Spacecraft autonomous attitude determination method based on natural celestial body radio radiation

Abstract

The invention discloses a spacecraft autonomous attitude determination method based on natural celestial body radio radiation, and the method comprises the steps: firstly obtaining a noise signal of the radio radiation of a natural celestial body, and calculating the attitude of a spacecraft based on the noise signal, a pre-stored antenna pattern, and a solar ephemeris. The method can be used as a backup attitude determination scheme under extreme conditions of sensor faults and the like, does not depend on manually transmitted navigation signals, is not afraid of strong light interference, and is high in reliability and good in robustness.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and in particular to a method for autonomous attitude determination of spacecraft based on radio radiation from natural celestial bodies. Background Technology

[0002] Existing spacecraft are typically equipped with dedicated attitude sensors such as star sensors, sun sensors, magnetometers, and fiber optic gyroscopes, achieving high-precision attitude measurement through single or multiple sensor combinations. Among these, star sensors are the most accurate conventional attitude measurement components, achieving measurement precision on the order of arcseconds, and are widely used for high-precision attitude determination in spacecraft. However, they are expensive, have a large size and weight, and require complex star map database maintenance, making them unsuitable for operation in environments with strong light sources such as direct sunlight. Sun sensors measure the solar vector at the spacecraft's location; ideally, with multiple measurement surfaces working together, a full-sky field of view can be obtained in sunny areas. However, they require additional dedicated attitude determination hardware, which means independent cost, weight, power consumption burden, as well as additional interfaces and potential points of failure. Fiber optic gyroscopes directly measure the attitude angular velocity information in the spacecraft's inertial frame, offering advantages such as all-weather operation and no field-of-view constraints. However, their inherent error drift characteristics prevent them from operating independently for extended periods, requiring periodic calibration using absolute attitude measurement equipment such as star sensors, creating a cost-cumulative dependency. Magnetometers are effective in near-Earth orbit, but they rely on variable geomagnetic field models and are subject to specific working environments and conditions.

[0003] It can be seen that star sensors and sun sensors are the most important sensors in existing spacecraft. Once they fail, the spacecraft lacks a highly reliable backup attitude determination method. Summary of the Invention

[0004] To address some or all of the problems in the existing technology, the first aspect of this invention provides a spacecraft autonomous attitude determination method based on radio radiation from natural celestial bodies, as a backup attitude determination means in extreme situations such as sensor failure. The spacecraft autonomous attitude determination method includes: Acquiring noise signals from the radio emissions of natural celestial bodies; and The spacecraft's attitude is calculated based on the noise signal, the pre-stored radiation patterns of each antenna, and the solar ephemeris.

[0005] Furthermore, acquiring noise signals from the radio radiation of natural celestial bodies includes: Capturing broadband electromagnetic radiation signals emitted by natural celestial bodies using multiple directional antennas, wherein the multiple directional antennas are arranged non-coplanarly; and The broadband electromagnetic radiation signal is preprocessed and converted into a digital voltage signal.

[0006] Furthermore, 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.

[0007] Furthermore, calculating the spacecraft's attitude includes: Based on the noise signal, a first observation vector representing the direction of a natural celestial body is determined. Based on the solar ephemeris, determine the theoretical direction vector of the natural celestial body; and Based on the first observed 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.

[0008] Furthermore, calculating the spacecraft's attitude includes: Based on the noise signal, a first observation vector representing the direction of a natural celestial body is determined. The observation vector is combined with a second observation vector provided by the magnetometer, wherein the second observation vector is not collinear with the first observation vector; and Based on the combined observation vectors, the three-axis attitude matrix of the spacecraft is determined by a vector attitude determination algorithm.

[0009] Furthermore, determining the first observation vector representing the direction of a natural celestial body based on the noise signal includes: The vector that minimizes the overall difference between the expected power value and the actual measured value is taken as the first observation vector, wherein the expected power value is calculated based on the vector and the radiation pattern function of the directional antenna.

[0010] A second aspect of the present invention provides a spacecraft autonomous attitude determination system based on radio radiation from natural celestial bodies, comprising: A signal receiving module comprising at least two non-coplanar directional antennas arranged on a spacecraft platform, the signal receiving module being used to capture broadband electromagnetic radiation signals emitted by natural celestial bodies; The signal processing module includes a preprocessing unit, a power detection unit, and an analog-to-digital converter. The signal processing module is used to convert the broadband electromagnetic radiation signal into a digital voltage signal. An attitude calculation module, which is communicatively connected to the signal processing module, is used to calculate the spacecraft attitude based on the digital voltage signal.

