JP2002362500A - Space structure and its development system as well as satellite solar power station - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a structure capable of being stored in a compact form and developed in space. SOLUTION: A hexagonal module M is constructed into a framed structure with beams 101 rigidly joined together. Twenty modules M are connected together in a predetermined pattern with joints J. The modules M of prismatic contour in a folded state are adapted to be developed in such procedures as to prevent the modules M from interfering with one another. In a developed process, the adjacent modules are connected and fixed to one another with couplers, and finally they become the space structure in a spherical shell form.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

The present invention relates to a space structure utilizing a structure of a carbon allotrope represented by C60 fullerene, a deployment system thereof, and a solar power generation satellite.
In particular, it is suitable for constructing large space structures such as solar power generation, space bases, spacecraft, orbiting fuel stations, and the like.

[0002]

2. Description of the Related Art At present, a plan is underway to construct a space station with the cooperation of various countries and to make crew members stay in space for use as solar power satellites and various experimental buildings, or as a base in outer space. ing. Also,
When used as a space station, a space structure with a certain degree of strength is required, but when launching from the ground, it is necessary to be able to store it compactly and obtain a large volume by deploying it in space. Have been.

For example, the above solar power satellite (SPS: Solar Po
wer Satellite) was first proposed by NASA in 1978, and various forms have been developed thereafter. FIG. 42 is a configuration diagram illustrating an example of a conventional solar power generation satellite. The solar power generation satellite 900 includes a plurality of power generation modules 901 and a main structure 902 having a tether structure.
An arm 903 connecting the main structure 902 and the power generation module 901, an attitude control device 904 provided at a joint between the arm 903 and the main structure 902, and a main structure 902.
And a power transmission antenna 905 provided at an end of the power transmission antenna. 43 is a configuration diagram showing the power generation module shown in FIG. 42, (a) is a front view, and (b) is a cross-sectional view.
The power generation module 901 includes a dish-shaped reflector 906 having a thin film structure, and a condenser 90 provided in front of the reflector 906.
7 is comprised.

The solar power satellite 900 is launched by a rocket into a geosynchronous satellite orbit at an altitude of 36000 km and assembled in outer space. The attitude control device 9
04 controls the attitude of the reflector 906 so that the reflector 906 faces the direction of the sun. The solar rays condensed by the reflector 906 are collected by a concentrator 90 having a solar cell panel.
7 and is transmitted from the power transmitting antenna 905 to the ground in the form of a high-frequency microwave (2.45 GHz). On the ground, the transmitted microwave is received by an antenna and converted to AC power by rectification. With such a solar power generation satellite, it is possible to generate several GW per unit.

[0005]

A conventional solar power generation satellite 9
00 is a reflection plate 9 for receiving sunlight from the front.
06 and the condenser 907 need to be attitude-controlled. Therefore, there is a problem that the configuration of the solar power generation satellite 900 is complicated. Further, from the viewpoint of the launch capacity of the rocket, including space structures other than the solar power generation satellite 900, those which are compact when deployed and which are deployed in space and function as high-strength structures are desired. I have. Thus, the present invention provides a new structure in outer space, and its application is not limited to the above-mentioned solar power generation satellite, and has a possibility of various developments.

[0006]

In order to achieve the above-mentioned object, a space structure according to claim 1 has a circular section in cross section, and a plurality of hexagonal hexagons which are components of a hollow shell structure. The module, a rotation connection means for rotatably connecting the sides of the module, and a connection fixing means capable of connecting and fixing the sides when the sides of the module are in contact with each other according to the expansion of the module, When the module is unfolded and deformed into a hollow shell structure from the state in which the module is folded by the rotation connection means, the module is combined with the rotation connection means and the connection fixing means so that the modules do not interfere with each other. It is.

[0007] This space structure is constructed as a model of a carbon allotrope having a spherical shell structure discovered in 1985. Such a carbon allotrope is a soccer ball-shaped C60 (buckminsterfullerene). Is known as a typical one. In addition to C60, C70, C120 (bucky dumbbell), C200
(Doughnut fullerenes) and the like are known, and all have a three-dimensional hollow structure. According to the present invention, it is possible to provide a space structure having a hollow shell structure having a circular cross section by adopting a structure of a carbon allotrope at a molecular level as the space structure.

