Conformational spacecraft

Abstract

To provide a conformational spacecraft capable of constituting an omnidirectional space antenna capable of achieving both light weight and shape accuracy of a thin film surface, observing radio waves arriving from all directions, and identifying radio weave sources present in various directions and distances with high resolution.SOLUTION: This conformation spacecraft is constructed by expanding expansion-type beam members and tension members around the spacecraft from the form of the spacecraft that stores a plurality of expansion-type beam members and a plurality of tension members. The expansion-type beam members expanded from the state of being installed in the spacecraft are arranged, in a plurality of rotation symmetry axes of a virtual polyhedron constructed with the rough center of the spacecraft set as an original point, at roughly equal distances in the direction of the rotation symmetry axes. The tension members support two places of the respective ends of a nearest expansion-type beam member with tensile force. The respective ends of the expansion-type beam member are simultaneously supported by three or more tension members.SELECTED DRAWING: Figure 1

Description

[Technical field]

[0001] The present invention relates to a spacecraft having a three-dimensional structure. [Background technology]

[0002] Electromagnetic waves are emitted by the accelerated motion of electrons, and as they propagate, they are absorbed and scattered by various atoms, inducing the emission of new electromagnetic waves. Therefore, electromagnetic waves contain a variety of information, not only about electronic anthropogenic activities on Earth and in space, but also about the materials, activities, and formation processes of celestial bodies such as stars, planets, and nebulae. Therefore, observing electromagnetic waves and analyzing their radiation sources and propagation processes plays an extremely important role in understanding anthropogenic activities and the global environment, and in elucidating celestial activities and formation processes.

[0003] In recent years, constellations of low-orbit Earth observation satellites have been developed, and observation data with high frequency and high resolution has become available for visible light and electromagnetic waves of nearby wavelengths. Furthermore, in order to understand the state of radiation sources and propagation paths that are farther away or cannot be understood by visible light, observations of electromagnetic waves (hereinafter referred to as radio waves) with wavelengths longer than infrared are being made by satellites equipped with antennas.

[0004] Incidentally, in radio wave observations in space, it is necessary to identify radio wave radiation sources in various directions and distances with high resolution and analyze only the necessary information. Therefore, development of large-aperture space antenna technology that can achieve high resolution by making the aperture area of ​​the antenna deployed in space sufficiently large relative to the wavelength is being promoted around the world. In addition, a thin-film deployment structure that can reduce the space transportation cost by making the payload small and lightweight and can configure a space antenna with high shape precision is promising for large-aperture space antenna technology.

[0005] Patent Document 1 is known regarding a spacecraft equipped with an antenna. In Patent Document 1, for the purpose of obtaining a highly reliable deployable deployable antenna reflector for space equipment, it proposes "a deployable antenna that is composed of a metal mesh constituting the radio wave reflecting surface of the antenna, a network cable including a shape-retaining cable that retains the shape of the metal mesh, a back cable, and a tie cable that connects the shape-retaining cable and the back cable, and a plurality of deployment ribs and a deployment hinge that deploys the deployment ribs, and that deploys the metal mesh and the network cable from a folded state, and that after deployment, includes a plurality of deployment masts for holding the metal mesh and the network cable, and a tendon cable that connects the deployment masts, and that the deployment ribs and the network cables are arranged separately without being connected to each other, and that deploy independently when the deployment masts are deployed." [Prior art documents] [Patent documents]

[0006] [Patent Document 1] JP 2008-236500 A Summary of the Invention [Problem to be solved by the invention]

[0007] According to Patent Document 1, it is possible to obtain a deployable antenna reflector for space equipment that can be deployed with high reliability. On the other hand, the deployable antenna of Patent Document 1 has room for improvement in the configuration of a large aperture space antenna that achieves both light weight and shape accuracy of the thin film surface.

[0008] In other words, when constructing a large aperture antenna with high geometric precision in order to configure the antenna by extending a cantilevered deployment rib from a spacecraft, it is expected that the weight will increase because it is necessary to increase the cross-sectional area of ​​the deployment rib in order to increase the bending rigidity of the deployment rib.

[0009] There is also room for improvement in the configuration of an omnidirectional space antenna that has directivity in all directions. A planar antenna is constructed by extending a cantilevered deployment rib from the spacecraft in a plane, which gives it high directivity in the perpendicular direction to the plane, but does not realize an omnidirectional antenna that has directivity in all directions. Therefore, it is not possible to identify radio wave radiation sources that exist in all directions with high resolution.

[0010] In view of the above, the present invention aims to provide a three-dimensional spacecraft that combines light weight with precision in the shape of the thin film surface, and that can form an omnidirectional space antenna that can observe radio waves coming from all directions and identify radio wave sources in various directions and at various distances with high resolution. [Means for solving the problem]

[0011] In view of the above, the present invention is defined as "a three-dimensional spacecraft configured by deploying the deployable beam members and force members around the spacecraft from a spacecraft configuration that stores a plurality of deployable beam members and a plurality of tension members, the deployable beam members being deployed from a state provided in the spacecraft, the deployable beam members being arranged at approximately equal distances in the direction of the rotational symmetry axes of a virtual polyhedron configured with the approximate center of the spacecraft as the origin, the tension members supporting two points of each end of the nearest deployable beam members by tension, and each end of the deployable beam members being supported simultaneously by three or more of the tension members." Effect of the Invention

[0012] According to the present invention, it is possible to provide a three-dimensional spacecraft that combines light weight with precision in the shape of the thin film surface, and that can form an omnidirectional space antenna that can observe radio waves coming from all directions and identify radio wave sources in various directions and at distances with high resolution. [Brief description of the drawings]