[0011] Furthermore, the signal receiving module reuses the spacecraft's communication antenna.

[0012] Furthermore, the signal processing module reuses the signal processing units in the spacecraft's communication unit, box, and / or radar system.

[0013] Furthermore, the attitude calculation module reuses the spacecraft's onboard computer.

[0014] This invention provides a spacecraft autonomous attitude determination method and system based on radio radiation from natural celestial bodies. It uses radio signals from natural celestial bodies for spacecraft attitude determination, serving as a backup attitude determination solution in extreme situations such as sensor failure. The method does not rely on artificially transmitted navigation signals and is unaffected by strong light interference, exhibiting high reliability and enhancing the spacecraft's survivability. Furthermore, the method requires only a few antennas, and even if some antennas fail, as long as there are at least two remaining antennas that are not collinear, the system can still operate in degraded mode, demonstrating high robustness. Simultaneously, the system's hardware can reuse existing equipment on the spacecraft, such as existing telemetry, tracking, and data transmission antennas in idle states. No new hardware needs to be added specifically for navigation and attitude determination, resulting in a simple system structure, resource reuse, and reduced system weight, power consumption, and cost. The method and system can serve both emergency survival for near-Earth orbit spacecraft and autonomous attitude determination for deep space probes, demonstrating strong versatility. Testing shows that its attitude determination accuracy can reach the order of 2°, meeting the application requirements for emergency safety attitude determination and coarse attitude reference establishment for spacecraft. Attached Figure Description

[0015] To further illustrate the above and other advantages and features of the various embodiments of the present invention, a more specific description of the various embodiments of the present invention will be presented with reference to the accompanying drawings. It is to be understood that these drawings depict only typical embodiments of the invention and are therefore not intended to limit its scope. In the drawings, identical or corresponding parts will be indicated by identical or similar reference numerals for clarity.

[0016] Figure 1 This diagram illustrates a flowchart of a spacecraft autonomous attitude determination method based on radio radiation from natural celestial bodies, according to an embodiment of the present invention; and Figure 2 The 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

[0017] 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.

[0018] 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.

[0019] 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.

[0020] 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.

[0021] Building upon this foundation, the inventors conducted further research, deriving the possibility of capturing broadband electromagnetic radiation from natural celestial bodies and the resolution for attitude determination based on it, thus confirming the feasibility of the proposed scheme. Taking the sun as an example, the antenna received power is calculated based on flux density theory as follows: The sun's radiation in the microwave frequency band can be approximated as blackbody radiation, following the Rayleigh-Jeans law: , in, This refers to the spectral brightness, measured in W / m² / Hz / sr, where sr refers to steradian, a unit used to measure solid angles. Boltzmann's constant; The solar brightness temperature (T) is used in the microwave band. The solar radiation characteristics cannot be described using the photosphere temperature of the visible light band; instead, the concept of radio brightness temperature should be used. This brightness temperature T is an equivalent temperature, representing the temperature an ideal blackbody needs to reach to produce the observed radio radiation intensity. This refers to the typical brightness temperature of the sun at 2 GHz in the S-band. The frequencies corresponding to the microwave band are in Hz; and It is the speed of light.

[0022] For a directional antenna, the received noise power is: in: The coefficient corresponds to the received power loss of a single-polarization antenna; Let be the effective area of ​​the antenna, where Where is the wavelength corresponding to the center frequency of the antenna, and G is the antenna gain; The system receive bandwidth is expressed in Hz; and For the solid angle of the sun, where The radius of the sun, This represents the average distance between the Earth and the Sun.

[0023] Calculations were performed using typical S-band parameters of spacecraft, where the frequency... ,bandwidth and antenna gain The center frequency of the antenna corresponds to the wavelength. Effective antenna area Solar solid angle Spectral brightness and the received noise power .

[0024] The antenna received power is calculated based on antenna temperature theory as follows: When a spacecraft's receiving antenna observes a solar radiation source at a small solid angle, the equivalent antenna temperature contributed by the sun is calculated as follows: , in: The antenna temperature contributed to the sun, measured in Kelvin (K). To provide warmth for the sun's brightness; The solid angle of the sun, measured in steradian degrees; The solid angle of a directional antenna beam, expressed in steradian degrees, where G is the antenna gain; This refers to antenna efficiency.

[0025] The formula for calculating received power is: , in, This represents the receiver bandwidth, measured in Hz.

[0026] First, calculate the antenna beam solid angle: ; Then, calculate the antenna temperature contributed by the sun: ;as well as Finally, calculate the received power: .

[0027] As can be seen, the two independent calculation methods yielded similar results, verifying the reliability of the calculation.