The module may have a frame structure composed of a plurality of beams, or may have a plate shape. The rotary connecting means may have any structure as long as the modules are rotatably connected. Further, the fixed connection means only needs to have a structure capable of self-connection between the sides when the sides of the modules are in contact with each other, and various known or novel techniques having the function can be used.

Further, in order to prevent the modules from interfering with each other, it is possible to judge by simulating a plurality of specific connection configurations. If the modules are connected to each other by the combination determined by such a simulation, it is possible to prevent the module from being unable to be deployed in the outer space at the time of deployment.

In the space structure according to a second aspect of the present invention, in the space structure, the module is formed into a frame structure including six beams, and wires or other wires connecting the nodes of the frame structure. A reinforcing member is provided. By providing the reinforcing member in the frame structure, the bending rigidity in the module plane can be increased. In addition to the wire, a hollow or solid rod can be used as the reinforcing member.

The space structure according to a third aspect of the present invention is the space structure, wherein joints are provided at nodes of the frame so as to be foldable in a plane. In this way, by configuring the hexagonal module to be foldable in the plane, it is possible to store the module more compactly when launching with a rocket.

According to a fourth aspect of the present invention, in the space structure, a center joint is provided at the center of two pairs of opposing beams in the frame structure including six beams. An extension means extending and contracting by an actuator is provided between joint portions shared by a pair of beams.

When the telescopic member is extended by the actuator, the joints of the two pairs of beams which the telescopic members pass over are opened, and accordingly, the beams are opened at the central joint and the joints, and the module is deployed. According to this structure, the space structure can be further compactly folded.

According to a fifth aspect of the present invention, in the space structure, when the module is deployed, a force balance at the time of deployment under zero gravity is obtained by simulation, and the position of the space before and after the deployment is determined. The rotary connection means and the fixed connection means are combined and attached to a module based on a development pattern in which the displacement becomes small.

When the module is deployed under zero gravity, there is a possibility that the position of the space structure is largely changed by the inertial force at the time of module deployment. For this reason, the balance of the force at the time of module deployment is simulated, and the deployment pattern is determined based on this result. And
The rotation connection means and the fixed connection means are combined and attached to each module so as to be deployed according to this deployment pattern. In this way, the movement of the module during deployment can be suppressed to a very small value.

Further, as an example of the above-mentioned space structure, the following structure can be mentioned. The space structure according to claim 6 has 20 hexagonal modules capable of forming a spherical shell shape, and five of the modules (M1 to M5) are connected in a C-shape with each side thereof. The next module (M6) is connected to the outermost side of the C-shaped module (M5) located at the rear end of the C-shape and includes the module (M6).
Modules (M6 to M15) are connected in a zigzag with one side in a substantially straight line, and the next five modules (M16 to M20) are connected in an inverted C-shape with each side, and Connected modules (M6-M
15) Of the modules (M15) located at the end,
The inverted C of the module (M16) located at the tip of the inverted C character
A plurality of rotary connecting means for connecting to the outermost outer side of the character, and a plurality of connecting and fixing means for connecting the modules that contact each other at each side when each module is developed with the rotary connecting means. It is.

The order of deployment of the space structure can be determined by a deployment simulation so that the modules do not interfere with each other. Specifically, from the state where the first module (M1) to the last module (M20) are sequentially stacked, the rotational connection means between the modules is represented by J1, J2... J18, J19. J1, J2, J3, J5, J4, J7, J6 one by one, then J8 and J9, J10 and J1
1, J12 and J13, J14 and J15, J16
And J17 in the order, and J18, J19
The modules are developed one by one in the order of (15). Note that the deployment method includes the purpose of minimizing the movement of the space structure before and after deployment, so that the deployment order does not need to be as long as the module does not interfere.

The space system deployment system according to claim 7, wherein the space structure, temporary holding means for holding the module in a folded state and releasing the holding based on a release command, A deploying means for deploying the module around the rotational connection means based on the command; and a control means installed on the space structure and issuing a release command and a deploy command to the holding means and the deploying means. .

When a release command is issued by the control means, the temporary holding means releases the holding of the module. Examples of such temporary holding means include explosive bolts and fuses, or those that drive pins using a motor or solenoid. Next, when a deployment command is issued by the control means, the module is deployed around the rotation connection means. As the deploying means, various known techniques that enable a rotation operation of a motor, a solenoid, or the like can be used. When the modules are deployed, the modules are fixed by the fixing connection means, and a space structure having a hollow shell structure is formed.

