Space craft, ground station, and antenna

Abstract

To provide a space craft capable of estimating an arrival direction of an electromagnetic wave with a desired spatial resolution by using a multilayer interference measurement method that is not limited by a distance between element antennas.SOLUTION: A space craft includes: an antenna part 20 including a first layer film 11 obtained by providing a plurality of slit-shaped openings 13 in a film through which an electromagnetic wave 16 is impermeable, and a second layer film 14 provided with a power reception element 15 that detects an electromagnetic wave that has transmitted through the plurality of openings and interfered; and a processing part 21 that estimates an arrival direction of the electromagnetic wave from a signal intensity of the electromagnetic wave that the power reception element has detected.SELECTED DRAWING: Figure 4

Description

[Technical Field]

[0001] The present invention relates to a spacecraft, a ground station, and an antenna for conducting electromagnetic wave observations. [Background technology]

[0002] Electromagnetic wave observation systems used for astronomical observation, remote sensing, etc. require highly accurate determination of the direction of arrival of electromagnetic waves from the electromagnetic wave source. The spatial resolution of the interferometer measurement method (called rotational interferometry) used in VLBI (Very Long Baseline Interferometry) observation systems, which is a type of radio interferometer, is a function of the distance D between the element antennas and the observation wavelength λ, and is expressed as λ / D (radian).

[0003] Furthermore, a method is being considered in which electromagnetic waves are observed by a single satellite (spacecraft) using the principle of rotational interferometry of a VLBI observation system.As an example of an element antenna deployed on a single spacecraft, Patent Document 1 discloses an antenna system for space deployment that is folded when transported into space and is deployed in space, and that includes multiple panels that are movable relative to each other. [Prior art documents] [Patent documents]

[0004] [Patent Document 1] Japanese Patent Application Laid-Open No. 2015-80213 Summary of the Invention [Problem to be solved by the invention]

[0005] VLBI measurements, developed for the purpose of determining the positions of celestial bodies, enable radio source detection with high spatial resolution by placing element antennas on the ground at distances of 1,000 km or more and using the Earth's rotation to overlap points where the interference between the signal strengths of antenna pairs is strong. In contrast, when using rotational interferometry to identify the direction of arrival of radio waves from a radio source on the ground using a spacecraft equipped with element antennas as in Patent Document 1, the distance D between the element antennas depends on the size of the spacecraft itself, and the spatial resolution is restricted by the distance D between the element antennas, resulting in a low degree of freedom in spatial resolution. Therefore, the inventors of the present application discovered a new multi-layer interferometry method that can estimate the direction of arrival of electromagnetic waves on a spacecraft alone without being restricted by the distance D between the element antennas, and considered installing it on a spacecraft. [Means for solving the problem]

[0006] A spacecraft according to one embodiment of the present invention comprises an antenna section including a first layer membrane having a plurality of slit-shaped openings formed in a membrane that is opaque to electromagnetic waves having a wavelength to be observed, and a second layer membrane having a receiving element that detects the electromagnetic waves that have passed through the plurality of openings and interfered, and a processing section that estimates the direction of arrival of the electromagnetic waves from the signal strength of the electromagnetic waves detected by the receiving element. [Effects of the Invention]

[0007] The present invention provides a spacecraft that can estimate the direction of arrival of electromagnetic waves with a desired spatial resolution based on the positional relationship between the slit-shaped opening and the receiving element. Other objects and novel features will become apparent from the description of this specification and the accompanying drawings. [Brief explanation of the drawings]

[0008] [Figure 1] FIG. 1 is a schematic configuration diagram of a spacecraft. [Figure 2A] FIG. 10 is a diagram for explaining a multilayer interferometry method. [Figure 2B] FIG. 10 is a diagram for explaining a multilayer interferometry method. [Figure 2C] FIG. 10 is a diagram for explaining a multilayer interferometry method. [Figure 3] FIG. 10 is a diagram for explaining a multilayer interferometry method. [Figure 4] FIG. 1 is a block diagram of an electromagnetic wave observation system. [Figure 5] 1 shows an example of the configuration of a first layer film and a second layer film that constitute an antenna part. [Figure 6] This is an example of a triangular opening. [Figure 7] 1 is an example of an equivalent circuit of a software defined radio. [Figure 8] FIG. 2 is a functional block diagram of an electromagnetic wave arrival direction estimation process performed by a processor. [Figure 9A] 10 is a configuration example of an antenna unit in which the power receiving element group includes five power receiving elements. [Figure 9B] FIG. 10 is a diagram showing the incidence angle dependency of the signal strength of each power receiving element. [Figure 10A] FIG. 2 is a functional block diagram of an electromagnetic wave arrival direction estimation process performed by a processor. [Figure 10B] 10 is a configuration example of an antenna unit provided with a reference receiving element. [Figure 10C] FIG. 10 is a diagram for explaining a method for estimating an incident angle. [Figure 11] FIG. 2 is a functional block diagram of an electromagnetic wave arrival direction estimation process performed by a processor. [Figure 12] FIG. 2 is a functional block diagram of an electromagnetic wave arrival direction estimation process performed by a processor. [Figure 13] FIG. 1 is a side view of a spacecraft. [Figure 14] A bird's-eye view of the spacecraft. [Figure 15] This is the flow for estimating the direction of a radio wave source using an electromagnetic wave observation system. [Figure 16] FIG. 16 is an explanatory diagram of the flow of FIG. [Figure 17] This is a flow chart for estimating the direction of a radio wave source by using both rotational interferometry and multilayer interferometry. [Figure 18] FIG. 10 is a diagram for explaining Modification 1. [Figure 19] FIG. 10 is a diagram for explaining Modification 2. [Figure 20]FIG. 10 is a diagram for explaining Modification 3. [Figure 21] 1 is a configuration example of an electromagnetic wave monitoring system. [Figure 22] FIG. 10 is a diagram for explaining Modification 4. DETAILED DESCRIPTION OF THE INVENTION

