, with the X0 axis pointing in the direction of satellite flight, the Y0 axis pointing in the direction of the negative normal to the orbital plane, and the S... i The axis represents the solar vector in the inertial coordinate system, S. i The axis points in the direction of the normal to the satellite's solar panels, R z The axis represents the unit vector pointing towards the Earth's center, R. z The axes point towards the Earth's center. The A-axis represents the optical axis design direction of the first star sensor, and the B-axis represents the optical axis design direction of the second star sensor. Figure 2 The coordinate axes shown will be explained later.

[0047] The following details steps S110 to S140 as described above:

[0048] refer to Figure 2 As shown, in step S110, based on the solar vector S in the inertial coordinate system... i Establish a sun-oriented coordinate system OX based on the position vector in the inertial coordinate system (not shown in the figure). i2 Y i2 Z i2 Where O represents the satellite's center of mass, X i2 The axis is determined by the solar vector S. i and the geocentric vector R z The cross product, Z i2 The axis points to the solar vector S i The negative direction of Y i2 The axis is X i2 Axis and Z i2 The axis is obtained according to the right-hand rule. For example, X... i2 Axis, Y i2 Axis and Z i2 The axes form an orthogonal coordinate system, and this application uses the sun-oriented coordinate system OX. i2 Y i2 Z i2 This serves as the basic coordinate system for satellite attitude control. In some embodiments, the inertial coordinate system includes the J2000 level equatorial geocentric coordinate system.

[0049] In some embodiments, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 The steps also include: normalizing the solar vector S in the inertial coordinate system using the following formula (1). i Obtain the normalized solar vector S z ; and according to the formula (2) below, the negative direction of the normalized solar vector Sz is taken as the solar orientation coordinate system OX. i2 Y i2 Z i2 Z i2 axis.

[0050]

[0051] Z i2 =-S z (2)

[0052] For example, the solar-facing side of a satellite is usually defined as the -Z plane. Then, the direction of the normal to the satellite's solar panels includes the direction from the satellite to the sun, and the vector of the normal to the satellite's solar panels in the satellite's body coordinate system is [00-1].

[0053] In some embodiments, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system.i2 Y i2 Z i2 The steps also include: based on the position vector R in the inertial coordinate system i Use the following formula (3) to establish the unit vector R pointing towards the Earth's center. z .

[0054]

[0055] In some embodiments, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 The steps also include: establishing the sun-oriented coordinate system OX using the following formulas (4)-(5) i2 Y i2 Z i2 X i2 axis.

[0056]

[0057] X i =S z ×R z (5)

[0058] Where S z R represents the normalized solar vector. z represents a unit vector pointing towards the Earth's center, and × represents the vector cross product operation.

[0059] In some embodiments, a solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system. i2 Y i2 Z i2 The steps also include: establishing the sun-oriented coordinate system OX using the following formulas (6)-(7) according to the right-hand rule. i2 Y i2 Z i2 Y i2 axis:

[0060]

[0061] Y i =Z i2 ×X i2 (7)

[0062] Z i2 Represents the sun-oriented coordinate system OX i2 Y i2 Z i2 Z i2 Axis, X i2Represents the sun-oriented coordinate system OX i2 Y i2 Z i2 X i2 axis.

[0063] In some embodiments, the satellite attitude determination method of this application further includes: based on the sun-oriented coordinate system OX i2 Y i2 Z i2 Constructing the satellite's attitude matrix The satellite's attitude angles are calculated based on the attitude matrix. For example, the sun-oriented coordinate system OX described above in this application... i2 Y i2 Z i2 This can be performed by the satellite's onboard computer, which uses the solar orientation coordinate system OX specified in this application. i2 Y i2 Z i2 An attitude matrix is ​​constructed, and then the satellite's attitude angles and angular velocities are calculated. Furthermore, the satellite's attitude can be controlled based on these attitude angles. The attitude matrix can be transformed into the desired attitude angles or the desired attitude quaternions for guidance law calculation.

[0064] In step S120, based on the sun-oriented coordinate system OX i2 Y i2 Z i2 Construct the mounting matrix of the star sensor. For example, the mounting matrix of the star sensor is used to represent the vector relationship between the star sensor coordinate system and the satellite body coordinate system.

[0065] Figure 3 This is a schematic diagram of a satellite structure equipped with a first star sensor and a second star sensor according to one embodiment of this application. (Reference) Figure 3 As shown, in some embodiments, the star sensor includes a first star sensor 320 and a second star sensor 330, and the mounting matrix A of the first star sensor 320 is shown. bmASTR Mounting matrix A of the second-star sensor 330 bmBSTR Represented by the following equations:

[0066]

[0067]

[0068] Where cosd(·) represents the cosine of the angle, and the numbers in cosd(·) represent the angle values. For example, the angle values ​​in the star sensor mounting matrix can be set to different values ​​so that different star sensors can cover different areas of the sky. Subsequently, the angle of the satellite can be calculated by combining the parameters of the star sensor mounting matrix. Figure 3The embodiment shown has two star sensors. Generally, the attitude determination accuracy of a dual-star sensor is better than that of a single-star sensor. This application does not limit the number of star sensors.

