A spacecraft autonomous attitude determination method based on natural celestial radio radiation

Abstract

The application discloses a spacecraft autonomous attitude determination method based on natural celestial radio radiation, which firstly acquires a noise signal of natural celestial radio radiation, and calculates the attitude of the spacecraft based on the noise signal, a pre-stored antenna directional diagram and a sun ephemeris table. The method can be used as a backup attitude determination scheme in extreme cases such as sensor failure, and does not depend on artificial navigation signals and is not afraid of strong light interference, and has high reliability and good robustness.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace technology, and in particular to a spacecraft autonomous attitude determination method based on natural celestial radio radiation. BACKGROUND

[0002] The existing spacecraft is usually equipped with star sensors, sun sensors, magnetometers, fiber-optic gyroscopes and other special attitude sensors, and high-precision attitude measurement is achieved through single or multi-sensor combination. Among them, the star sensor is the most accurate component in the conventional attitude measurement component, and the measurement accuracy can reach the order of angle seconds, and it is widely used in high-precision attitude determination of spacecraft. However, it has high cost, large volume and weight, and complex star map database maintenance requirements, and cannot work in strong light source scenes such as direct sunlight. The sun sensor is used to measure the sun vector at the location of the spacecraft. In the ideal case, through the cooperation of multiple measurement surfaces, a full-sky field of view can be obtained in the sunlit area. However, it is a special attitude determination hardware that needs to be installed additionally, which means an independent cost, weight, power consumption burden, and additional interface and failure point. The fiber-optic gyroscope is used to directly measure the attitude angular velocity information of the spacecraft in the inertial system, and has the advantages of all-weather and no field of view constraints. However, the inherent error drift characteristic makes it unable to work independently for a long time, and it must rely on absolute attitude measurement equipment such as star sensor for regular correction, forming a cost superimposed dependent relationship. The magnetometer is effective in low earth orbit, but it depends on the variable geomagnetic field model and is subject to specific working environment and conditions.

[0003] It can be seen that the star sensor and the sun sensor in the existing spacecraft are the most important sensors, and once a fault occurs, the spacecraft lacks a high-reliability backup attitude determination means. SUMMARY

[0004] In view of some or all of the problems in the prior art, the first aspect of the present application provides a spacecraft autonomous attitude determination method based on natural celestial radio radiation, which is a backup attitude determination means in extreme cases such as sensor failure, and the spacecraft autonomous attitude determination method comprises:

[0005] acquiring a noise signal of radio radiation of a natural celestial body; and

[0006] calculating the attitude of the spacecraft based on the noise signal, pre-stored directional patterns of each antenna, and a sun ephemeris table.

[0007] Further, acquiring the noise signal of the radio radiation of the natural celestial body comprises:

[0008] capturing a wideband electromagnetic radiation signal emitted by the natural celestial body through multiple directional antennas, wherein the multiple directional antennas are arranged non-coplanarly; and

[0009] The wideband electromagnetic radiation signal is preprocessed and converted into a digital voltage signal.

[0010] Further, the preprocessing includes:

[0011] The wideband electromagnetic radiation signal is amplified, filtered, and down-converted.

[0012] The down-converted signal is converted into an analog voltage signal.

[0013] The analog voltage signal is sampled, quantized, and converted into a discrete voltage quantity.

[0014] Further, calculating the attitude of the spacecraft includes:

[0015] Determining a first observation vector representing the direction of the natural celestial body based on the noise signal.

[0016] Determining a theoretical direction vector of the natural celestial body based on the ephemeris.

[0017] Determining a rotation matrix of the spacecraft based on the first observation vector and the theoretical direction vector, and further determining a yaw angle and a pitch angle of the spacecraft.

[0018] Further, calculating the attitude of the spacecraft includes:

[0019] Determining a first observation vector representing the direction of the natural celestial body based on the noise signal.

[0020] Combining the observation vector with a second observation vector provided by a magnetometer, wherein the second observation vector is not collinear with the first observation vector.

[0021] Determining a three-axis attitude matrix of the spacecraft based on the combined observation vector through a vector-based attitude determination algorithm.

[0022] Further, determining a first observation vector representing the direction of the natural celestial body based on the noise signal includes:

[0023] Determining the first observation vector as a vector that minimizes the overall difference between an expected power value and an actual measured value, wherein the expected power value is calculated based on the vector and a directional pattern function of the directional antenna.