[0013] [Figure 1] FIG. 1 is a diagram showing an example of the configuration of a three-dimensional spacecraft according to a first embodiment of the present invention. [Diagram 2] FIG. 6 is a diagram showing a configuration example of a three-dimensional spacecraft according to a second embodiment of the present invention. [Diagram 3] FIG. 11 is a diagram showing a configuration example of a three-dimensional spacecraft according to a third embodiment of the present invention. [Figure 4] FIG. 2 is a diagram showing the deployment process from an icosahedral spacecraft 3 to a three-dimensional spacecraft 10. [Diagram 5] FIG. 2 is a diagram showing the deployment process from a tetrahedral spacecraft 3 to a three-dimensional spacecraft 10. [Figure 6] FIG. 2 is a diagram showing the deployment process from an octahedral spacecraft 3 to a three-dimensional spacecraft 10. [Figure 7] FIG. 5 is a cross-sectional view of the three-dimensional structure shown in FIG. 4, taken through the center of the spacecraft 3. [Figure 8] FIG. 4 is a cross-sectional view of the storage section 8 taken along the line D-D. [Figure 9] FIG. 2 is a diagram for explaining how the deployable beam member 1 can be provided with an antenna function. [Figure 10] A diagram showing the payload configuration of a spacecraft. [Figure 11] FIG. 1 is a diagram showing an example of a directional pattern of a dipole antenna array. [Figure 12] FIG. 1 is a diagram showing the configuration of a directional pattern of a circularly polarized wave. [Figure 13] FIG. 8 is a diagram showing a thin film 21 provided on each face of the icosahedron of FIG. 7. [Figure 14] FIG. 2 is a diagram showing an example of an antenna structure in which a deployable beam member 1 has a plurality of antenna elements. [Figure 15] 1A to 1C are diagrams showing examples of rotation axes in various multi-faceted three-dimensional spacecraft 10. [Figure 16] 1 is a diagram showing the observed globe as the spacecraft 3 rotates along its axis of rotation. [Figure 17] A diagram showing the relationship between the number of resolvable electromagnetic wave sources when a 6-channel array antenna is used. [Figure 18] Flowchart for omnidirectional electromagnetic wave observation using Spacecraft 3. [Figure 19] A diagram showing an example of a system configuration for omnidirectional electromagnetic wave observation. [Figure 20]FIG. 13 is a diagram showing an example in which a photon energy absorbing portion that absorbs photons is formed in a thin film portion. [Figure 21] FIG. 13 is a diagram showing an example in which a photon energy reflecting portion that reflects photons is configured in a thin film portion. [Figure 22] FIG. 13 is a diagram showing an example in which a photon energy reflecting portion or a photon energy transmitting portion that reflects or transmits photons is configured in a thin film portion. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0014] Hereinafter, embodiments of the present invention will be described with reference to the drawings.

[0015] In the following description of the embodiments, various configuration examples, deployment methods, antenna configurations, and others of a three-dimensional spacecraft are broadly described. First, in embodiments 1 to 3, various configuration examples of a three-dimensional spacecraft are described. In embodiment 4, deployment from a spacecraft to a three-dimensional spacecraft is described. In embodiment 5, configuring an antenna function in a three-dimensional spacecraft is described. In embodiment 6, the meaning of rotation about the rotational symmetry axis 5 of the spacecraft 3 is described. In embodiment 7, various functions realized by using a thin film are described. EXAMPLES

[0016] In the present invention, the three-dimensional spacecraft is configured as a virtual regular polyhedron with the approximate center of the spacecraft as the origin. As an example, in the first embodiment, the three-dimensional spacecraft is configured as a virtual regular tetrahedron with the approximate center of the spacecraft as the origin. In the following description, the term "regular polyhedron" is used, but this does not necessarily mean that the spacecraft is "regular" in the strict sense, as long as it is a polyhedron.

[0017] Fig. 1 is a diagram showing an example of the configuration of a three-dimensional spacecraft according to a first embodiment of the present invention. The three-dimensional spacecraft 10 in Fig. 1 includes a plurality of deployable beam members 1, tension members 2, and a spacecraft 3, and the deployable beam members 1 and tension members 2 are used to form a virtual regular tetrahedral three-dimensional structure 10 around the spacecraft 3.

[0018] In order to configure a virtual regular tetrahedron, four sets of deployable beam members 1, 1A, 1B, 1C, and 1D, are used and are arranged at approximately equal distances in the direction of multiple rotational symmetry axes 5A, 5B, 5C, and 5D of a virtual regular polyhedron configured with the approximate center of the spacecraft 3 as the origin O. As a result, the four sets of deployable beam members 1A, 1B, 1C, and 1D are arranged pointing in the vertex directions of the virtual regular tetrahedron, and the ends 7A, 7B, 7C, and 7D of each of the deployable beam members 1A, 1B, 1C, and 1D are positioned at the vertices of the virtual regular tetrahedron.

[0019] To form an imaginary regular tetrahedron, six tension members 2, 2A, 2B, 2C, 2D, 2E, and 2F, are used to support by tension two points at each end 7A, 7B, 7C, and 7D of the nearest expandable beam members 1A, 1B, 1C, and 1D. As a result, each end 7A, 7B, 7C, and 7D of the expandable beam members is simultaneously supported by three or more tension members 2A, 2B, 2C, 2D, 2E, and 2F.

[0020] By arranging them as described above, a three-dimensional spacecraft 10 is formed with the spacecraft 3 as the central structure. The deployable beam member 1 is initially assembled integrally with the spacecraft 3, and then deployed in the direction shown in the figure. This deployment will be described in the following examples.