[0028] Further analysis of the signal-to-noise ratio and system noise power was conducted. Assuming the system noise temperature is a typical value The system noise power is approximately The relative power change caused by the solar signal, i.e., the signal-to-noise ratio, is as follows: , Although seemingly low, this can be improved. Specifically, while solar radio radiation exhibits random noise in the time domain, its average power level generated by a particular antenna is a stable and predictable physical quantity over a short period when the relative geometry between the spacecraft and the sun remains unchanged. This power level can be accurately estimated through signal integration, i.e., time averaging, thus improving the accuracy of total power measurements. The improvement factor for integrating by 1 second is... Then the signal-to-noise ratio after integration is Based on this, the signal strength is perfectly sufficient for detection and extraction.

[0029] The measurement principle and mathematical model of the attitude determination method provided by this invention are as follows: Assuming the spacecraft is equipped with N non-coplanar antennas, and the solar direction is a unit vector s in the body coordinate system, then the solar radiation power data model obtained by the spacecraft analytically is: , in, The noise power measured for the i-th antenna is in W, and it is a direct measurement value. This is the system gain coefficient, measured in W / K, which is obtained through ground calibration or on-orbit calibration. Theoretically, The received power, ignoring noise, is when the sun's direction is exactly aligned with the antenna's maximum gain direction under normalized antenna gain. ; Let be the normalized gain of the i-th antenna in the solar direction s, and its value ranges from 0 to 1; The solar brightness temperature, in Kelvin (K), is a known astronomical parameter. The system noise temperature, in Kelvin (K), is a system design parameter; and To measure noise, it follows a mean of 0 and a variance of . The Gaussian distribution of white noise is known as white noise.

[0030] Based on the aforementioned model, the Cramer-Rao lower bound theory is derived, wherein the probability model includes measurement vectors. Conditional probability density: , in, The residual between the measured value and the ideal value; This indicates that the smaller the residual, the higher the probability density of the event, and vice versa; and This represents the joint probability of all N antennas, i.e., the measurement results of all antennas are considered simultaneously.

[0031] The log-likelihood function is calculated as follows: , The Fisher information matrix is ​​derived as follows: , Finally, through rigorous derivation, we can obtain: , The lower bound of the attitude estimation variance can then be calculated: .

[0032] Furthermore, accuracy analysis is performed based on actual antenna characteristics. A practical antenna pattern model is considered: , in, Beamwidth is analyzed in the context of a typical dual-antenna system, where antenna 1 has a beamwidth of [missing information]. Antenna 2 direction is ,in The included angle of the antenna. and beamwidth .

[0033] Based on this, in symmetrical positions The sensitivity is calculated as follows: Then in Place, Then the sum of squares of sensitivity is: .

[0034] Finally, the attitude resolution value is calculated. The signal-to-noise ratio parameter after integration is: , The theoretical accuracy is calculated as follows: Considering practical engineering factors, the antenna pattern calibration error is approximately 20%, the system gain instability is approximately 15%, the model simplification error is approximately 10%, and the total error amplification factor is approximately 1.6. Therefore, the actual accuracy is calculated as follows: , Through algorithm optimization and system design, it is possible to achieve... This angle meets the requirements for coarse attitude determination. Therefore, the noise power of radio radiation from natural celestial bodies can be effectively received and resolved by the spacecraft antenna, with an attitude determination accuracy on the order of 5°, which meets the requirements for coarse attitude determination.

[0035] The technical solution of the present invention will be further described below with reference to the accompanying drawings of the embodiments.