[0024] The second aspect of the present application provides a spacecraft autonomous attitude determination system based on natural celestial body radio radiation, comprising:

[0025] A signal receiving module including at least two pairs of non-coplanar directional antennas arranged on a spacecraft platform, the signal receiving module being configured to capture a wideband electromagnetic radiation signal emitted by a natural celestial body.

[0026] a signal processing module, which comprises a pre-processing unit, a power detection unit and an analog-to-digital converter, and is configured to convert the wideband electromagnetic radiation signal into a digital voltage signal; and

[0027] a pose calculation module, which is communicatively connected to the signal processing module and configured to calculate a spacecraft pose based on the digital voltage signal.

[0028] Further, the signal receiving module multiplexes the communication antenna of the spacecraft.

[0029] Further, the signal processing module multiplexes the signal processing unit in the communication unit, box or radar system of the spacecraft.

[0030] Further, the pose calculation module multiplexes the on-board computer of the spacecraft.

[0031] The present application provides a spacecraft autonomous pose determination method and system based on natural celestial radio radiation, which determines the pose of a spacecraft based on the radio signal of a natural celestial body and can be used as a backup pose determination scheme in extreme situations such as sensor failure. The method does not rely on artificially emitted navigation signals and is not afraid of strong light interference, has high reliability and enhances the survivability of the spacecraft. In addition, the method only needs to use a few antennas, and even if some antennas fail, as long as the number of remaining antennas is not less than 2 and they are not collinear, the system can still work in degraded mode, and has high robustness. At the same time, the hardware of the system can be reused on the spacecraft, such as existing measurement and control, data transmission antennas in idle state, etc. New hardware devices need to be added for navigation and pose determination. The system structure is simple, resources are reused, and the system weight, power consumption and cost are reduced. The method and system can serve not only the emergency survival of near-earth orbit spacecraft, but also the autonomous pose determination of deep space probes, and have strong universality. The pose determination accuracy of the method and system can reach the order of 2°, which can meet the application requirements of spacecraft emergency safety pose determination and coarse attitude reference establishment. BRIEF DESCRIPTION OF DRAWINGS

[0032] To further clarify the above and other advantages and features of the embodiments of the present application, a more particular description of the embodiments of the present application will be rendered by reference to the appended drawings. It is appreciated that these drawings depict only typical embodiments of the application and are therefore not to be considered limiting of its scope. In the drawings, the same or corresponding elements are denoted by the same or similar reference signs.

[0033] Figure 1 a flowchart of a spacecraft autonomous pose determination method based on natural celestial radio radiation according to an embodiment of the present application; and

[0034] Figure 2The diagram shows a schematic representation of an autonomous attitude determination system for spacecraft based on radio radiation from natural celestial bodies, according to an embodiment of the present invention. Detailed Implementation

[0035] In the following description, the invention is described with reference to various embodiments. However, those skilled in the art will recognize that the embodiments may be practiced without one or more specific details or in conjunction with other alternatives and / or additional methods or components. In other instances, well-known structures or operations are not shown or described in detail so as not to obscure the inventive points of the invention. Similarly, for illustrative purposes, specific numbers and configurations are set forth to provide a comprehensive understanding of embodiments of the invention. However, the invention is not limited to these specific details. Furthermore, it should be understood that the embodiments shown in the drawings are illustrative representations and are not necessarily drawn to scale.

[0036] In this specification, references to "an embodiment" or "this embodiment" mean that a particular feature, structure, or characteristic described in connection with that embodiment is included in at least one embodiment of the invention. The phrase "in one embodiment" appearing throughout this specification does not necessarily refer to the same embodiment in all instances.

[0037] It should be noted that the embodiments of the present invention describe the method steps in a specific order; however, this is only for illustrating the specific embodiment and not for limiting the order of the steps. On the contrary, in different embodiments of the present invention, the order of the steps can be adjusted according to actual needs.