The space structure deployment system according to claim 8 is the space structure deployment system, further comprising: communication means for performing communication between the space structure and the ground; Command means for instructing the control means to issue a release command or a deployment command via communication means. If there is no person working in outer space, it is necessary to launch the folded space structure with a rocket and deploy it in outer space. Therefore, a command base is installed on the ground, and a release command or a deployment command is issued to the control means remotely.

A space structure according to a ninth aspect of the present invention is the space structure, wherein a pressure block of an inflatable film structure having an available internal space is provided inside the space structure. By arranging the space structure around the pressurized block having the inflatable film structure and, for example, providing a protection wall on the module, it is possible to protect the pressurized block from collisions such as space debris. Further, by providing the radiation shielding plate, it becomes possible to suppress the exposure in the pressurized block. As described above, the space structure can be used for various applications as a protective function of the pressurized block.

According to a tenth aspect of the present invention, there is provided a solar power module, wherein a solar cell panel is installed on the module of the space structure. By installing a solar cell panel on this space structure, a solar power generation module having a circular portion at least in cross section is formed. For example, when the space structure has a spherical shell structure, the solar power generation module can be installed in a substantially spherical shape, so that the hemispherical surface can receive sunlight without controlling the attitude.

Further, in the solar power generation module according to the present invention, a plurality of reflecting mirrors are provided on the module of the space structure, and a servomotor for controlling the angles of these reflecting mirrors is provided. A solar cell panel is installed.

The reflecting mirror of the module is controlled by a servomotor so that sunlight can be focused on a solar cell panel installed inside. The collected sunlight is photoelectrically converted in an internal solar cell panel. The shape of the reflecting mirror may be any of a rectangle, a rectangle, and a circle. The number and arrangement of the servomotors may be provided according to the number and shape of the reflecting mirrors.

According to a twelfth aspect of the present invention, in the solar power module, the solar cell panel is installed on a small space structure module installed inside the space structure. The sunlight condensed by the reflector is photoelectrically converted by a solar cell panel installed in a space structure module.

The solar power satellite according to claim 13 is:
The solar power generation module includes a plurality of solar power generation modules, a main structure supporting the solar power generation satellite, and power transmission means for converting power generated by the solar power generation satellite into high-frequency microwaves and transmitting power from an antenna. The electric power generated by the solar power generation module is transmitted from the antenna to the ground as high-frequency microwaves. On the ground, this microwave is received by the receiving antenna and converted into electric power again.

A spacecraft according to a fourteenth aspect is the spacecraft mother ship, and the spacecraft is mounted on the mother ship.
13 and an electric propulsion device that is attached to the mother ship and that obtains propulsion by electric power generated by the solar power generation module.

Although solar energy can be obtained indefinitely by using solar energy as the propulsion energy of the spacecraft, if the solar power generator installed in the spacecraft fails, the spacecraft is repaired or abandoned. There must be. Therefore, by using a solar power generation module that does not require attitude control as an energy source, it is possible to reduce the occurrence of a spacecraft failure.

[0029]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS The present invention will be described below in detail with reference to the drawings. The present invention is not limited by the embodiment.

[Structure of Space Structure] FIG. 1 is a perspective view showing a space structure according to the first embodiment of the present invention.
FIG. 2 is a developed plan view showing the space structure shown in FIG. This space structure 100 is a spherical shell-like structure composed of a plurality of modules M having a hexagonal frame structure, and each module M is connected by a joint J that can be deployed in a butterfly shape. In FIG. 2, each module is indicated by reference numeral “M” and each joint is indicated by reference numeral “J”, and the reference numerals are added to the module and joint numbers. It should be noted that a joint number having a rotation angle of 138 ° is further denoted by reference numeral “a”, and a joint number having a rotation angle of 222 ° is further denoted by reference numeral “b”.

Each module M is provided with a coupler C for connecting and fixing with other modules M.
In the figure, each coupler is indicated by a symbol “C”, and the symbol of the coupler C is added to the symbol. Further, the male coupler C is further denoted by reference numeral “a”, and the female coupler C is denoted by reference numeral “b”. Note that the module M may be configured by a plate-shaped body instead of the frame structure as shown in FIG. 1 (not shown).