[0009] Figure 1 shows the schematic configuration of a spacecraft. Spacecraft 1 is a three-dimensional spacecraft with a virtual regular polyhedron shape formed with the approximate center of the spacecraft as the origin. A deployable beam (hereinafter referred to as beam) 3 is built into satellite casing 2, and as the beam 3 is deployed in space, a thin film 4 is extended. For example, multiple patch antennas 5 can be arranged on the thin film 4. By synchronizing the phase of the patch antennas 5, a highly directional transmitting and receiving antenna can be configured. Details of the structure of such a three-dimensional spacecraft are disclosed in Japanese Patent Application No. 2023-105999 by the same applicant.

[0010] The interferometric measurement of this embodiment takes advantage of the spacecraft 1's three-dimensional structure and its multilayered structure. For example, the first layer can be the thin film 4, and the second layer can be the surface of the satellite housing 2. This reduces the number of required films. Alternatively, in a spacecraft with multiple thin films 4, the first layer can be the outer thin film (the one farther from the satellite housing 2) and the second layer can be the inner thin film (the one closer to the satellite housing 2). The interferometric measurement of this embodiment generates diffracted waves from the incoming electromagnetic waves in the first layer, and measures the interference pattern between the diffracted waves using an antenna installed in the second layer. Because the interference pattern changes depending on the incident angle θ of the electromagnetic waves, the direction of arrival of the electromagnetic waves can be estimated based on the interference pattern. The interferometric measurement method of this embodiment is called multilayer interferometry.

[0011] (Principle of multilayer interferometry) The multilayer interferometry method will be described using Figures 2A to 2C. Figure 2A is a schematic diagram of the configuration of an antenna unit for performing interferometry using the multilayer interferometry method. The first layer 11 is a film that is opaque to electromagnetic waves at the observation wavelength and has multiple slit-shaped openings 13 formed therein. The portion of the film that is opaque to electromagnetic waves is sometimes called a shielding portion 12. The second layer 14 is provided with a power receiving element 15 that detects electromagnetic waves that have passed through the openings 13 in the first layer 11. Note that for simplicity of explanation, Figure 2A shows an example in which the first layer 11 and the second layer 14 are arranged parallel to each other, but this arrangement is not limiting. Furthermore, when the slit-shaped opening 13 is rectangular, it is preferable that the power receiving element 15 extend in the same direction as the opening 13. However, any shape is acceptable as long as the interference of the diffracted wave from the first layer 11 can be detected by the power receiving element 15 in the second layer 14.

[0012] FIG. 2B shows a schematic diagram of how electromagnetic waves transmitted through the first layer 11 are detected by the power receiving element 15. Because the source of the electromagnetic wave 16 to be observed is located far enough away from the antenna, the electromagnetic wave 16 arrives as a coherent plane wave. For simplicity, an example is shown in which the equipotential surface of the electromagnetic wave 16 is parallel to the first layer 11, i.e., the incident angle θ of the electromagnetic wave 16 to the antenna is 0°. The electromagnetic wave 16 is diffracted when it passes through the opening 13 in the first layer 11. Diffracted wave 17a from opening 13a, diffracted wave 17b from opening 13b, and diffracted wave 17c from opening 13c are detected by the power receiving elements 15a and 15b, respectively. The intensity of the electromagnetic waves received by the power receiving elements 15a and 15b is high only at a certain incident angle θ because the diffracted waves from the respective openings interfere with each other. When the spacecraft 1 is rotated and the incident angle θ is continuously changed, the intensity of the electromagnetic waves received by the power receiving elements 15a and 15b changes periodically. This is shown in Figure 2C. The position where the signal intensity shows a maximum is determined by the relationship between the incident angle θ and the attitude angle of the spacecraft 1. Therefore, the incident angle θ can be identified from the distribution where the signal intensity at the power receiving elements shows a maximum while the spacecraft 1 is rotated and the incident angle θ is continuously changed.

[0013] Here, when considering the case where the diffracted wave 17 from the aperture 13 is detected by the power receiving element 15, the signal strength S can be expressed as in (Equation 1).

[0014]

number

[0015] where A l =(a l ,0) is the position on the aperture plane, B m =(b m ,0) represents the position on the receiving element, k, θ, λ, and ω represent the wave number, incident angle, wavelength, and frequency of the electromagnetic wave, respectively, t represents time, and j represents the imaginary unit.

[0016] (Equation 1) is the component C that represents the interference fringes. m and the component T representing time fluctuations m These can be divided into (Equation 2) and (Equation 3), respectively.

[0017]

number

[0018]

number

[0019] (Equation 2) is the component C that represents the interference fringes. m is the opening position a l and receiving element position b m and the wavelength λ of the electromagnetic wave.

[0020] For this reason, multilayer interferometry has the following characteristics. First, whereas spatial resolution in rotational interferometry is a function determined by the distance D between the element antennas, multilayer interferometry allows for the use of measurement parameters such as the distance between the first and second layer films 11 and 14 (inter-layer distance δ), the position of the sensing surface, the film pattern, the arrangement of the two films, and the position and number of the receiving elements 15. In other words, when determining the direction of arrival of electromagnetic waves from an electromagnetic source on a spacecraft alone, the antenna section can be designed with various parameters without being restricted in spatial resolution by the distance D between the element antennas as in rotational interferometry, making it possible to set the desired spatial resolution. For example, the incident angle θ (see Figure 2C) at which signal strength is high varies depending on the inter-layer distance δ. Therefore, the antenna section can be optimized by designing the first and second layer films 11 and 14 according to the desired spatial resolution.