[0038] Currently, there is a lack of a highly reliable, fully autonomous backup attitude determination method that is independent of external artificial signals, unaffected by strong light, and capable of utilizing existing platform hardware resources under extreme conditions. To address this issue, the inventors discovered that celestial bodies, such as the Sun, Earth, and Moon, emit broadband electromagnetic radiation. This broadband electromagnetic radiation, acting as a plane wavefront, arrives at different antennas on the same spacecraft almost simultaneously in physical space. Due to the different pointing directions of each antenna in its physical structure, the angle between the broadband electromagnetic radiation emitted by celestial bodies such as the solar radiation wavefront and the main beam direction of each antenna varies. They are also different. According to antenna theory, an antenna acts as a spatial filter, and the power of its output signal... With the angle of incidence function This means that the antenna's radiation pattern is directly related. Based on this, the spatial orientation information of natural celestial bodies can be encoded into electrical signals of different intensities by each antenna through its inherent physical directivity, and the attitude of the spacecraft can be calculated based on the intensity of these electrical signals.

[0039] On this basis, the inventors further researched the possibility of capturing the broadband electromagnetic radiation of natural celestial bodies and the resolution of the attitude determination based thereon, and confirmed the feasibility of the scheme. Taking the sun as an example, the antenna receiving power based on the flux density theory is calculated as follows:

[0040] The radiation of the sun in the microwave band can be approximated as blackbody radiation, which follows the Rayleigh-Jeans law:

[0041] ,

[0042] wherein,

[0043] is the spectral brightness, with the unit of W / m² / Hz / sr, wherein sr refers to a unit for measuring a solid angle;

[0044] is the Boltzmann constant;

[0045] is the solar brightness temperature, in the microwave band, the radiation characteristics of the sun cannot be described by the photosphere temperature in the visible light band, and the concept of radio brightness temperature should be adopted. The brightness temperature T is an equivalent temperature, which indicates how high a temperature an ideal blackbody needs to reach to produce the observed radio radiation intensity. is the typical value of the brightness temperature of the sun in the S band, i.e., 2 GHz;

[0046] is the corresponding frequency in the microwave band, with the unit of Hz; and

[0047] is the speed of light.

[0048] For a directional antenna, the received noise power is:

[0049]

[0050] wherein,

[0051] the coefficient corresponds to the receiving power loss of a single-polarized antenna;

[0052] is the effective area of the antenna, wherein is the wavelength corresponding to the center frequency of the antenna, and G is the antenna gain;

[0053] is the system receiving bandwidth, with the unit of Hz; and

[0054] is the solid angle of the sun, wherein is the radius of the sun, The average distance between the Earth and the Sun.

[0055] Taking typical spacecraft S-band parameters for calculation, the frequency , bandwidth , and antenna gain , the antenna center frequency corresponds to the wavelength , the antenna effective area , the solar solid angle , the spectral brightness , and the received noise power .

[0056] The antenna received power based on the antenna temperature theory is calculated as follows:

[0057] When the spacecraft's receiving antenna observes a small solid angle of the sun's radiation source, the equivalent antenna temperature contributed by the sun is calculated as follows:

[0058] ,

[0059] Where:

[0060] T is the antenna temperature contributed by the sun, in K;

[0061] T is the solar brightness temperature;

[0062] Ω is the solid angle of the sun, in steradians;

[0063] Ω is the solid angle of the directional antenna beam, in steradians, where G is the antenna gain;

[0064] η is the antenna efficiency.

[0065] The received power calculation formula is:

[0066] ,

[0067] Where,

[0068] B is the receiver bandwidth, in Hz.

[0069] First, calculate the antenna beam solid angle:

[0070] ;

[0071] Then, calculate the antenna temperature contributed by the sun:

[0072] ; and

[0073] Finally, the received power is calculated:

[0074] .

[0075] It can be seen that the two independent calculation methods obtain similar results, verifying the reliability of the calculation.

[0076] Further analysis of the signal-to-noise ratio, the system noise power , assuming that the system noise temperature is a typical value , then the system noise power is about , then the relative power change caused by the solar signal, i.e. the signal-to-noise ratio, is as follows:

[0077] ,

[0078] Although it seems low, it can be improved. Specifically, the solar radio radiation appears as random noise in the time domain, but the average power level it produces in a specific antenna is a stable and predictable physical quantity in a short period of time when the relative geometry of the spacecraft and the sun does not change. By integrating the signal, i.e. the time-averaged method, the power level can be accurately estimated, improving the accuracy of the total power measurement. The improvement factor of integrating 1 second is , then the signal-to-noise ratio after integration is . Based on this, the signal strength is sufficient for detection and extraction.