[0036] Figure 1 This diagram illustrates a flow chart of a spacecraft autonomous attitude determination method based on radio radiation from natural celestial bodies, according to an embodiment of the present invention. Figure 1 As shown, a spacecraft autonomous attitude determination method based on radio radiation from natural celestial bodies includes: First, in step 101, radio emission signals are acquired. Noise signals from the radio emission of natural celestial bodies are acquired. In one embodiment of the invention, broadband electromagnetic radiation signals emitted by natural celestial bodies are captured using multiple directional antennas. These broadband electromagnetic radiation signals are then preprocessed and converted into digital voltage signals, wherein the multiple directional antennas are arranged non-coplanarly. In one embodiment of the invention, the preprocessing mainly involves amplifying, filtering, and down-converting the broadband electromagnetic radiation signals, converting them into analog voltage signals, and finally sampling and quantizing the analog voltage signals to convert them into discrete voltage quantities. The voltage quantities from multiple channels are used for attitude calculation; and Finally, in step 102, the spacecraft attitude is calculated. Based on the noise signal, the pre-stored radiation patterns of each antenna, and the solar ephemeris, the spacecraft attitude is calculated. In one embodiment of the invention, firstly, a first observation vector representing the direction of a natural celestial body is determined based on the noise signal. In one embodiment of the invention, determining the first observation vector involves constructing an optimal estimation problem, i.e., finding a unit vector S such that when the unit vector S is substituted into the radiation pattern function of each antenna... At that time, the calculated expected power value Compared with actual measured value The overall difference is minimized, where the unit vector S represents the direction of the sun in the body coordinate system. In one embodiment of the invention, the least squares method is used to solve for the optimal solution. It should be understood that other optimal solution methods can also be used in other embodiments of the invention. In this case, a precise first observation vector representing the direction of the sun can be obtained. Furthermore, based on the aforementioned solar ephemeris, the theoretical direction vector of the celestial body can be determined. Therefore, the attitude determination problem of a spacecraft, that is, the rotation problem of the body coordinate system relative to the inertial coordinate system, is transformed into determining a rotation matrix. , making At this point, the spacecraft's attitude can be determined. However, when using a single natural celestial body vector, such as the solar vector, the rotation matrix... This is not unique, meaning the spacecraft can freely rotate around the solar vector, and the roll angle is uncertain. Therefore, only two-dimensional attitude information can be obtained, namely, the yaw and pitch angles of the spacecraft. Based on this, in one embodiment of the invention, 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 sensors such as a magnetometer, and based on the combined observation vector, a vector attitude determination algorithm is used to determine the three-axis attitude matrix of the spacecraft, wherein the second observation vector is not collinear with the first observation vector. In one embodiment of the invention, the vector attitude determination algorithm includes QUEST, which can uniquely calculate the complete three-axis attitude matrix of the satellite.

[0037] Based on the spacecraft autonomous attitude determination method described above Figure 2 This diagram illustrates a structural schematic of a spacecraft autonomous attitude determination system based on radio radiation from natural celestial bodies, according to an embodiment of the present invention. Figure 2 As shown, a spacecraft autonomous attitude determination system based on radio radiation from natural celestial bodies 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 broadband electromagnetic radiation signals emitted by natural celestial bodies; the signal processing module 202 is used to convert the broadband electromagnetic radiation signals into digital voltage signals; and the attitude calculation module 203 is used to calculate the spacecraft's attitude based on the digital voltage signals.

[0038] In one embodiment of the present invention, the signal receiving module 201 comprises at least two non-coplanar directional antennas on the spacecraft platform. The directional antennas can be other antennas already used on the spacecraft, such as S-band telemetry and control antennas or X-band communication antennas, which are multiplexed when their communication missions are idle to capture noise signals from radio radiation emitted by celestial bodies. During physical installation, the phase centers of the directional antennas are not at the same point, and there is a fixed, known angle between their maximum radiation directions, i.e., the direction of the main beam, for example, greater than 30 degrees, thus forming a fixed spatial baseline in the spacecraft's coordinate system.

[0039] In one embodiment of the present invention, the signal processing module 202 includes a radio frequency (RF) front-end, a power detection unit, and an analog-to-digital converter (ADC). The RF front-end is used for signal preprocessing and may include, for example, a low-noise amplifier, a filter, a down-converter, etc. The power detection unit is used to measure the total noise power received by each antenna, and the ADC converts the analog noise power into a digital quantity. In one embodiment of the present invention, the signal processing module reuses the signal processing units in the spacecraft's communication unit, box, and / or radar system.

[0040] In one embodiment of the present invention, the attitude calculation module 203 includes hardware and software components. The hardware component includes an embedded processor (such as a CPU or FPGA) and its associated memory, which can reuse the spacecraft's onboard computer. The software component includes radiation pattern data of each antenna pre-stored in memory or calculated online, recording the gain value of each antenna at different incident angles. And the solar ephemeris, and the attitude calculation algorithm as described above running in the processor memory.

[0041] In one embodiment of the present invention, the signal receiving module 201, i.e., multiple directional antennas, is connected to the radio frequency front-end of the signal processing module 202 via radio frequency cables. The digital output terminal of the signal processing module 202 is connected to the processor of the attitude calculation module 203 via a data bus, such as UART, I2C, CAN, etc. The signal receiving module 201 is responsible for sensing electromagnetic signals radiated by natural celestial bodies; the signal processing module 202 is responsible for converting physical signals into processable digital quantities; and the attitude calculation module 203 calculates the spacecraft attitude based on the digital quantities, combined with known antenna pattern quantities and attitude calculation algorithms.