[0021] The outline of the three-dimensional spacecraft 10 of the first embodiment has been described above, but will be described in more detail below. In the first embodiment, a regular tetrahedron is assumed as a virtual regular polyhedron configured with the approximate center of the spacecraft 3 as the origin O. There are multiple rotational symmetry axes 5 of the regular tetrahedron, but here, four rotational symmetry axes from the approximate center to the vertices of the regular tetrahedron are considered. The four deployable beam members 1 stored in the spacecraft 3 deploy at approximately equal distances toward the respective rotational symmetry axes 5, and the tension members 2 support two points of each end 7 of the nearest deployable beam member 1 by tension, and each end 7 of the deployable beam member 1 is supported by the three tension members 2 simultaneously, so that a compressive axial force along the rotational symmetry axes 5 acts on the deployable beam members 1. In other words, the positions of the deployable beam member 1 and the tension member 2 are arranged so that the compressive force from the deployable beam member 1 and the tension from the tension member 2 are balanced, and an approximately regular tetrahedral three-dimensional structure 10 is constructed around the spacecraft 3, having four omnidirectional openings formed by the three tension members 2.

[0022] Here, when considering the stability of a three-dimensional structure using Maxwell's law of structural stability, expanded by Calladine (1978), the number of compression members r and the number of tension members t required for structural stability of n nodes in three-dimensional space are given by the relationship in the following equation (1).

[0023]

number

[0024] In Example 1, the number of nodes n is the number of spacecraft 3 (1) and the number of ends 7 of deployable beam member 1 (4), totaling n=5, and the sum of the numbers of compression members and tension members required for this structure to be stable is 9. In contrast, the deployable beam member 1, which is a compression member, has r=4, and the effective number of tension members is t=6, which is the number of sides of a regular tetrahedron. Therefore, since the left side of equation (1) (r+t=10) is greater than the right side (3n-6=9), it can be concluded that the structure is stable.

[0025] Furthermore, by making the tension members 2 corresponding to the edges of the regular tetrahedron light-weight wires or cables having a tension function, the tension members 2 can be configured to be lighter than the deployable beam members 1. When a three-dimensional structure is configured using only compression members based on a conventional space structure, nine compression members are required according to this embodiment, but only four compression members are required according to the first embodiment of the present invention. Therefore, the spacecraft according to the first embodiment of the present invention is lighter than the conventional space structure.

[0026] In this way, according to the first embodiment of the present invention, it is possible to configure a spacecraft with a three-dimensional structure that is lightweight and has a high degree of shape accuracy and has large openings in all directions. EXAMPLES

[0027] In the present invention, the three-dimensional spacecraft is configured as a virtual regular polyhedron with the approximate center of the spacecraft as the origin. As an example, in the second embodiment, the three-dimensional spacecraft is configured as a virtual regular octahedron with the approximate center of the spacecraft as the origin.

[0028] Fig. 2 is a diagram showing a configuration example of a spacecraft with a three-dimensional structure according to a second embodiment of the present invention. The difference in configuration from Fig. 1 is the number of deployable beam members 1 and tension members 2, which is due to the use of a regular octahedron, but the basic concept of Fig. 1 is followed.

[0029] In addition, in relation to the fact that the spacecraft is formed into a regular octahedron in FIG. 2, the axis of rotational symmetry 5 is three perpendicular orthogonal axes 6 (6A, 6B, 6C), and the deployable beam members 1 are arranged at approximately equal distances in the positive and negative directions of the axis of rotational symmetry 5, which is formed with the approximate center of the spacecraft 3 as the origin O, with the first deployable beam members 1A1, 1B1, 1C1 and the second deployable beam members 1A2, 1B2, 1C2 being positioned at approximately equal distances in the positive and negative directions of the axis of rotational symmetry 5, thereby forming a three-dimensional spacecraft 10.

[0030] In the second embodiment, a regular octahedron is assumed as a virtual regular polyhedron with the approximate center of the spacecraft 3 as the origin O. There are multiple rotational symmetry axes of a regular tetrahedron, but here, three orthogonal rotational symmetry axes 6 extending from the approximate center to the faces of the regular octahedron are considered. The six deployable beam members 1 stored in the spacecraft 3 deploy at approximately equal distances in the positive and negative directions of the respective rotational symmetry axes 6, and the tension members 2 support two points of each end 7 of the nearest deployable beam member 1 by tension, and each end 7 of the deployable beam member 1 is supported simultaneously by the four tension members 2, so that a compressive axial force along the rotational symmetry axis 5 acts on the deployable beam member 1. In other words, the positions of the deployable beam members 1 and the tension members 2 are arranged so that the compressive force by the deployable beam members 1 and the tension by the tension members 2 are balanced, and a three-dimensional structure of an approximately regular octahedron having eight omnidirectional openings formed by the three tension members 2 is configured around the spacecraft 3.

[0031] The stability of this three-dimensional structure is examined using the extended Maxwell's law of equation (1). In Example 2, the number of nodes n is the number of spacecraft 3 (1) and the number of ends 7 of deployable beam member 1 (6), totaling n=7, and it can be seen that the sum of the numbers of compression members and tension members required for this structure to be stable is 15. In contrast, the deployable beam member 1, which is a compression member, has r=6, and the effective number of tension members is t=12, which is the number of sides of a regular octahedron. Therefore, since the left side of equation (1) (r+t=18) is greater than the right side (3n-6=15), it can be concluded that the structure is stable.

[0032] Furthermore, by using lightweight wires having a tension function for the tension members 2 that correspond to the edges of the regular octahedron, the tension members 2 can be configured to be lighter than the deployable beam members 1. When a three-dimensional structure is configured using only compression members based on a conventional space structure, 15 compression members are required, but according to the second embodiment of the present invention, only six compression members are required. Therefore, the spacecraft 3 according to the second embodiment of the present invention is lighter than the conventional space structure.

[0033] In this way, according to the present invention, it is possible to configure a spacecraft that is both lightweight and has high shape accuracy and has large openings in all directions. EXAMPLES

[0034] In the present invention, the three-dimensional spacecraft is configured as a virtual regular polyhedron with the approximate center of the spacecraft as the origin. As an example, in the third embodiment, the three-dimensional spacecraft is configured as a virtual regular icosahedron with the approximate center of the spacecraft as the origin.