[0069] In step S130, the installation matrix is ​​used in the sun-oriented coordinate system OX. i2 Y i2 Z i2 The calculations include the first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector.

[0070] In step S140, the satellite's attitude is controlled so that the first included angle is maintained within a first preset range, and the second included angle is maintained within a second preset range, thereby making the star sensor effective throughout the entire orbital period. Exemplarily, the satellite's attitude can be controlled by adjusting its attitude angles. The first preset range includes greater than 35°, and the second preset range includes greater than 85°, by aligning the star sensor with the solar vector S. i The first angle between the star sensor and the geocentric vector is always greater than the star sensor's solar suppression angle of 35°, and the second angle between the star sensor and the geocentric vector is always greater than 85°. As a result, sunlight and atmospheric light will not enter the star sensor's shield and will not affect the star sensor's field of view. The star sensor remains effective throughout the entire orbital cycle and will not switch.

[0071] This application is based on the OX coordinate system for directional orientation. i2 Y i2 Z i2 A simulation experiment was conducted, for example, with reference to Figure 3 As shown, the initial structural coordinate system of satellite 310 includes the rectangular coordinate system OX. s Y s Z s Where the origin O is located at the center of the star-rocket separation surface, X s The axis is perpendicular to the Y-axis s Axis and Z s The plane enclosed by the axes, X s The direction of the axis is perpendicular to the origin and upwards, Z. s The direction of the axis is perpendicular to the paper and outwards from the origin. s The axis is located within the star-rocket separation plane and is parallel to X. s Axis and Z s The axes form a right-handed coordinate system. In the simulation experiment, the initial structural coordinate system OX of the satellite is used... s Y s Z s The sun-oriented coordinate system OX described above i2 Y i2 Z i2 By aligning the two satellites, their attitude can be controlled.

[0072] Figure 4 This is a schematic diagram showing the simulation results of the angles between the optical axis of the first star sensor and the geocentric vector and the solar vector in one embodiment of this application. Figure 5 This is a schematic diagram illustrating the simulation results of the angles between the optical axis of the second star sensor and the geocentric vector and the solar vector, respectively, in one embodiment of this application. In the simulation experiment, this application recorded the results of the simulation of the first star sensor (e.g.,...). Figure 4 The star sensor A) and the second star sensor (such as Figure 5 The long-term angle between the star-sensor B and the solar vector and the geocentric vector. Figure 4 and Figure 5 In the simulation results diagram, the horizontal axis represents time, in seconds (s), and the vertical axis represents angles, in degrees. (Reference) Figure 4 As shown, the angle between the first star sensor and the geocentric vector is consistently greater than 103°, and the angle with the solar vector is equal to 90°. (Reference) Figure 5 As shown, the angle between the second star sensor and the geocentric vector is consistently greater than 92°, and the angle with the solar vector is equal to 90°. Therefore, the solar orientation coordinate system OX based on this application... i2 Y i2 Z i2 Installing star sensors and controlling the satellite's attitude ensures that the dual-star sensors are effective throughout the entire orbital period. The dual-star sensors are not affected by Earth's atmosphere, sunlight, or other factors, and their attitude output remains normal.

[0073] For example, for solar imagers or survey telescopes that are not sensitive to the payload's line of sight, the aforementioned solar orientation coordinate system OX can be established. i2 Y i2 Z i2 This ensures that the star sensors on the satellite are always effective, preventing them from becoming ineffective due to the influence of the Earth or the Sun.

[0074] The technical solution of this application is based on the satellite's mission requirements and the installation relationship of the satellite's solar panels. A solar orientation coordinate system OX is established based on the solar vector and position vector in the inertial coordinate system. i2 Y i2 Z i2 , where X i2 Axis, Y i2 Axis and Z i2 The axes form an orthogonal coordinate system, which is used as the basic coordinate system for satellite attitude control; based on the OX coordinate system... i2 Y i2 Z i2The system constructs the installation matrix of the star sensor and calculates the first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector. Throughout the satellite's orbital period, the first and second angles are maintained within preset ranges by controlling the satellite's attitude. This ensures the star sensor is unaffected by sunlight, atmospheric radiation, and stray light from the sun and Earth, maintaining its continuous effectiveness without switching. This improves the reliability of the satellite's attitude determination algorithm and attitude measurement system, guarantees on-orbit attitude accuracy, and enhances operational safety. The satellite computer reduces the time spent determining the communication and arbitration status of the star sensor, simplifying the star sensor selection logic in the satellite computer code, improving code robustness, and increasing the speed at which the satellite computer processes satellite attitude data.