[0079] The measurement principle and mathematical model of the attitude determination method provided by the application are as follows. Assuming that the spacecraft is equipped with N non-coplanar antennas, the direction of the sun in the body coordinate system is a unit vector s, then the solar radiation power data model obtained by the spacecraft is:

[0080] ,

[0081] wherein,

[0082] is the noise power measured by the i-th antenna, with a unit of W, which is a direct measurement value;

[0083] is the system gain coefficient, with a unit of W / K, which is obtained by ground calibration or on-orbit calibration. In theory, is the power received without considering noise when the sun direction is exactly aligned with the maximum gain direction of the antenna under the normalized antenna gain, i.e. ;

[0084] is the normalized gain of the i-th antenna in the direction of the sun s, which takes a value of 0 to 1;

[0085] Tsunis the sun brightness temperature in K, which is a known astronomical parameter;

[0086] Tsysis the system noise temperature in K, which is a system design parameter; and

[0087] is the measurement noise, which is subject to a Gaussian distribution with mean 0 and variance , i.e., it is white noise.

[0088] Cramer-Rao lower bound theory derivation based on the model, wherein the probability model includes a conditional probability density of a measurement vector :

[0089] ,

[0090] wherein,

[0091] is the residual error of the measurement value and the ideal value;

[0092] indicates that the smaller the residual error, the greater the probability density of the event, and vice versa, the greater the residual error, the smaller the probability density; and

[0093] indicates the joint probability of all N antennas, i.e., considering the measurement results of all antennas at the same time.

[0094] The log-likelihood function is calculated as follows:

[0095] ,

[0096] The Fisher information matrix is derived as follows:

[0097] ,

[0098] Finally, through rigorous derivation, it can be obtained that:

[0099] ,

[0100] Further, the attitude estimation variance lower bound can be calculated as:

[0101] .

[0102] Further, based on the actual antenna characteristics, the precision is analyzed. Considering an engineering practical antenna pattern model:

[0103] ,

[0104] wherein, beam width, for a typical dual-antenna system, wherein the direction of antenna 1 is , the antenna 2 direction is where is the antenna included angle, and the beam width .

[0105] Based on this, the sensitivity at the symmetric position is calculated as follows:

[0106]

[0107] At , , the sum of the sensitivities is:

[0108] .

[0109] Finally, the attitude resolution value is calculated. The integral signal-to-noise ratio parameter is:

[0110] ,

[0111] The theoretical accuracy is calculated as follows:

[0112]

[0113] Considering the actual engineering factors, the antenna pattern calibration error is about 20%, the system gain instability is about 15%, the model simplification error is about 10%, and the total error amplification factor is about 1.6, the actual accuracy is calculated as:

[0114] ,

[0115] And through algorithm optimization and system design, the angle can be realized, which meets the rough attitude determination requirement. Therefore, the noise power of natural celestial radio radiation can be effectively received and resolved by the spacecraft antenna, and the attitude determination accuracy can reach the order of 5°, meeting the rough attitude determination requirement.

[0116] The technical solutions of the present application will be further described in combination with the embodiment drawings.

[0117] Figure 1 A flowchart of a spacecraft autonomous attitude determination method based on natural celestial radio radiation according to an embodiment of the present application is shown. As Figure 1 shown, a spacecraft autonomous attitude determination method based on natural celestial radio radiation comprises:

[0118] First, in step 101, a radio radiation signal is acquired. A noise signal of radio radiation of a natural celestial body is acquired. In one embodiment of the present application, a wide-band electromagnetic radiation signal emitted by a natural celestial body is captured by a plurality of directional antennas, and then pre-processed and converted into a digital voltage signal, wherein the plurality of directional antennas are arranged in a non-coplanar manner. In one embodiment of the present application, the pre-processing mainly refers to amplifying, filtering, and down-converting the wide-band electromagnetic radiation signal into an analog voltage signal, and finally sampling and quantizing the analog voltage signal into discrete voltage values. Voltage values of multiple channels are used to solve the attitude of the spacecraft; and