[0035] Figure 3 is a diagram showing a configuration example of a spacecraft with a three-dimensional structure according to a third embodiment of the present invention. The difference in configuration from Figures 1 and 2 is that the number of deployable beam members 1 and tension members 2 is different because of the use of a regular icosahedron, but the basic concept of Figure 1 is followed.

[0036] In addition, in relation to the fact that the spacecraft is formed into a regular icosahedron in FIG. 3, the axis of rotational symmetry 5 is three perpendicular orthogonal axes 6 (6A, 6B, 6C), and the first deployable beam members 1A1, 1B1, 1C1 and the second deployable beam members 1A2, 1B2, 1C2 are arranged at approximately equal distances in the positive and negative directions on the axis of rotational symmetry 5, which is defined with the approximate center of the spacecraft 3 as the origin O, so that their longitudinal directions are perpendicular to the axis of rotational symmetry 5, thereby forming a three-dimensional spacecraft 10.

[0037] In the third embodiment, a regular icosahedron is assumed as a virtual regular polyhedron with the approximate center of the spacecraft 3 as the origin O. Although there are multiple rotational symmetry axes of the regular icosahedron, three orthogonal rotational symmetry axes 5A, 5B, and 5C extending from the approximate center toward the faces of the regular hexahedron are considered here. The six deployable beam members 1 (six members: first deployable beam members 1A1, 1B1, and 1C1 and second deployable beam members 1A2, 1B2, and 1C2) stored in the spacecraft 3 are arranged in the positive and negative directions of the respective rotational symmetry axes 5A, 5B, and 5C, and deploy in the directions perpendicular to the rotational symmetry axes 5A, 5B, and 5C. The tension members 2 support the closest two ends 7 of the expandable beam member 1 with tension, and as each end 7 of the expandable beam member 1 is supported simultaneously by four tension members, a compressive axial force acts on the expandable beam member 1 in a direction perpendicular to the axes of rotational symmetry 5A, 5B, 5C.

[0038] That is, the positions of the deployable beam members 1 and the tension members 2 are arranged so that the compressive force of the deployable beam members 1 and the tension of the tension members 2 are balanced, and a three-dimensional structure having 14 omnidirectional openings formed by the tension members 2 is configured around the spacecraft 3. When the deployable beam members 1 arranged in the positive and negative directions of the rotational symmetry axes 5A, 5B, and 5C are brought closer to each other, the three-dimensional structure asymptotically approaches a regular octahedron, and when they are moved away from each other, the three-dimensional structure asymptotically approaches a cubic octahedron. In this case, a three-dimensional structure that is approximately a Jessen icosahedron in which the relationship between compressive force and tension is balanced is configured.

[0039] The stability of this three-dimensional structure is examined using the extended Maxwell's law of equation (1). In this embodiment, the number of nodes n is 12, which is the number of ends 7 of the expandable beam member 1, and it can be seen that the sum of the number of compression members and tension members required for this structure to be stable is 30. In contrast, the expandable beam member 1, which is a compression member, has a length of r=6, and the effective number of tension members is t=24, which is the number of sides of a regular octahedron. Therefore, since the left side of equation (1), r+t=30, is equal to the right side, 3n-6=30, it can be concluded that the structure is structurally stable.

[0040] Furthermore, by using lightweight wires having a tension function for the tension members 2 that correspond to the edges of the Jessen icosahedron, the tension members 2 can be configured to be lighter than the deployable beam members 1. When a three-dimensional structure is configured using only compression members based on a conventional space structure, 30 compression members are required, but according to the third embodiment of the present invention, only six compression members are required. Therefore, the spacecraft according to the third embodiment of the present invention is lighter than the conventional space structure.

[0041] In this way, according to the third embodiment of the present invention, it is possible to configure a spacecraft that is both lightweight and has high shape accuracy, and has three orthogonal axes and a large opening in all directions. EXAMPLES

[0042] In the present invention, the spacecraft 3 is deployed into the three-dimensional spacecraft 10. In the fourth embodiment, the deployment process of each of the polyhedrons will be described.

[0043] First, Figure 4 shows the deployment process from an icosahedral spacecraft 3 to a three-dimensional spacecraft 10. The left side of the figure shows the state in which the tension members 2 and deployable beam members 1 are assembled and integrated into the spacecraft 3, and from this state the tension members 2 and deployable beam members 1 are deployed one after another to finally form a three-dimensional spacecraft 10 made of the tension members 2 and deployable beam members 1 around the spacecraft 3. Note that in a later step, thin films will be placed on each surface formed by the tension members 2 and deployable beam members 1, but at this stage the thin films have not been placed and only the framework has been deployed.

[0044] Regarding this deployment, we assume that the spacecraft 3 is a cube-shaped CubeSat. By using a shape memory alloy or a convex tape material that exerts a restoring force when bent for the deployable beam member 1, the deployable beam member 1 can be stored in a cube shape along the spacecraft 3 in the initial stage (pre-deployment stage) on the left side of Fig. 4.

[0045] When stored, the tension member 2 is bent and no tension is applied. That is, in the initial stage of deployment, the deployable beam member 1 exerts its own restoring force to deploy the device. As the device expands, the tension member 2 is stretched and applies tension to the end 7 of the deployable beam member 1.

[0046] Ultimately, the positions of the deployable beam member 1 and the tension member 2 are arranged so that the compressive force of the deployable beam member 1 and the tension force of the tension member 2 are balanced. This allows the spacecraft to be stored in a form that is easy to store in a rocket, making it possible to construct a spacecraft with a large opening in all directions in an economical form that does not incur transportation costs.