[0075] This application also includes a star sensor-based satellite attitude determination system, installed within the satellite's onboard computer. The system includes a memory and a processor. The memory stores instructions executable by the processor, which then executes these instructions to implement the star sensor-based satellite attitude determination method described above.

[0076] Figure 6 This is a system block diagram of a satellite attitude determination system based on a star sensor according to an embodiment of this application. (Reference) Figure 6 As shown, the star-sensor-based satellite attitude determination system 600 may include an internal communication bus 601, a processor 602, a read-only memory (ROM) 603, a random access memory (RAM) 604, and a communication port 605. When applied to a satellite service computer, the star-sensor-based satellite attitude determination system 600 may also include a hard disk 606. The internal communication bus 601 enables data communication between the components of the star-sensor-based satellite attitude determination system 600. The processor 602 can perform judgments and issue prompts. In some embodiments, the processor 602 may consist of one or more processors. The communication port 605 enables data communication between the star-sensor-based satellite attitude determination system 600 and external devices. In some embodiments, the star-sensor-based satellite attitude determination system 600 can send and receive information and data from a network through the communication port 605. The star sensor-based satellite attitude determination system 600 may also include different types of program storage units and data storage units, such as a hard disk 606, a read-only memory (ROM) 603, and a random access memory (RAM) 604, capable of storing various data files used for computer processing and / or communication, as well as possible program instructions executed by the processor 602. The processor executes these instructions to implement the main part of the method.

[0077] The above-described satellite attitude determination method based on star sensors can be implemented as a computer program, stored in hard disk 606, and loaded into processor 602 for execution to implement the satellite attitude determination method based on star sensors of this application.

[0078] This application also includes a computer-readable medium storing computer program code that, when executed by a processor, implements the star sensor-based satellite attitude determination method described above.

[0079] When a satellite attitude determination method based on a star sensor is implemented as a computer program, it can also be stored as an article of manufacture in a computer-readable storage medium. For example, computer-readable storage media can include, but are not limited to, magnetic storage devices (e.g., hard disks, floppy disks, magnetic stripes), optical discs (e.g., compact discs (CDs), digital multifunction discs (DVDs)), smart cards, and flash memory devices (e.g., electrically erasable programmable read-only memory (EPROM), cards, sticks, key drives). Furthermore, the various storage media described herein can represent one or more devices and / or other machine-readable media used for storing information. The term "machine-readable medium" can include, but is not limited to, wireless channels and various other media (and / or storage media) capable of storing, containing, and / or carrying code and / or instructions and / or data.

[0080] It should be understood that the embodiments described above are merely illustrative. The embodiments described herein may be implemented in hardware, software, firmware, middleware, microcode, or any combination thereof. For hardware implementation, the processor may 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, microcontrollers, microprocessors, and / or other electronic units designed to perform the functions described herein, or combinations thereof.

[0081] Some aspects of this application can be executed entirely by hardware, entirely by software (including firmware, resident software, microcode, etc.), or by a combination of hardware and software. The aforementioned hardware or software may be referred to as a "data block," "module," "engine," "unit," "component," or "system." The processor may 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 this application may manifest as computer products residing in one or more computer-readable media, including computer-readable program code. For example, computer-readable media may include, but are not limited to, magnetic storage devices (e.g., hard disks, floppy disks, magnetic tapes, etc.), optical discs (e.g., compressed CDs, digital multifunction DVDs, etc.), smart cards, and flash memory devices (e.g., cards, sticks, key drives, etc.).

[0082] A computer-readable medium may contain a propagated data signal containing computer program code, for example, on baseband or as part of a carrier wave. This propagated signal may take various forms, including electromagnetic, optical, and so on, or suitable combinations thereof. A computer-readable medium can be any computer-readable medium other than a computer-readable storage medium, which can be connected to an instruction execution system, apparatus, or device to enable communication, propagation, or transmission of a program for use. The program code located on the computer-readable medium can be propagated through any suitable medium, including radio, cable, fiber optic cable, radio frequency signals, or similar media, or any combination of the above media.

[0083] The basic concepts have been described above. Obviously, for those skilled in the art, the above disclosure is merely illustrative and does not constitute a limitation of this application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and corrections to this application. Such modifications, improvements, and corrections are suggested in this application, and therefore remain within the spirit and scope of the exemplary embodiments of this application.