[0119] Finally, in step 102, the attitude of the spacecraft is solved. Based on the noise signal, pre-stored directional diagram of each antenna, and the sun ephemeris, the attitude of the spacecraft is calculated. In one embodiment of the present application, first, a first observation vector representing the direction of the natural celestial body is determined based on the noise signal. In one embodiment of the present application, the first observation vector is determined by constructing an optimal estimation problem, i.e., finding a unit vector S, such that when the unit vector S is substituted into the directional diagram function of each antenna the overall difference between the calculated expected power value and the actual measured value is the smallest, wherein the unit vector S represents the direction of the sun in the body coordinate system. In one embodiment of the present application, the least square method is used to solve the optimal solution. It should be understood that in other embodiments of the present application, other optimal solution solving methods can also be used. At this time, an accurate first observation vector representing the direction of the sun is obtained. At the same time, based on the sun ephemeris, a theoretical direction vector of the natural celestial body can be determined, and then the attitude determination problem of the spacecraft, i.e., the rotation problem of the body coordinate system relative to the inertial coordinate system, is converted into determining a rotation matrix such that , and thus the attitude determination of the spacecraft can be completed. However, when a single natural celestial body vector, such as the sun vector, is used, the rotation matrix Not unique, i.e. the spacecraft can freely rotate around the sun vector, the roll angle is uncertain, so at this time only two-dimensional pose information can be obtained, i.e. the yaw angle and the pitch angle of the spacecraft are determined. Based on this, in an embodiment of the present application, further fusion with other sensors is performed to obtain a three-dimensional attitude matrix. Specifically, the first observation vector is combined with a second observation vector provided by a magnetometer and the like, and based on the combined observation vector, a three-axis attitude matrix of the spacecraft is determined through a vector pose determination algorithm, wherein the second observation vector is not collinear with the first observation vector. In an embodiment of the present application, the vector pose determination algorithm includes QUEST and the like, which can uniquely solve the complete satellite three-axis attitude matrix.

[0120] Based on the spacecraft autonomous pose determination method as described above, Figure 2 An embodiment of the present application shows a structural schematic diagram of a spacecraft autonomous pose determination system based on natural celestial radio radiation, as shown in Figure 2 As shown, a spacecraft autonomous pose determination system based on natural celestial radio radiation includes a signal receiving module 201, a signal processing module 202, and an attitude calculation module 203. The signal receiving module 201 is used to capture the wideband electromagnetic radiation signal emitted by the natural celestial body, the signal processing module 202 is used to convert the wideband electromagnetic radiation signal into a digital voltage signal, and the attitude calculation module 203 is used to calculate the spacecraft attitude based on the digital voltage signal.

[0121] In an embodiment of the present application, the signal receiving module 201 is composed of at least two pairs of non-coplanar directional antennas arranged on the spacecraft platform. The directional antennas can use other existing antennas on the spacecraft, such as S-band TT&C antennas, X-band communication antennas, etc., which are multiplexed when their communication tasks are idle to capture the noise signal of the radio radiation of the natural celestial body. The phase centers of the directional antennas are not at the same point when physically installed, and there is a fixed and known angle between the directions of their maximum radiation, i.e. the main beams, for example, greater than 30 degrees, thereby forming a fixed spatial baseline in the spacecraft body coordinate system.

[0122] In an embodiment of the present application, the signal processing module 202 includes a radio frequency front end, a power detection unit, and an analog-to-digital converter. The radio frequency front end is used for signal preprocessing, which can include a low-noise amplifier, a filter, a frequency downconverter, etc., the power detection unit is used to measure the total noise power received by each antenna, and the analog-to-digital converter converts the analog noise power into a digital quantity. In an embodiment of the present application, the signal processing module multiplexes the signal processing unit in the communication unit, box, or radar system of the spacecraft.

[0123] In an embodiment of the present application, the attitude calculation module 203 comprises both hardware and software parts. The hardware part includes an embedded processor (such as a CPU or FPGA) and its associated memory, which can be reused from the on-board computer of the spacecraft. The software part includes the directional pattern data of each antenna, which records the gain value of each antenna at different incident angles, pre-stored in the memory or calculated online and stored in the memory and the sun ephemeris, and the attitude calculation algorithm as described above running in the processor memory.

[0124] In an embodiment of the present application, the signal receiving module 201, i.e. multiple directional antennas, is connected to the RF front end of the signal processing module 202 through RF cables, and the digital output of the signal processing module 202 is connected to the processor of the attitude calculation module 203 through a data bus such as UART, I2C, CAN, etc. The signal receiving module 201 is responsible for sensing the electromagnetic signals of natural celestial bodies; the signal processing module 202 is responsible for converting physical signals into digital quantities that can be processed; and the attitude calculation module 203 calculates the attitude of the spacecraft based on the digital quantities, in combination with the known quantities of the antenna directional patterns and the attitude calculation algorithm.