[0047] Next, Fig. 5 shows the deployment process from a tetrahedral spacecraft 3 to a three-dimensional spacecraft 10. Fig. 6 shows the deployment process from an octahedral spacecraft 3 to a three-dimensional spacecraft 10. In Figs. 5 and 6, the left side shows the spacecraft 3 in an integrated state with tension members 2 and deployable beam members 1, from which the tension members 2 and deployable beam members 1 are deployed one after another, until the three-dimensional spacecraft 10, which is a tetrahedron and an octahedron, is formed around the spacecraft 3 by the tension members 2 and deployable beam members 1. Figs. 5 and 6 show the extension in the direction of the axis of rotational symmetry.

[0048] 5 and 6, it is assumed here that the spacecraft 3 is a cube-shaped CubeSat. The deployable beam member 1 can be stored inside the spacecraft 3 by configuring it to expand and contract in the direction of the rotation axis 5. Here, for example, by inserting an axial spring into the deployable beam member 1, it is possible to configure the deployable beam member 1 to have a restoring force in the extension direction. When stored, the tension member 2 is bent and no tension is applied.

[0049] That is, in the initial stage of deployment, the deployable beam member 1 deploys due to its own restoring force. As the deployment progresses, the tension member 2 is stretched, applying tension to the end 7 of the deployable beam member 1.

[0050] Ultimately, the positions of the deployable beam member 1 and the tension member 2 are arranged so that the compressive force of the deployable beam member 1 and the tension force of the tension member 2 are balanced. This allows the spacecraft to be stored in a form that is easy to store in a rocket, making it possible to construct a spacecraft with a large opening in all directions in an economical form that does not incur transportation costs.

[0051] A specific deployment mechanism proposal for the deployment of the icosahedron in Fig. 4 will be explained using Fig. 7 and Fig. 8. First, Fig. 7 is a diagram showing the three-dimensional structure diagram in Fig. 4 in a cross section passing through the center of the spacecraft 3, and according to this diagram, the deployable beam member 1 that deploys in a direction perpendicular to the axis of rotational symmetry 5 is stored in a storage section 8 (8C1 in the figure) provided in the spacecraft 3 in Fig. 7. The D-D cross section of the storage section 8 is illustrated in Fig. 8.

[0052] In the structure shown in Fig. 8, the expandable beam members 1 are spirally wound like mosquito coils and stored in the storage section 8, and two sets of expandable beam members 1 extending in the positive and negative directions on the same straight line are stored as a pair so that they extend the same distance at the same time when extended. In this figure, 10 is a shaft, 11 is a roller, and 9 is a guide roller.

[0053] Here, the deployable beam member 1 is stored by being wound around the roller 11. The deployable beam member 1 is formed of, for example, a convex tape material so that a restoring force acts against such bending. When deployed, in FIG. 8, the roller 11 rotates counterclockwise and the deployable beam member 1 extends in the longitudinal direction. By deploying the deployable beam member 1 using its own restoring force and the tension from the tension member 2, the positions of the deployable beam member 1 and the tension member 2 are finally arranged so that the compressive force from the deployable beam member 1 and the tension from the tension member 2 are balanced. By storing the deployable beam member 1 in a form that is easy to store in a rocket, the spacecraft 3 having a large opening in all directions can be constructed in an economical form that does not incur transportation costs. EXAMPLES

[0054] In the fifth embodiment, a description will be given of configuring an antenna function in the three-dimensional spacecraft 10.

[0055] Fig. 9 explains how to provide an antenna function to the deployable beam member 1. Here, a cross section of the deployable beam member 1 is shown, and a power feed point 13 and dipole antennas 12 (12A, 12B, 12C) are arranged within this cross section. Multiple sets of these antennas are prepared according to the wavelength λ of the electromagnetic waves to be captured.

[0056] In relation to the arrangement of the antenna in FIG. 9, the deployable beam member 1 has a certain length because it is necessary to configure a large three-dimensional structure 10 around the spacecraft 3. In addition, since it has a cross-sectional area large enough not to buckle, a conductive antenna element 12 can be installed along the deployable beam member 1 as shown in the figure. For example, as shown in the example, a dipole antenna can be configured by setting the deployment center as a feeding point 13 and installing conductive antenna elements 12 in the extension direction to the left and right. It is known that the wavelength λ of the resonant frequency of a dipole antenna is approximately half the length of the dipole antenna, and by providing dipole antenna elements having multiple lengths on the deployable beam member 1 as shown in the figure, an antenna 12 corresponding to multiple frequencies can be configured on the deployable beam member 1.

[0057] When a spacecraft is equipped with multiple antennas corresponding to such multiple frequencies, the payload configuration of the spacecraft can be configured as shown in the following Figure 10. Here, to configure a six-channel array antenna, six deployable beam members 1 (six members consisting of the first deployable beam members 1A1, 1B1, 1C1 and the second deployable beam members 1A2, 1B2, 1C2) in Figure 2 or Figure 3 are used, and a group of three dipole antennas, 12A, 12B, and 12c in the illustrated example, are arranged on each of these deployable beam members 1. In this figure, 14 and 15 represent the traveling direction and wavefront of the electromagnetic wave.

[0058] In contrast, the payload configuration 20 inside the spacecraft 3 includes a switch 16 for changing over signals received by a 6-channel array antenna, an amplifier 17, a software radio 18, a processor 19 for signal processing, and the like.

[0059] Fig. 10 shows the payload configuration inside the spacecraft 3 as a block diagram. The multiple dipole antennas 12 provided on the deployable beam members 1 can be switched for use by a frequency-changing RF (radio frequency) switch 16. If the number of deployable beam members 1 is six, the number of antennas that can be used simultaneously is six. The electromagnetic wave signals observed by the six dipole antennas 12 are converted into electric signals, amplified by an amplifier 17, and then received by, for example, a software radio 18. Each software radio 18 samples synchronously, and the traveling direction of the electromagnetic waves can be obtained by, for example, performing correlation processing on a processor.

[0060] Furthermore, when the six dipole antennas 12 are synchronized as described above, they can be used as a six-channel dipole antenna array. If the distance between the antennas is sufficiently small compared to the wavelength, the directional pattern of the dipole antenna array is expressed as shown in Figure 11 below.