FIG. 3 is an explanatory diagram showing an example of the module shown in FIG. Each module M has six beams 101
To form a hexagonal frame structure. The connecting portion 102 of each beam 101 may be a hinged connection, but is preferably a rigid connection to transmit a moment applied to the beam 101. Further, a solid rod or a rod having an open cross section can be used for each beam 101, but it is preferable to adopt a cylindrical or rectangular closed cross section in order to secure torsional rigidity. Further, the material of the module M is aluminum alloy or CFRP (carbon fiber reinforced plastic).
s) It is preferred to use a laminate. Since aluminum alloy and CFRP have low thermal conductivity, heat insulation measures can be omitted or a simple structure can be adopted.

In particular, when using an aluminum alloy,
There is an advantage that the workability is good and the weight can be reduced. Also,
The use of CFRP has an advantage that the load due to thermal deformation can be suppressed because the coefficient of linear expansion is small. CF
Regarding the torsional rigidity of RP, the ratio of ± 45 degree layer is 50%
Good results can be obtained by setting the degree to the above range.
Also, the coefficient of linear expansion of CFRP can be adjusted by the laminated structure. Further, when an elastic coefficient equivalent to that of the aluminum alloy is given, the weight of the module M can be reduced as compared with the aluminum alloy.

A steel wire 103 as a reinforcing member is stretched between the connecting portions 102 of the module M in a star shape. By using the wire 103, the weight of the module M can be reduced, and a structure in which the module M is folded and deployed can be adopted. When the wires 103 are stretched, it is necessary to control the tension of each wire 103. By reinforcing the module M with the wire 103, the rigidity of the module M is improved. Further, a rod may be used as the reinforcing member in addition to the wire 103 (not shown). When a rod is used, there is an advantage that a tensile load and a compressive load can be shared. The material of the rod is aluminum alloy or CF similarly to the beam 101.
Use RP. Preferably, the rod and the beam 101 are rigidly connected from the viewpoint of improving rigidity.

The reinforcing member has the effect of increasing the in-plane rigidity of the module M and improving the natural frequency and the amount of deformation. The rigidity of the reinforcing member is 1/10 to 1/1 / the beam rigidity.
1 is sufficient. FIG. 4 is an explanatory view showing a modification of the mounting form of the reinforcing member. The wire 103 or the rod as the reinforcing member may be provided between the coupling portions 102 facing each other so as to cross at the center of the module M as shown in FIG. Further, as shown in FIG. 3B, the holes may be provided in an hourglass shape. Further, as shown in FIG. 3C, a wire 103 may be passed between a pair of connecting portions 102 facing each other, and the hourglass-shaped wires 103 may be provided on both sides around the wire 103. . The arrangement structure of these wires 103 is
May be determined based on the natural frequency and the amount of deformation required for the above.

FIG. 5 is an explanatory view showing the joint structure shown in FIG. As shown in FIG. 1A, the joint J is configured such that the fixed-side module Mf and the deployment-side module Ma can be relatively rotated about a rotation axis. The rotation of the module Ma is performed by the actuator 105 provided on the rotation shaft. As the actuator 105, a motor, a spring, a shape memory alloy, an air pressure, or the like can be used. Elements that convert the rotational motion of the motor into the unfolding motion include a speed reducer, a combination of a ball screw and a cam, and the like. further,
A cam can be used as a motion conversion element when using a spring and a shape memory alloy. When a spring is used as the actuator 105, the energy for deployment can be extremely small, and the structure of the actuator 105 including the electric system can be simplified.

As a specific example of such an actuator 105, as shown in FIG. 3B, the output shaft of a speed reducer 106 having a high reduction ratio is connected to the rotation shaft 107 of the joint J, and the speed reducer 106 And an input shaft connected to the rotation shaft of the motor 108. The motor 108 and the speed reducer 106 are provided at an end of the joint J. In the figure, the module Ma that rotates 138 degrees is indicated by a solid line, and the module Ma that rotates 222 degrees.
Is indicated by a dotted line. Further, the virtual center of rotation of the module M is located at a position eccentric from the center of rotation of the joint J.

If the motor 108 and the speed reducer 106 are provided at the center of the joint J, they do not hinder the unfolding operation of the module M or the attachment of an element such as a solar cell panel to the module M (not shown). ).
Further, if the deployment operation of the module M fails due to the failure of the actuator 105 or other reasons, it becomes impossible to deploy the module M after that module. Therefore, it is preferable to arrange a plurality of the actuators 105 in parallel.