[0021] On the other hand, in multilayer interferometry, electromagnetic waves that pass through the aperture 13 are detected, so the signal strength of the electromagnetic waves that reach each individual receiving element 15 is weak. However, this problem can be easily solved by increasing the number of receiving elements. For example, if multiple receiving elements 15 are arranged so that the signal strength is maximized at the same incident angle, the outputs from those receiving elements 15 can be combined. Also, if the incident angles at which the signal strength of multiple receiving elements 15 is maximized differ depending on their placement, the timing of the outputs from each receiving element 15 can be adjusted using a phase shifter or delay circuit so that the signal strength of multiple receiving elements 15 at different placements is maximized at the same incident angle, and the outputs from the multiple timing-adjusted receiving elements 15 can be combined. This improves the signal strength output by the antenna unit.

[0022] 3 shows an example in which a power receiving element group 18 is arranged on the second layer film 14 instead of the power receiving element 15. In this case, as shown in FIG. 2C, when the incident angle θ of each power receiving element constituting one power receiving element group is continuously changed, the received signal strength changes periodically. However, the incident angle θ at which the signal strength peaks varies depending on the position of the power receiving element. In other words, the signal strength detected by each power receiving element constituting one power receiving element group differs, reflecting the interference pattern between diffracted waves.

[0023] Therefore, by comparing the magnitude of the signal strength of the power receiving elements 15al, 15ac, and 15ar that make up the power receiving element group 18a while rotating the spacecraft 1 and continuously changing the incident angle θ, it is possible to determine the direction of arrival of the electromagnetic waves from the relative signal strengths of the three power receiving elements and the attitude angle of the spacecraft 1. This makes it possible to further improve the spatial resolution.

[0024] (Electromagnetic Wave Observation System) Figure 4 shows a block diagram of an electromagnetic wave observation system that estimates the direction of arrival of electromagnetic waves. A spacecraft has various functions, but the electromagnetic wave observation system is a system that extracts the parts related to the interferometer measurement of this embodiment. The electromagnetic wave observation system includes an antenna unit 20, a processing unit 21, and a satellite control unit 25. As described above, the antenna unit 20 includes the first layer 11 with the shielding portion 12 and the aperture 13, and the second layer 14 with the power receiving element 15. The processing unit 21 includes a software-defined radio (SDR) 23 and a processor 24. The signal received by the power receiving element 15 is transmitted to the SDR 23 through a waveguide 22. The waveguide 22 can be a coaxial cable or a waveguide. The SDR 23 converts the transmitted analog signal into a digital signal, and the processor 24 determines the direction of the radio wave source from the acquired digital signal. The satellite control unit 25 includes an attitude detection device 26 that detects the current attitude and a rotation control device 27 that rotates the spacecraft 1. The attitude detection device 26 can use a star sensor, a sun sensor, an earth sensor, a magnetic sensor, an angular velocity sensor, etc. The rotation control device 27 can use a reaction wheel, a momentum wheel, a magnetic torquer, a thruster, etc. The antenna unit 20 and the processing unit 21 will now be described.

[0025] (Antenna part 20) FIG. 5 shows an example of the configuration of the first layer 11 and the second layer 14 that constitute the antenna unit 20. The film 31 in the first layer 11 corresponds to the shielding unit 12 (see FIG. 2A) and is a dielectric. For example, it may be a solar panel that generates operating power for the spacecraft 1. By using a solar panel as the shielding unit 12, the power consumed in controlling the rotation and attitude of the spacecraft 1 can be compensated for by using the power generated by the solar panel, thereby improving power efficiency. The film 31 need only be opaque to electromagnetic waves and may be either a reflector or an absorber. The film 31 has a slit 32 corresponding to the opening 13 and a patch antenna 33 (see FIG. 1). The film 34 in the second layer 14 is a metal film and may be, for example, the surface of the satellite casing 2 of the spacecraft 1. To detect electromagnetic waves that have passed through the slit 32, a group of power receiving elements is provided, with a slot antenna 35 as the power receiving element 15. The films 31 and 34 may be conductors or insulators, and a power receiving element appropriate for the material can be used. In this way, the antenna section 20 is composed of a gap that partially transmits electromagnetic waves or a shield that partially blocks them, and an antenna that receives the electromagnetic waves. By transmitting or reflecting radio waves with the first layer of antenna, the interference of diffracted waves can be measured with the second layer of antenna. Furthermore, the rigidity of the film is not limited. With a film of high rigidity, the shape of the opening is less likely to deform, making it possible to measure more stable interference signals.

[0026] 5 shows an example of a slit whose longitudinal direction is the Y direction as the opening 13, but the longitudinal direction may be oriented in any direction. In this case, it is desirable that the slot antenna 35 disposed in the second layer film 14 is disposed so that the change in intensity due to interference of diffracted waves of the electromagnetic wave transmitted through the slit 32 is reflected as strongly as possible. Therefore, for example, if the longitudinal direction of the slit 32 is disposed along the X direction, it is desirable to dispose the slot antenna 35 so that the longitudinal direction also follows the X direction.

[0027] The appropriate width of the slit can be determined depending on the wavelength of the incoming electromagnetic wave. The width of the slit can be adjusted according to the wavelength λ of the electromagnetic wave whose direction of arrival is to be detected. For example, the width can be set to about 0.2 times the wavelength λ. Furthermore, multiple slits of various widths can be provided in the antenna section. This makes it possible to detect the direction of arrival of electromagnetic waves with various wavelengths.