[0125] The spacecraft autonomous attitude determination method and system can be applied to emergency earth orientation of geosynchronous orbit (GEO) communication satellites, autonomous navigation of deep space probes such as Mars orbiters, and attitude determination of low earth orbit (LEO) cubesats, etc. Specifically, when a geosynchronous orbit (GEO) communication satellite encounters a high-energy particle event in space, causing the star sensor to reset failure, the spacecraft autonomous attitude determination system can be started to receive radiation from the Earth (a strong noise source) and the cosmic background using the earth-pointing TT&C antenna and the space-pointing TT&C antenna. By comparing the power differences received by each antenna, the system can calculate the pointing of the satellite relative to the Earth and control the satellite to maintain earth orientation, ensuring that the communication link is not interrupted until the ground personnel perform troubleshooting. Deep space probes are outside the earth deep space network TT&C arc segment during their flight to Mars. At this time, the spacecraft autonomous attitude determination system is continuously working, and the sun (the strongest source) and the Earth (a gradually weakening source) are used as dual beacons. By accurately measuring the intensity of these two sources on the probe's different antennas and combining with the orbit dynamics model, the system can not only calculate the three-axis attitude of the probe, but also assist in orbit determination by observing the continuous change of the Earth direction. The low-cost cubesat of the low earth orbit (LEO) cannot carry a star sensor due to budget and space limitations, so it uses the spacecraft autonomous attitude determination method and system as its main attitude determination means. It continuously receives noise from the sun using its existing UHF band TT&C antenna and S band TT&C antenna. By calculating the coarse attitude information, it can provide two-axis attitude data for the cubesat to support its completion of the sun orientation, communication and other basic tasks.

[0126] While the foregoing describes various embodiments of the application, such description should be considered as exemplary and not restrictive. Various combinations, modifications and alterations of the described embodiments are apparent to those skilled in the art from this disclosure, and are intended to be within the scope of the application. Accordingly, the breadth and scope of the application should not be limited by any of the above-described exemplary embodiments, but should be defined only in accordance with the following claims and their equivalents.

Claims

1. A spacecraft autonomous attitude determination method based on radio radiation from natural celestial bodies, characterized in that, include: The broadband electromagnetic radiation signals emitted by natural celestial bodies are captured by multiple directional antennas, and the broadband electromagnetic radiation signals are preprocessed and converted into digital voltage signals. Based on the digital voltage signal, a first observation vector representing the direction of a natural celestial body is determined. Based on the solar ephemeris, a theoretical direction vector of the natural celestial body is determined. Based on the first observation vector and the theoretical direction vector, the rotation matrix of the spacecraft is determined, and then the yaw angle and pitch angle of the spacecraft are determined. as well as The first observation vector is combined with the second observation vector provided by the magnetometer. Based on the combined observation vector, the three-axis attitude matrix of the spacecraft is determined by a vector attitude determination algorithm, wherein the second observation vector is not collinear with the first observation vector.

2. The spacecraft autonomous attitude determination method as described in claim 1, characterized in that, The multiple directional antennas are arranged in a non-coplanar manner.

3. The spacecraft autonomous attitude determination method as described in claim 1, characterized in that, The multiple directional antennas are reused with the spacecraft's communication antenna.

4. The spacecraft autonomous attitude determination method as described in claim 1, characterized in that, The preprocessing includes: The broadband electromagnetic radiation signal is amplified, filtered, and down-converted. Convert the down-converted signal into an analog voltage signal; and The analog voltage signal is sampled and quantized to convert it into discrete voltage values.

5. The spacecraft autonomous attitude determination method as described in claim 4, characterized in that, The broadband electromagnetic radiation signal is preprocessed by the signal processing module of the spacecraft's communication unit and / or radar system, wherein the signal processing module of the communication unit includes a radio frequency front-end, a power detection unit, and an analog-to-digital converter.

6. The spacecraft autonomous attitude determination method as described in claim 1, characterized in that, Determining the first observation vector representing the direction of a natural celestial body based on the digital voltage signal includes: By substituting unit vectors in different directions into the radiation pattern function of each antenna, the expected power value is calculated; and The unit vector that minimizes the overall difference between the expected power value and the actual measured value will be used as the first observation vector.

7. The spacecraft autonomous attitude determination method as described in claim 1, characterized in that, The spacecraft's attitude is calculated using its onboard computer.

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