[0061] Each dipole antenna 12 has a doughnut-shaped directional pattern centered on the axis, but in this embodiment, the spacing between each dipole antenna is short compared to the wavelength, so when the dipole antennas are arranged in an array, interference results in a directional pattern as shown in Figure 11.

[0062] In the case of a 6-channel dipole antenna array, since there are axes of symmetry in four directions, four directional patterns can be configured as shown in Fig. 11 (a), (b), (c), and (d) depending on how the reference is taken. Since (a), (b), (c), and (d) each have different antenna null directions, i.e., directions with low reception sensitivity, the null direction can be selected according to the reception target. In this way, according to the present invention, an omnidirectional antenna directional pattern can be arbitrarily configured.

[0063] In FIG. 11, the axis of rotational symmetry, the null receiving direction, the deployable beam member 1 that constitutes the dipole antenna, the directional pattern of the dipole antenna, the distance between the antennas, etc. are displayed three-dimensionally to make it easier to understand their mutual relationships.

[0064] Furthermore, in the above-mentioned 6-channel dipole antenna array, the axis of symmetry is C3, i.e., it has a rotational symmetry of 120 degrees, so by receiving or transmitting with a 120 degree shift, it is possible to configure a circularly polarized directional pattern as shown in Figure 12 below.

[0065] Similarly, in this example, four circularly polarized wave patterns (a), (b), (c), and (d) can be configured depending on the reference. (a), (b), (c), and (d) each have a different receiving direction with high receiving sensitivity, so the direction can be selected according to the receiving target. In this way, according to the present invention, a spacecraft can be configured that can radiate or receive electromagnetic waves of various frequency ranges and directivities in all directions.

[0066] In Fig. 9, an example in which the deployable beam members 1 are used to realize the function of an antenna is explained, but this can be further realized as shown in Fig. 13. In the explanation up to this point, it has been explained that the three-dimensional spacecraft 10 is made into a polyhedral three-dimensional spacecraft 10 by forming a framework of the deployable beam members 1 and tension members 2 around the spacecraft 3, but in reality, in a later step, a thin film is provided on each surface formed by the deployable beam members 1 and tension members 2 to make it into a substantially planar structure.

[0067] In the next example of antenna installation, an antenna is formed on the surface made of this thin film. Figure 13 shows a case in which a thin film 21 is installed on each surface of the surface shown in Figure 7, and a patch antenna 22 is further provided on the thin film 21.

[0068] According to the configuration example of the skeleton structure in Fig. 7, a structurally stable thin film 21 can be provided on a plane formed by tension members 2 as shown in Fig. 13, so that a patch antenna array can be configured in which thin film patch antennas 22 are laid out. By synchronizing the phase of such a patch antenna array, a directivity pattern with sharp directivity as shown in Fig. 13 can be obtained. In this configuration, since eight aperture planes are provided in all directions, it is possible to configure a transmitting or receiving antenna with sharp directivity in all directions.

[0069] Furthermore, Fig. 14 shows an example of an antenna structure in which multiple antenna elements are provided on the deployable beam member 1. As shown on the left side of the AA cross section in the right diagram of Fig. 14, multiple dipole antenna elements 12 are installed in the longitudinal direction of the deployable beam member 1.

[0070] 14, the deployable beam member 1 has a certain length because it is necessary to configure a large three-dimensional structure around the spacecraft 3. In this embodiment, a dipole antenna array is configured on the deployable beam member 1 by installing multiple dipole antenna elements 12 in the longitudinal direction of the deployable beam member 1 with a certain inter-element distance.

[0071] According to this configuration, by arranging the inter-element distance to be sufficiently small compared to the wavelength, it is possible to configure a pattern with sharp directivity for reception and transmission.

[0072] The above-mentioned later-stage three-dimensional spacecraft 10 is not a simple framework, but has a thin film formed on the surface that forms a polyhedron. Therefore, this structure has a thin film on a plane formed by three or more tension members 2 connected via each end 7 of the deployable beams 1, and an antenna can be installed on one or more of the deployable beams 1, tension members 2, and the thin film, making it possible to radiate or receive electromagnetic waves. EXAMPLES

[0073] In the sixth embodiment, the meaning of rotation about the axis of rotational symmetry 5 of the spacecraft 3 will be described. The three-dimensional spacecraft 10 described above receives electromagnetic waves by rotating about the axis of rotational symmetry 5.

[0074] The spacecraft 3 receives and analyzes electromagnetic waves, including rotation information, by rotating along multiple rotational symmetry axes 5, and downlinks the analysis results to a ground station, where the ground station can investigate changes in analysis patterns according to the orbital conditions and rotation information of the spacecraft, making it possible to apply the system to identifying electromagnetic wave characteristics and electromagnetic wave source positions. Also, if the spacecraft 3 has sufficient computational resources, the electromagnetic wave characteristics and electromagnetic wave source positions may be identified in orbit.

[0075] FIG. 15 shows examples of rotation axes in the various multi-faceted three-dimensional spacecraft 10 described so far. As shown here, there are multiple rotation axes. For example, when the array antenna 12 relating to the icosahedral three-dimensional structure shown in FIG. 3 is constructed around the spacecraft 3, it has 13 rotational symmetry axes. Here, Cn (n is an integer of 2 or more) represents an operation for a rotation of 2π / n [rad] around the axis of the arrow shown in FIG. 15. In this way, since the array antenna 12 is constructed to have high rotational symmetry, electromagnetic wave observation can be performed while rotating along the 13 rotational symmetry axes in FIG. 15 by using a three-axis reaction wheel provided inside the satellite. In other words, including the operation without rotation (identity transformation) E, the array antenna can be observed under a total of 14 operations 0={4C3, 3C4, 6C2, E}.