[0028] The opening 13 provided in the first layer film 11 will now be described. The opening 13 transmits the incoming electromagnetic waves and generates diffracted waves. Therefore, the opening 13 does not have to be a physically empty void as illustrated in FIG. 5; it may be filled with a material that transmits electromagnetic waves, such as glass. Conversely, even if there is a physically empty void, such as a metal mesh, if the electromagnetic waves are reflected and cannot be transmitted, it does not qualify as an opening in this embodiment. Furthermore, the opening may be an opening that transmits electromagnetic waves having the wavelength of the observation target but does not transmit electromagnetic waves having wavelengths other than the observation target. Furthermore, the opening 13 is not limited to a rectangular shape like the slit 32. For example, it may be a square or a triangular opening 36 as shown in FIG. 6. A pattern is generated by the interference of the electromagnetic waves transmitted through the slit-shaped openings on each side of the opening 36, and the spatial distribution of the interference pattern changes depending on the angle of incidence θ of the electromagnetic waves. Therefore, it can be used as the opening 13. In this way, if an interference pattern can be generated with one opening 36, only one opening may be provided in the film 31. The opening shown in FIG. 6 can also be considered as a combination of a plurality of slit-shaped openings with different longitudinal directions.

[0029] Regarding the positional relationship between the first layer 11 and the second layer 14, it is sufficient if the interference of the diffracted wave from the first layer 11 can be detected by the receiving element of the second layer 14. When there is a single receiving element, candidates for the incident angle of the electromagnetic wave can be estimated from the change in signal strength caused by rotating the spacecraft 1, and the incident angle can be uniquely detected by further rotation. When there are multiple receiving elements, candidates for the incident angle of the electromagnetic wave can be estimated based on the relative signal strength between the multiple receiving elements, and the incident angle of the electromagnetic wave can be uniquely detected from the strength of the signal strength of the receiving element measured when the spacecraft 1 is further rotated. Therefore, they do not need to be positioned parallel to each other as shown in Figure 2A. Furthermore, they may be arranged so that the diffracted wave from the first layer 11 is detected by the receiving element of the second layer 14 after being reflected by another film surface. Alternatively, the electromagnetic wave may be reflected by the side cross section of the opening 13 and transmitted through the opening 13.

[0030] (Processing unit 21) As shown in FIG. 4, an analog signal from the receiving element 15 is transmitted to the SDR 23 via the waveguide 22. Analog signals of the same phase (for example, the output signals of the receiving elements 15a and 15b in FIG. 4) may be combined in the waveguide 22. If necessary, the phase of the analog signal is adjusted using a phase shifter or a delay circuit. FIG. 7 shows an example of an equivalent circuit of the SDR 23. The equivalent circuit shown in FIG. 7 can extract, as its output, the real and imaginary parts of the time waveform of the received intensity of the electromagnetic wave detected by the receiving element 15 when expressed as a complex number.

[0031] The analog signal from the power receiving element 15 is band-limited by a band-pass filter 41 and then amplified by an amplifier 42. Meanwhile, a local oscillator 43 and a phase shifter 44 generate first and second high-frequency signals that are orthogonal to each other, and the first and second high-frequency signals are combined with the analog signal from the power receiving element 15 by mixers 45a and 45b, respectively. The combined signals pass through a low-pass filter 46, an amplifier 47, and a band-pass filter 48, respectively, and are converted into a digital signal by an analog-to-digital converter 49.

[0032] The processor 24 estimates the direction of arrival of the electromagnetic wave from the time waveform of the signal strength obtained for each receiving element 15 by the SDR 23. An example of a method for estimating the direction of arrival of the electromagnetic wave by the processor 24 is described below. The processor 24 functions as a functional unit that provides a predetermined function by executing processing in accordance with a program loaded into a main memory (not shown). In the following description, when describing processing by a program, the functional unit may be described as the main body, but the main body of the hardware in these cases is the processor 24.

[0033] First, an example of signal processing will be illustrated when the antenna unit 20 has the receiving element group shown in Fig. 3 on the second layer film 14. Fig. 8 is a functional block diagram of the process of estimating the direction of arrival of electromagnetic waves by the processor 24, and implements the method of estimating the incident angle θ exemplified in the explanation of Fig. 3. One receiving element group is assumed to have N receiving elements 50-1 to 50-N.

[0034] Analog signals from receiving elements 1 to N are input to SDR 23, where a time waveform is calculated. Here, the processing of SDR 23 described above is represented as time waveform calculation unit 51, and its processing corresponds to the processing of the equivalent circuit shown in Figure 7. Average power calculation unit 52 calculates the average power (signal strength) from the time waveform of each receiving element 1 to N. The average power is calculated using data for a period at least equal to or longer than the signal period. Comparison unit 53 determines in advance the dependence of the signal strength of receiving elements 1 to N on changes in incident angle θ, and estimates the incident angle θ based on the ranking or ratio of the magnitude of the average power of each receiving element 1 to N. Figure 9A shows an example of antenna unit 20 when N = 5. Figure 9B shows the dependence of signal strength on incident angle θ for each receiving element 50-1 to 5, simulated by setting predetermined parameters for the aperture length la, aperture spacing da, and inter-membrane distance δ of aperture 13. For example, when the power receiving element 50-3 has the maximum signal strength, it can be estimated that the incident angle θ is a value indicating one of the ranges A1, A2, and A3, and the incident angle θ is output as the median value of the ranges A1, A2, and A3. For the sake of explanation, let (A1, A2, A3) = (-14°, 0°, 14°).