[0084] Furthermore, this application uses specific terms to describe embodiments of the application. For example, "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic related to at least one embodiment of the application. Therefore, it should be emphasized and noted that "an embodiment," "one embodiment," or "an alternative embodiment" mentioned twice or more in different locations in this specification do not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0085] In some embodiments, numbers describing the quantity of components and attributes are used. It should be understood that such numbers used in the description of embodiments are modified in some examples with the terms "approximately," "approximately," or "generally." Unless otherwise stated, "approximately," "approximately," or "generally" indicates that the numbers are allowed to vary by ±20%. Accordingly, in some embodiments, the numerical parameters used in the specification and claims are approximate values, which may be changed depending on the characteristics required by individual embodiments. In some embodiments, numerical parameters should take into account specified significant digits and employ a general method of digit reservation. Although the numerical ranges and parameters used to confirm their breadth of scope in some embodiments of this application are approximate values, in specific embodiments, such values ​​are set as precisely as feasible.

Claims

1. A satellite attitude determination method based on a star sensor, characterized in that, include: A solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the same inertial coordinate system. i2 Y i2 Z i2 Where O represents the satellite's center of mass, X i2 The axis is obtained by the cross product of the solar vector and the geocentric vector, Z. i2 The axis points in the negative direction of the solar vector, Y. i2 The axis is composed of the X i2 Axis and the Z i2 The axes are obtained according to the right-hand rule; the solar orientation coordinate system OX is... i2 Y i2 Z i2 This serves as the basic coordinate system for attitude control of the satellite. According to the aforementioned sun-oriented coordinate system OX i2 Y i2 Z i2 Construct the mounting matrix of the star sensor; According to the installation matrix in the solar orientation coordinate system OX i2 Y i2 Z i2 The first angle between the star sensor and the solar vector, and the second angle between the star sensor and the geocentric vector are calculated. The attitude of the satellite is controlled so that the first included angle is kept within a first preset range and the second included angle is kept within a second preset range, thereby making the star sensor effective throughout the entire orbital period.

2. The satellite attitude determination method as described in claim 1, characterized in that, A solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the same inertial coordinate system. i2 Y i2 Z i2 The steps also include: The solar vector S in the inertial coordinate system is normalized using the following formula. i Obtain the normalized solar vector S z : According to the formula: Z i2 =-S z The normalized solar vector S z The negative direction is taken as the solar orientation coordinate system OX. i2 Y i2 Z i2 Z i2 axis.

3. The satellite attitude determination method as described in claim 2, characterized in that, A solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the same inertial coordinate system. i2 Y i2 Z i2 The steps also include: based on the position vector R in the inertial coordinate system i Use the formula: Establish a unit vector R pointing towards the Earth's center. z .

4. The satellite attitude determination method as described in claim 3, characterized in that, A solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the same inertial coordinate system. i2 Y i2 Z i2 The steps also include: establishing the solar orientation coordinate system OX using the following formula. i2 Y i2 Z i2 X i2 axis: X i =S z ×R z Where S z R represents the normalized solar vector. z The vector represents the unit vector pointing towards the Earth's center, and × represents the vector cross product operation.

5. The satellite attitude determination method as described in claim 4, characterized in that, A solar orientation coordinate system OX is established based on the solar vector in the inertial coordinate system and the position vector in the same inertial coordinate system. i2 Y i2 Z i2 The steps also include: establishing the solar orientation coordinate system OX using the following formula. i2 Y i2 Z i2 Y i2 axis: Y i =Z i2 ×X i2 Z i2 The solar orientation coordinate system OX represents i2 Y i2 Z i2 Z i2 Axis, X i2 The solar orientation coordinate system OX represents i2 Y i2 Z i2 X i2 axis.

6. The satellite attitude determination method as described in claim 5, characterized in that, Also includes: According to the aforementioned sun-oriented coordinate system OX i2 Y i2 Z i2 Construct the attitude matrix of the satellite Calculate the attitude angles of the satellite based on the attitude matrix; The steps for controlling the attitude of the satellite include: controlling the attitude angle of the satellite.

7. The satellite attitude determination method as described in claim 1, characterized in that, The inertial coordinate system includes the J2000 flat equatorial geocentric coordinate system.

8. The satellite attitude determination method as described in claim 1, characterized in that, The star sensor includes a first star sensor and a second star sensor, and the mounting matrix A of the first star sensor is... bmASTR include: and the mounting matrix of the second star sensor AbmBSTR include: Where cosd(·) represents the cosine value of the angle, and the numbers in cosd(·) represent the angle values.

9. A satellite attitude determination system based on a star sensor, installed within the satellite's onboard computer, characterized in that, The satellite attitude determination system includes: Memory is used to store instructions that can be executed by the processor; A processor for executing the instructions to implement the satellite attitude determination method as described in any one of claims 1-8.

10. A computer-readable medium storing computer program code, characterized in that, The computer program code, when executed by a processor, implements the satellite attitude determination method as described in any one of claims 1-8.

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