[0076] By utilizing the above-mentioned features, it becomes possible to perform omnidirectional electromagnetic wave observations using rotational operations as shown below. Figure 16 shows the observed global sphere when the spacecraft 3 rotates along its axis of rotation.

[0077] By rotating the spacecraft 3 along its axis of rotation, it becomes possible to analyze the electromagnetic wave source in the observational global area around the spacecraft 3. However, for simplicity, we assume that the electromagnetic wave source located at the direction vector s+σ from the spacecraft is stationary on the two-dimensional observation plane (u, v). In an electromagnetic interferometer, the position of the electromagnetic wave source is identified using a pattern of strong and weak interference fringes (hereinafter, fringes) that are generated when the electromagnetic waves arriving from the electromagnetic wave source to a pair of array antennas are interfered with and output. In other words, a phase difference occurs in the arriving electromagnetic waves due to the difference between the distance vector B(t) between the array antennas and the direction vector of the electromagnetic wave source. For example, if the electromagnetic wave source is between the fringe lines as shown in the figure, the interference output is weak, and if it is on the fringe line, the output is strong.

[0078] As mentioned above, since the spacecraft 3 can perform measurements while rotating, the baseline vector is a function of time. Therefore, the direction vector of the fringes on the observation plane is also a function of time, resulting in the generation of interference output strengths (hereinafter, fringe patterns) that vary according to rotation. In the case of a six-channel array antenna, a total of 15 fringe patterns can be obtained. Therefore, by observing and analyzing these 15 fringe patterns under a total of 14 rotation operations, it becomes possible to map the direction of the electromagnetic wave source within the full solid angle of 4π steradians.

[0079] Fig. 17 shows the relationship between the number of resolvable electromagnetic wave sources when the 6-channel array antenna of the present invention is adopted. In conclusion, the number of resolvable electromagnetic wave sources is 5 in observations using three satellites with 1-channel antennas, whereas it is 20 in observations using the 6-channel dipole antenna array of the present invention (without rotation), and it is several thousand in observations using the 6-channel dipole antenna array of the present invention (taking rotation into account).

[0080] As described above, according to this embodiment, the spacecraft can observe electromagnetic waves under a total of 14 operations 0={4C3, 3C4, 6C2, E}. First, according to this embodiment, the spacecraft has six dipoles orthogonal to the three axes X, Y, and Z, so the observed electromagnetic field data can be written as a six-dimensional vector as shown in the following equation (2).

[0081]

number

[0082] In formula (2), [] T represents the transpose of a vector, and E represents the electric field obtained by the dipole antenna. Therefore, the observation data obtained under operation O can be expressed by the following equation (3).

[0083]

number

[0084] In the interferometry, the covariance matrix CF(p, O) of the observation data can be calculated from equation (3) as shown in the following equation (4).

[0085]

number

[0086] where () H is the complex conjugate transpose. The covariance matrix CF(p, O) of the observation data is N 2 = 36 elements, but because it is a symmetric matrix, the information dimension obtained is N(N+1) / 2 = 21. The number of resolvable electromagnetic wave sources that can be determined based on second-order statistics is one less than the information dimension, so the maximum number without rotation is 20.

[0087] Furthermore, if an invariant phase reference signal is available in operation O and the observation time is not taken into account, the covariance between the observed signals under each operation can be calculated. That is, if the number of antennas is 6, the number of observed signals is N = 6 x 14 = 84, and if all observed signals are assumed to be independent, the number of dimensions of the obtained information is 3570. Therefore, the number of resolvable electromagnetic wave sources is a maximum of 3569 when rotation is taken into account.

[0088] The left bar graph in Fig. 17 shows the number of resolvable electromagnetic wave sources of a spacecraft that performs electromagnetic wave source exploration using a constellation of three satellites that is already in operation in outer space as an example of a conventional spacecraft. Since each satellite uses information from one antenna, the information dimension is 3x4 / 2, and the number of resolvable electromagnetic wave sources is 5. As shown in the right bar graph in Fig. 17, according to the present invention, electromagnetic waves including rotation information are received and analyzed, making it possible to analyze electromagnetic wave sources in all directions in detail.

[0089] A flowchart of omnidirectional electromagnetic wave observation using the spacecraft 3 according to this embodiment is shown in Fig. 18. An example of the system configuration for omnidirectional electromagnetic wave observation is shown in Fig. 19.

[0090] According to the system configuration example of Fig. 19, a spacecraft 3 moving on an orbit around the Earth receives electromagnetic waves from electromagnetic wave sources related to human activities on Earth and electromagnetic waves from electromagnetic wave sources related to human activities in outer space, and also communicates (downloads data and updates commands) with a ground station (base station) installed on Earth. Note that the spacecraft 3 is placed in various orbits such as low orbit, medium orbit, geostationary orbit, and cislunar orbit depending on the object of observation. Here, a placement assuming a low orbit is shown.

[0091] In this arrangement, omnidirectional electromagnetic wave observation using the spacecraft 3 is performed in the following procedure. First, in processing step S11, the ground station uplinks to the spacecraft 3 commands regarding vehicle rotation control, observation means, observation time, and signal processing method at the spacecraft site. Next, in processing step S12, the spacecraft 3 observes electromagnetic waves while controlling the rotation of the vehicle in accordance with the commands. The observation targets all electromagnetic waves around the spacecraft, such as electromagnetic waves arriving from electromagnetic radiation sources related to human activities on Earth and electromagnetic waves arriving from electromagnetic radiation sources related to human activities in outer space.

[0092] Next, in process step S12, the spacecraft performs signal processing of the electromagnetic waves in accordance with the command. This observation process is continued in process step S12 until the observation time ends.