[0035] Next, the spacecraft 1 is rotated to change the satellite attitude angle 55, and the incident angle θ is estimated in the same way. Here, consider the case where the arrival direction of the electromagnetic waves when the satellite attitude angle 55 = 0° is direction A2. In this case, even if the satellite attitude angle 55 is tilted to -14° or 14°, the signal strength becomes strong, but if it is tilted to 28°, the direction of the antenna unit makes it difficult to receive the electromagnetic waves, and the signal strength becomes weak. By changing the satellite attitude angle 55 in this way, the incident angle θ of the electromagnetic waves and the signal strength of the receiving element 50 also change.

[0036] In contrast, if we consider the case where the electromagnetic wave arrival direction is A3 when the satellite attitude angle 55 = 0°, the signal strength will be strong even if the satellite attitude angle 55 is tilted by 14° or 28°, but will be weak if tilted by -14°. In this way, by changing the satellite attitude angle 55, it is possible to narrow down the possible incident angles.

[0037] In this way, the angle estimation unit 54 estimates the direction of arrival of the electromagnetic waves based on the incident angle θ estimated by the comparison unit 53 and the satellite attitude angle 55 from the attitude detection device 26. In other words, it is possible to uniquely determine the direction of arrival of the electromagnetic waves from the estimated incident angle θ and the satellite attitude angle 55. Note that, although an example has been shown in which the average power is calculated from the time waveform of each of the power receiving elements 1 to N as the signal strength, the maximum amplitude value obtained from the time waveform of the signal may also be calculated as the signal strength.

[0038] FIG. 10A is a functional block diagram of the electromagnetic wave arrival direction estimation process by the processor 24, which implements another estimation method for the incident angle θ. FIG. 10B shows the configuration of the antenna unit 20 that performs the electromagnetic wave arrival direction estimation process shown in FIG. 10A. In this example, a receiving element (e.g., a patch antenna 33 (see FIG. 5)) provided on the first layer film 11 is used as the reference receiving element 50-R in the estimation process. The estimation method for the incident angle θ in this example will be explained using FIG. 10C. As shown in FIG. 10C, since the inclination of the equiphase surface 65 of the electromagnetic wave differs depending on the incident angle θ, the phase of the diffracted wave 17 generated at each aperture 13 changes depending on the incident angle θ. Therefore, the interference patterns generated by the receiving elements 50-1 and 50-2 differ depending on the incident angle θ.

[0039] The analog signal from the reference receiving element is input to the SDR 23, which calculates a time waveform. The time waveform, whose phase component has been inverted by the conjugation unit 61, is converted into a frequency spectrum by an FFT (Fast Fourier Transformation) 62-R. Meanwhile, the analog signals from receiving elements 1 to N are also input to the SDR 23, which calculates a time waveform for each. The time waveforms of receiving elements 1 to N are each converted into a frequency spectrum by the FFT 62, and are multiplied by the frequency spectrum of the reference receiving element, whose phase component has been inverted, to calculate the cross spectrum of the signal of the reference receiving element and each receiving element. The phase difference calculation unit 63 calculates the phase difference between the signal of the reference receiving element and the signal of each receiving element from the input cross spectrum. The angle estimation unit 54 estimates the incident angle θ from the phase difference at each receiving element calculated by the phase difference calculation units 63-1 to 63-N, and estimates the direction of arrival of the electromagnetic wave based on the satellite attitude angle 55 from the attitude detection device 26. By using a reference receiving element, the incident angle can be estimated from the phase difference rather than the signal strength, and the direction of arrival of the electromagnetic wave can be accurately estimated even for low-intensity signals whose strength is difficult to detect.

[0040] Fig. 11 is a functional block diagram of the electromagnetic wave arrival direction estimation process by processor 24, which estimates the change in the arrival direction of the electromagnetic wave over time using any power receiving element. The estimation method is as explained with reference to Fig. 10C. In Fig. 11, for the same power receiving element (power receiving element 1 in this example), delay unit 71 is used to calculate the phase difference based on the frequency spectrum from a predetermined time ago. This makes it possible to measure the phase difference according to the change in the incident angle, estimate the incident angle from the phase difference, and estimate the arrival direction of the electromagnetic wave.

[0041] 12 is a functional block diagram of the process of estimating the direction of arrival of electromagnetic waves by the processor 24 when an electromagnetic wave source emits incoherent electromagnetic waves. In the case of incoherent electromagnetic waves, interference from multiple diffracted waves causes the signal strength to be strong only at a specific incident angle θ. Therefore, the strength of the frequency spectrum of the electromagnetic waves can be determined by the strength determination unit 75, and the direction of the electromagnetic wave source can be estimated.

[0042] An example of mounting the electromagnetic wave observation system of this embodiment on a spacecraft 1 will be described below. Figure 13 shows a side view of the spacecraft 1, and Figure 14 shows a bird's-eye view of the spacecraft 1. Assume that the frequency of the electromagnetic waves to be observed is 2 GHz (observation wavelength λ = 15 cm), and the dimensions of the spacecraft 1 are, for example, D = 120 cm, d = 20 cm, I1 = 40 cm, and I2 = 20 cm. Regarding the beams 3, the beams whose one ends are connected to the satellite casing 2 after deployment are referred to as vertical beams, and the beams arranged perpendicular to the vertical beams are referred to as horizontal beams. For each vertical beam, a long horizontal beam is attached to the tip of the vertical beam, and a short horizontal beam is attached in the middle of the vertical beam. A triangular first layer film 11 is attached to the tip of the long horizontal beam, and a triangular second layer film 14 is attached to the tip of the short horizontal beam. Furthermore, a dipole antenna can be installed on the beam 3.