[0042] The described spacecraft autonomous attitude determination method and system can be applied to emergency Earth orientation of geostationary orbit (GEO) communication satellites, autonomous navigation of deep space probes such as Mars orbiters, and attitude determination of low Earth orbit (LEO) CubeSats. Specifically, when a GEO communication satellite encounters a high-energy particle event in space, causing its star sensor to fail to reset, the spacecraft's autonomous attitude determination system can be activated. Using a telemetry and control antenna pointing towards Earth and a data transmission antenna pointing towards space, the system receives radiation from Earth (a strong noise source) and the cosmic background. By comparing the power differences received by each antenna, the system can calculate the satellite's orientation relative to Earth and control the satellite to maintain Earth orientation, ensuring uninterrupted communication until ground personnel troubleshoot the problem. During its journey to Mars, the deep space probe is outside the Earth-based deep space network's tracking and control arc. At this time, the spacecraft's autonomous attitude determination system continues to operate, using both the Sun (the strongest source) and Earth (a gradually weakening source) as dual beacons. By precisely measuring the intensity of these two sources on different antennas of the probe and combining this with an orbital dynamics model, the system can not only calculate the probe's three-axis attitude but also assist in orbit determination by observing continuous changes in Earth's orientation. The low-cost CubeSat in Low Earth Orbit (LEO) cannot carry star sensors due to budget and space constraints; therefore, it employs the aforementioned autonomous attitude determination method and system as its primary attitude determination means. It utilizes its existing UHF band tracking and control antenna and S-band data transmission antenna to continuously receive noise from the Sun. By calculating coarse attitude information, it can provide the CubeSat with two-axis attitude data, supporting its basic tasks such as solar orientation and communication.

[0043] Although various embodiments of the invention have been described above, it should be understood that they are presented by way of example only and not as limitations. It will be apparent to those skilled in the art that various combinations, modifications, and alterations can be made without departing from the spirit and scope of the invention. Therefore, the breadth and scope of the invention disclosed herein should not be limited by the exemplary embodiments disclosed above, but should be defined solely by the appended claims and their equivalents.

Claims

1. A method for autonomous attitude determination of a spacecraft based on natural celestial radio emission, characterized in that, The method comprises: acquiring a noise signal of radio radiation of a natural celestial body; and calculating an attitude of a spacecraft based on the noise signal, a pre-stored directional diagram of an antenna, and a sun ephemeris table.

2. The spacecraft autonomous attitude determination method of claim 1, wherein, The acquiring of the noise signal of radio radiation of a natural celestial body comprises: capturing a wideband electromagnetic radiation signal emitted by a natural celestial body through multiple directional antennas, wherein the multiple directional antennas are arranged in a non-coplanar manner; and pre-processing the wideband electromagnetic radiation signal and converting it into a digital voltage signal.

3. The spacecraft autonomous attitude positioning method of claim 2, wherein, The multiple directional antennas are multiplexed with a communication antenna of the spacecraft.

4. The spacecraft autonomous attitude positioning method of claim 2, wherein, The pre-processing comprises: amplifying, filtering, and down-converting the wideband electromagnetic radiation signal; converting the down-converted signal into an analog voltage signal; and sampling, quantizing, and converting the analog voltage signal into a discrete voltage quantity.

5. The spacecraft autonomous attitude positioning method of claim 4, wherein, The pre-processing of the wideband electromagnetic radiation signal is performed by a signal processing module, wherein the signal processing module is multiplexed with a signal processing unit in a communication machine and / or a radar system of the spacecraft.

6. The spacecraft autonomous attitude positioning method of claim 1, wherein, The calculating of the attitude of the spacecraft comprises: determining a first observation vector representing a direction of the natural celestial body based on the noise signal; determining a theoretical direction vector of the natural celestial body based on the sun ephemeris table; and 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.

7. The spacecraft autonomous attitude positioning method of claim 1, wherein, The calculating of the attitude of the spacecraft comprises: determining a first observation vector representing a direction of the natural celestial body based on the noise signal; 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; and determining a three-axis attitude matrix of the spacecraft based on the combined observation vector through a vector attitude determination algorithm.

8. The spacecraft autonomous attitude determination method of claim 6 or 7, wherein, The determining of the first observation vector representing the direction of the natural celestial body based on the noise signal comprises: determining a vector that minimizes the overall difference between an expected power value and an actually measured value as the first observation vector, wherein the expected power value is calculated based on the vector and a directional diagram function of the directional antenna.

9. The spacecraft autonomous attitude positioning method according to claim 6 or 7, characterized in that, The calculating of the attitude of the spacecraft is performed by an on-board computer of the spacecraft.

Images