[0093] Signal processing after the observation processing may involve, for example, the electromagnetic interference or covariance of electromagnetic waves described above in processing step S14. After the observation time ends, the signal processing results are downlinked from the spacecraft 3 to the base station in processing step S15. Finally, in processing step S16, the base station performs an analysis in consideration of the signal processing results, the orbital conditions of the spacecraft, and the vehicle rotation control conditions, to clarify the electromagnetic wave source characteristics and position. EXAMPLES

[0094] In the seventh embodiment, various functions realized by using a thin film will be described.

[0095] In this case, a thin film is provided on a plane formed by three or more tension members 2 connected via each end 7 of the deployable beam member 1, and one of the various functions realized by using this thin film is to provide a function of a photon energy receiver that receives photons. This allows the spacecraft to be operated using photon energy, making it possible to operate the spacecraft for a long period of time.

[0096] FIG. 20 shows a thin film provided on a plane formed by an expandable beam member 1 and a tension member 2, and the thin film portion constitutes a photon energy absorbing portion that absorbs photons from a star.

[0097] In this embodiment, eight openings in the plane of the tensile member 2 are used to provide a thin film, where photon energy absorbers are installed, that is, the absorbers absorb photons coming from the star and use the energy, for example, to generate electricity.

[0098] It is assumed that the spacecraft 3 will be operated while performing various rotation controls. According to the present invention, the tension members 2 on which the photon energy absorbers can be installed are symmetrically located in all directions of the spacecraft 3, so that there is a photon energy absorber facing the star at any rotation angle. This allows constant power generation regardless of the attitude of the spacecraft 3, and allows the spacecraft 3 to be configured for long-term operation.

[0099] Furthermore, a thin film is provided on the plane formed by three or more tension members 2 connected via each end 7 of the deployable beam member 1, and the second of the various functions realized by using this thin film is to provide a photon energy reflector that reflects photons. This allows the spacecraft to operate using the reflection of photon energy, making it possible to operate the spacecraft for a long period of time.

[0100] FIG. 21 shows a thin film provided on a plane formed by an expandable beam member 1 and a tension member 2, and the thin film portion constitutes a photon energy reflecting portion that reflects photons from a star.

[0101] In this embodiment, eight openings in a planar surface of the tensile member are utilized to provide a membrane into which photon energy reflectors are placed.

[0102] The photon energy reflector reflects photons coming from a star, and the spacecraft 3 can use the reaction force, for example, as a propulsion force. This spacecraft is assumed to be operated while performing various rotation controls, and as a result, the photon energy reflector faces the star at any rotation angle, and translational thrust can be constantly generated regardless of the attitude of the spacecraft, so that a spacecraft can be configured that can be propelled without using propellant and can operate for a long period of time.

[0103] Furthermore, as shown in Fig. 22, the photon energy reflecting portion can be configured to have a variable reflectance using, for example, a liquid crystal or an electrochromic device, so that torque control is also possible. That is, in Fig. 22, the photon energy reflecting portion can adjust the reflection and transmission of photons by varying the transmittance of photons using liquid crystal.

[0104] 22, photon energy is reflected at the upper part of the spacecraft 3 and transmitted at the lower part of the spacecraft 3, thereby generating a torque in the spacecraft 3 and controlling the attitude of the spacecraft 3. According to the present invention, tension members 2 on which photon energy reflectors can be installed are symmetrically located in all directions of the spacecraft, making it possible to control the attitude in any direction. [Explanation of symbols]

[0105] 10: Three-dimensional structure spacecraft 1: Expandable beam member 2: Tensile member 3: Spacecraft 5: Axis of rotational symmetry 7:End O:Origin

Claims

1. A spacecraft having a three-dimensional structure configured by deploying a plurality of deployable beam members and a plurality of tension members around the spacecraft, the spacecraft having a configuration in which the deployable beam members and the tension members are stored, The deployable beam members are deployed from a state in which they are provided in the spacecraft, and are arranged at substantially equal distances in a direction of a plurality of rotational symmetry axes of a virtual polyhedron formed with an approximate center of the spacecraft as an origin, a tension member supporting each of two ends of the nearest deployable beam member by tension, and each end of the deployable beam member being supported simultaneously by three or more tension members.

2. 2. The spacecraft of claim 1, The rotational symmetry axes are three orthogonal axes, a first deployable beam member and a second deployable beam member are disposed at approximately the same distance in the positive and negative directions of the axis of rotational symmetry, the axis being defined with an origin at approximately the center of the spacecraft.

3. 2. The spacecraft of claim 1, The rotational symmetry axes are three orthogonal axes, a first deployable beam member and a second deployable beam member are disposed at approximately the same distance in positive and negative directions with respect to the axis of rotational symmetry that is defined with the approximate center of the spacecraft as its origin, and the first deployable beam member and the second deployable beam member are disposed at approximately the same distance in positive and negative directions with their longitudinal directions perpendicular to the axis of rotational symmetry.

4. 2. The spacecraft of claim 1, a membrane on a plane formed by three or more of the tension members connected via each end of the deployable beam; wherein the deployable beams, the tension members, and the membrane comprise one or more antenna structures for emitting or receiving electromagnetic waves.

5. 5. The spacecraft of claim 4, The spacecraft receives electromagnetic waves while rotating along the multiple axes of rotational symmetry, thereby receiving and analyzing the electromagnetic waves including rotation information, and identifying the characteristics of the electromagnetic waves and the location of the source of the electromagnetic waves.

6. 6. The spacecraft of claim 5, The spacecraft downlinks the analysis results to a ground station, which then investigates changes in the analysis pattern according to the orbital conditions and rotation information of the spacecraft.

7. 2. The spacecraft of claim 1, a membrane on a plane formed by three or more of the tension members connected via each end of the deployable beam; The deployable beams, the tension members, and the thin film each include a photon energy receptor that receives photons.

8. 2. The spacecraft of claim 1, a membrane on a plane formed by three or more of the tension members connected via each end of the deployable beam; The deployable beams, the tension members, and the thin film each include a photon energy reflector that reflects photons.

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