[0043] The interferometric measurement in this embodiment takes advantage of the spacecraft 1's three-dimensional structure and multilayer membrane. For example, the first membrane 11 can be a thin film attached to a beam, and the second membrane 14 can be the surface of the satellite housing 2. Alternatively, in a spacecraft with multiple membranes deployed, as shown in Figure 13, the first membrane 11 can be the outer membrane (the membrane farther from the satellite housing 2) and the second membrane 14 can be the inner membrane (the membrane closer to the satellite housing 2). This takes advantage of the three-dimensional structure of the spacecraft 1, generating diffracted radio waves using the outer membrane surface. The interference fringes can then be measured using an antenna installed inside. Multilayer detection using gaps and slits can improve the angular resolution of direction-of-arrival detection. Using a two-layer membrane allows the inter-membrane distance to be used as a parameter, thereby increasing the number of interferences. It also allows for measurements at more observation points. On the other hand, using the first membrane 11 as a thin film and the second membrane 14 as the surface of the satellite housing 2 can result in a spacecraft with a simpler three-dimensional structure.

[0044] FIG. 15 shows a flow chart for estimating the direction of a radio wave source using the electromagnetic wave observation system of this embodiment while controlling the attitude of the spacecraft 1. The processing unit 21 (processor 24) sets the measurement content (S01), and the rotation control device 27 rotates the spacecraft 1 around a rotation axis parallel to the normal to the first layer 11 (S02). The processor 24 acquires attitude information (satellite attitude angle 55) from the attitude detection device 26 and acquires the signal strength from the power receiving element 15 (S03). The processor 24 compares the signal strength with a preset threshold (S04). If the signal strength is lower than the threshold, the rotation control device 27 rotates the spacecraft 1 around the rotation axis parallel to the normal to the first layer 11 (S02). The reason for this will be explained using FIG. 16. If the opening 13 in the first layer 11 has a shape that is elongated in only one direction, such as a slot, and the polarization direction of the electromagnetic wave 16 is perpendicular to the longitudinal direction of the slot, the electromagnetic wave 16 cannot pass through the opening 13, resulting in extremely low signal strength. Therefore, the rotation control device 27 rotates the first layer 11 in a direction that increases the signal strength received by the power receiving element, for example, so that the polarization direction of the electromagnetic wave 16 and the longitudinal direction of the slot are aligned. This maximizes the electromagnetic wave that passes through the opening 13. If sufficient signal strength is obtained from the power receiving element, the arrival direction of the electromagnetic wave is estimated using the multilayer interferometry method described above (S06).

[0045] Furthermore, if the openings 13 in the first layer film 11 are not only slot-shaped openings aligned in the longitudinal direction, but also multiple slot-shaped openings with different longitudinal directions as shown in Figure 6, for example, are formed in the first layer film 11, it is possible to suppress the decrease in signal strength caused by the polarization direction, which has the effect of making the flow in Figure 15 unnecessary.

[0046] Figure 17 shows a flow chart for estimating the direction of a radio wave source using both rotational interferometry using a pair of dipole antennas and multi-layer interferometry using the antenna unit of this embodiment when a dipole antenna is mounted on the beam 3 of the spacecraft 1. The processor 24 sets the measurement content (S11) and estimates the direction of arrival of the electromagnetic wave using rotational interferometry (S12). In rotational interferometry, the pair of dipole antennas receive electromagnetic waves from an electromagnetic wave source while the rotation control device 27 rotates the spacecraft 1 around the axis of rotational symmetry. The direction of arrival of the electromagnetic wave is estimated using the interference fringes that result when the received electromagnetic waves are output as interference. In the case of the spacecraft 1 shown in Figures 13 and 14, the longest distance between the pair of dipole antennas is D. Therefore, when the observation wavelength λ is 15 cm, the spatial resolution (λ / D) in step S02 is approximately 7.2°. The processor 24 controls the attitude of the spacecraft 1 so that the antenna unit 20 faces the direction of arrival of the electromagnetic waves estimated in step S02 (S13), and estimates the direction of arrival of the electromagnetic waves using multilayer interferometry (S14). By appropriately setting the parameters of the antenna unit 20 (see FIG. 9A), the spatial resolution using multilayer interferometry can be made higher than the spatial resolution using rotational interferometry. In this way, by using multilayer interferometry as a high-precision measurement mode, the efficiency of measurements can be improved.

[0047] A modification of this embodiment will now be described.

[0048] (Variation 1) 18 shows a dielectric lens 81 provided at the opening of the first layer film 11. When the incident angle θ of the electromagnetic wave changes, the focal position of the dielectric lens 81 shifts. By utilizing this property, the incident angle θ can be estimated from the shift in the focal position.

[0049] (Variation 2) 3, in which multiple power receiving elements are arranged on the detection surface, Fig. 19 shows an example in which a power receiving element 82 that can move on the detection surface is used. A motor, for example, is used as a drive source for the power receiving element 82. By moving the power receiving element to perform measurements, it is no longer necessary to install multiple power receiving elements, which leads to reduced manufacturing costs.

[0050] (Variation 3) Although a spacecraft having a tensegrity structure has been exemplified as the spacecraft of this embodiment, the present invention is not limited to this. Figure 20 shows an example in which the first layer membrane 11 is positioned at a distance from the second layer membrane 14 using an extension structure such as a tether 83. This allows the distance to be freely and variably set using a tether extension / contraction mechanism or the like. In the case of a spacecraft without a tensegrity structure, the entire spacecraft may rotate, or a part of the spacecraft including the antenna may rotate.

[0051] (Variation 4) FIG. 21 shows an example of the configuration of an electromagnetic wave monitoring system including a spacecraft 1 described as this embodiment and a modified example. The spacecraft 1, moving in an orbit 101 around the Earth, receives electromagnetic waves 111 from an electromagnetic wave source 110 related to anthropogenic activities on Earth, electromagnetic waves 121 from an electromagnetic wave source 120 related to anthropogenic activities in outer space, and electromagnetic waves 131 from a celestial body 130, and also communicates with a ground station (base station) 90 installed on Earth (data downlink 91 and command uplink 92). The spacecraft 1 is deployed in various orbits, such as low earth orbit, medium earth orbit, geostationary orbit, and cislunar orbit, depending on the object of observation. The deployment shown here assumes a low earth orbit. The object of observation is all electromagnetic waves around the spacecraft 1, including the electromagnetic waves exemplified in FIG. 21.

[0052] FIG. 22 shows the configuration of the spacecraft 1 and the ground station 90. The ground station 90 includes a processing unit 97, a transmitting unit 98, and a receiving unit 99. The spacecraft 1 includes the attitude detection device 26 and the rotation control device 27 shown in FIG. 4 as well as the transmitting unit 28 and the receiving unit 29. The receiving units 29 and 99 demodulate the high-frequency signals transmitted from the other party and decode them into signals that can be processed by a processor. The transmitting units 28 and 98 also encode the signals to be transmitted to the other party and modulate the encoded signals into high-frequency signals. The encoding / decoding method and the modulation / demodulation method are not limited. In this configuration, the estimation of the direction of arrival of electromagnetic waves using the spacecraft 1 shown in the flow chart of FIG. 17 is performed by the following procedure. The signal strength of the electromagnetic waves detected by the receiving element of the spacecraft 1 and information regarding the attitude of the spacecraft are downlinked from the transmitting unit 28 to the receiving unit 99 of the ground station 90. In the ground station 90, a command for rotation control of the spacecraft 1 is uplinked from the transmitter 98 to the receiver 29 of the spacecraft 1 based on the signal strength of the electromagnetic waves received by the receiver 99. The spacecraft 1 observes the electromagnetic waves while rotating its body in accordance with the command. The spacecraft 1 then downlinks again from the transmitter 28 to the receiver 99 of the ground station 90 the signal strength of the electromagnetic waves detected by the receiving element and information on the attitude of the spacecraft. Then, the processor 97 of the ground station 90 estimates the arrival direction of the electromagnetic waves based on the signal strength of the downlinked electromagnetic waves and the attitude of the spacecraft. Note that the details of the process by the processor 97 to estimate the arrival direction of the electromagnetic waves are the same as the process performed by the processor 21 shown in Figure 4.

[0053] In this way, in the fourth modification, an example has been shown in which the process of estimating the direction of arrival of electromagnetic waves in the electromagnetic wave monitoring system is performed by the ground station 90. The processing unit 97 may be mounted on the spacecraft 1 side, and commands for the processing by the processing unit 97 may be uplinked from the ground transmitter 98 to the spacecraft 1.

[0054] The processor 24 of the spacecraft in the embodiment or the processor of the ground station in the fourth modification performs processing to determine the direction of arrival of electromagnetic waves in accordance with a program loaded into a main memory (not shown). The program includes the following programs:

[0055] (1) A program that causes a processor to estimate the angle of incidence of electromagnetic waves relative to the first layer film 11 from the signal strength of electromagnetic waves that have passed through and interfered with the first layer film 11, detected by an antenna unit that includes a first layer film 11 having multiple slit-shaped openings 13 in a film that is opaque to electromagnetic waves having the wavelength of the object to be observed, and a second layer film 14 having a receiving element 15.

[0056] (2) When multiple receiving elements 15 are provided on the second layer film 14, a program that causes the processor to estimate the incident angle of the electromagnetic waves with respect to the first layer film 11 based on the relative signal strength of the electromagnetic waves detected by the multiple receiving elements 15.

[0057] (3) A program that causes the processor to estimate the angle of incidence of the electromagnetic wave with respect to the first layer film 11 based on the change in the signal strength of the electromagnetic wave detected by the receiving element 15 while the spacecraft is rotated by the rotation control device 27 that rotates the spacecraft 1 and the angle of incidence is changed.

[0058] (4) A program that causes the processor to estimate the direction of arrival of the electromagnetic wave from the satellite attitude angle detected by the attitude detection device 26, which detects the attitude of the spacecraft 1 while rotating the spacecraft 1 using the rotation control device 27, and the estimated angle of incidence of the electromagnetic wave with respect to the first layer film 11.

[0059] (5) A program that causes the processor to issue a command to rotate the spacecraft 1 around a rotation axis parallel to the normal to the first layer film 11 using the rotation control device 27 when the signal strength of the electromagnetic wave detected by the receiving element 15 is smaller than a threshold value.

[0060] The present invention is not limited to the above-described embodiments and includes various modifications. For example, the above-described embodiments have been described in detail to make the present disclosure easier to understand, and are not necessarily limited to those including all of the described configurations. Furthermore, it is possible to replace part of the configuration of one embodiment or modification with the configuration of another embodiment or modification, and it is also possible to add the configuration of another embodiment or modification to the configuration of one embodiment or modification. Furthermore, it is possible to add, delete, or replace part of the configuration of each embodiment or modification with other configurations.

[0061] Furthermore, the functional units described as being executed by software programs may be partly or entirely implemented in hardware, for example, by designing them as integrated circuits. Information such as programs, tables, and files that implement each function may be stored in a memory, a recording device such as a hard disk or SSD (Solid State Drive), or a recording medium such as an IC card, memory card, or optical disk. [Explanation of symbols]

[0062] 1: Spacecraft, 2: Satellite housing, 3: Deployable beam, 4: Thin film, 5: Patch antenna, 11: First layer membrane, 12: Shielding part, 13: Aperture, 14: Second layer membrane, 15: Receiving element, 16: Electromagnetic wave, 17: Diffracted wave, 18: Receiving element group, 20: Antenna part, 21: Processing part, 22: Waveguide, 23: Software radio, 24: Processor, 25: Satellite control part, 26: Attitude detection device, 27: Rotation control device, 28: Transmitter part, 29: Receiver part, 31, 34: Membrane, 32: Slit, 33: Patch antenna, 35: Slot antenna, 36: Aperture, 41: Bandpass filter, 42: Amplifier, 43: Local oscillator, 44: Phase shifter, 45: Mixer, 4 6: Low-pass filter, 47: Amplifier, 48: Band-pass filter, 49: Analog-to-digital converter (ADC), 50: Receiving element, 51: Time waveform calculation unit, 52: Average power calculation unit, 53: Comparison unit, 54: Angle estimation unit, 55: Satellite attitude angle, 61: Conjugation unit, 62: FFT, 63: Phase difference calculation unit, 65: Equiphase surface, 71: Delay unit, 75: Intensity judgment unit, 81: Dielectric lens, 82: Receiving element, 83: Tether, 90: Ground station, 91: Downlink, 92: Uplink, 97: Processing unit, 98: Transmitter, 99: Receiver, 101: Orbit, 110, 120: Electromagnetic wave source, 111, 121, 131: Electromagnetic wave, 130: Celestial body.

Claims

1. an antenna unit including a first layer film in which a plurality of slit-shaped openings are provided in a film that is opaque to electromagnetic waves having a wavelength to be observed, and a second layer film in which a receiving element is provided that detects the electromagnetic waves that have passed through the plurality of openings and interfered; a processing unit that estimates the direction of arrival of the electromagnetic wave from the signal strength of the electromagnetic wave detected by the power receiving element.

2. In claim 1, a plurality of the power receiving elements are provided on the second layer film; The processing unit estimates an incident angle of the electromagnetic wave with respect to the first layer film based on relative signal strengths of the electromagnetic waves detected by the plurality of power receiving elements.

3. In claim 1, a rotation control device for rotating the spacecraft, The processing unit estimates the incident angle of the electromagnetic wave with respect to the first layer film based on a change in signal intensity of the electromagnetic wave of the power receiving element detected while rotating the spacecraft using the rotation control device to change the incident angle of the electromagnetic wave with respect to the first layer film.

4. In claim 1, a satellite housing; a plurality of beams deployed from the satellite housing; a spacecraft in which the first layer membrane and the second layer membrane are each a thin film stretched by the plurality of beams, or the first layer membrane is a thin film stretched by the plurality of beams and the second layer membrane is a surface of the satellite casing.

5. In claim 1, an attitude detection device for detecting the attitude of the spacecraft; a rotation control device for rotating the spacecraft, The processing unit estimates the direction of arrival of the electromagnetic wave from the satellite attitude angle detected by the attitude detection device and the estimated angle of incidence of the electromagnetic wave with respect to the first layer film while rotating the spacecraft using the rotation control device.

6. In claim 5, When the signal strength of the electromagnetic wave detected by the power receiving element is smaller than a threshold value, the processing unit causes the rotation control device to rotate the spacecraft around a rotation axis parallel to a normal to the first layer film.

7. In claim 5, a satellite housing; a plurality of beams deployed from the satellite housing; A dipole antenna is provided on the beam, The processing unit estimates an arrival direction of electromagnetic waves based on interference outputs of electromagnetic waves received by the pair of dipole antennas while rotating the spacecraft using the rotation control device, and controls the rotation control device to orient the first layer film in the estimated arrival direction of the electromagnetic waves.

8. In claim 1, A spacecraft in which the longitudinal direction of the opening and the longitudinal direction of the power receiving element are arranged along the same direction.

9. In claim 1, The plurality of openings provided in the first layer film include openings oriented in different longitudinal directions.

10. In claim 1, The first layer film is a solar panel for a spacecraft.

11. A ground station that transmits and receives signals to and from a spacecraft having an antenna unit that includes a first layer film in which a plurality of slit-shaped openings are provided in an electromagnetic wave-impermeable film, and a second layer film in which a receiving element is provided that detects electromagnetic waves that have passed through the openings and interfered with the first layer film, and estimates the direction of arrival of electromagnetic waves having a wavelength to be observed, a receiving unit that receives the signal intensity of the electromagnetic wave detected by the receiving element and downlinked from the spacecraft; a processing unit that estimates the direction of arrival of the electromagnetic wave from the signal strength of the electromagnetic wave received by the receiving unit; A ground station comprising:

12. In claim 11, a plurality of the power receiving elements are provided on the second layer film of the antenna portion of the spacecraft, The processing unit is a ground station that estimates the angle of incidence of the electromagnetic waves with respect to the first layer film based on the relative signal strengths of the electromagnetic waves detected by the plurality of power receiving elements.

13. In claim 12, the receiving unit receives a satellite attitude angle of the spacecraft, The processing unit is a ground station that estimates the direction of arrival of the electromagnetic wave from the satellite attitude angle and the estimated incident angle of the electromagnetic wave with respect to the first layer film.

14. In claim 11, a transmitter for uplinking commands to the spacecraft; When the signal strength of the electromagnetic waves detected by the power receiving element is smaller than a threshold, the processing unit transmits a command to the spacecraft to rotate the spacecraft around a rotation axis parallel to a normal to the first layer film.

15. a first layer film having a plurality of slit-shaped openings formed in a film that is opaque to electromagnetic waves; a second layer film provided with a receiving element for estimating an incident angle of the electromagnetic wave with respect to the first layer film based on the signal strength of the electromagnetic wave that has been transmitted through the plurality of openings and interfered; An antenna comprising:

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