FIG. 6 is a perspective view showing the structure of the coupler shown in FIG. The coupler C shown in FIG. 1A includes a pin 111 having a groove 110 for latching, a conical cylinder 112 having a guiding function, and a fixed latch 113. The pin 111 is provided on the module M as the male side, and the cylinder 112 and the fixed latch 113 are provided on the module M on the female side. By unfolding the module M, the pins 111 of the predetermined module M are inserted into the cylinder 112 of the module M to be connected, and the fixed latch 113 is engaged with the groove 110 of the pin 111 to connect the modules M together. Fixed. Further, by providing an actuator 114 to the fixed latch 113,
11 may be opened. Actuator 11
For 4, a motor, an air cylinder, or the like can be used.

The coupler shown in FIG. 4B has a T-shaped handle 115 and a groove 1 into which the handle 115 is fitted.
16 with two plates 117 and each plate 11
7 is provided with a fixed latch 118 provided in the control unit 7. At the center of the handle 115, a triangular plate 119 is provided. Module M with handle 115 as male side
And the module 117 is provided on the module M with the plate 117 and the fixed latch 118 as the female side. When the module M is deployed, the triangular plate 119 of the handle 115 enters between the two plates 117 provided on the module M to be connected, thereby performing a guiding function. When the handle 115 is fitted in the groove 116, the handle 115 is locked by the fixing latch 118, and the modules M are connected and fixed. As described above, by providing an actuator,
The fixing latch can be opened (not shown).

The coupler shown in FIG. 3C is composed of a T-shaped handle 120 and a grip mechanism 121 for gripping the handle 120. The grip mechanism 121
It is driven by a motor or other actuator. The handle 120 is provided on the module M as a male side,
The grip mechanism 121 is provided on the scalpel side. When the module M is expanded, the handle mechanism 120 is opened and the grip mechanism 12 is opened.
1, the grip mechanism 121 grips the handle 120. Thereby, the modules M are connected and fixed. The grip mechanism 121 performs a gripping operation by detecting the handle 120 with a sensor or the like and operating an actuator based on the sensor signal. In addition,
A pusher may be provided in the grip, and the grip may mechanically perform a gripping operation when the handle presses the pusher with the unfolding operation. Although not shown, a simple guide plate having a trapezoidal shape or the like may be provided on the module M, and the module M may be connected by a magnet or a fixed latch in a state where the guide plates are assembled (all are not shown).

FIG. 7 is an explanatory view showing a state where the module is folded. The space structure 100 has a predetermined angle (138 °, 22 °) at the joint J shown in the developed view of FIG.
By rotating it by 2 °), it is stacked and folded on a hexagonal prism (see FIG. 7A). When launching the space structure 100 into the outer space, the module M is accommodated in the rocket in a folded state as described above. Each module M has
A temporary fixing device 125 for maintaining a stacked state is provided. The temporary fixing device 125 holds the stacked state of the modules M, prevents the modules M from vibrating at the time of launch, and damages the actuator, and also functions as a safety device in managing the module deployment order.

In this temporary fixing device 125, as shown in FIG. 7B, the module M and the module M are temporarily fixed by the fixing portion 126 and the explosion bolt 127, and when the module is deployed, the explosion bolt 127 is exploded. To release the fixation. In particular, when the explosion bolt 127 is used, there is an advantage that operation reliability and holding rigidity are high. Also,
In addition to the explosion bolt 127, the module can be temporarily fixed by a fuse or a combination of a fuse and a spring (not shown).

In addition to the forced separation type such as the explosion bolt 127, a notch pin type temporary fixing device can be used. FIG. 8 is an explanatory view showing a notch pin type temporary fixing device. In this temporary fixing device 128, a latch 129 is provided on one module M, and a notch 130 or a pin for hooking the latch 129 is provided on the other module M. Then, the latch 129 is rotated or slid by the actuator 131, and the restraint is released by the latch 129. The actuator 131 includes:
A motor, solenoid, paraffin, or the like can be used.

FIG. 9 is an explanatory diagram showing a modification of the module structure. The foldable module M is, as shown in FIG.
In this configuration, the telescopic device 141 is passed between the two. The telescopic device 141 has a structure in which the telescopic wire 143 can be wound and pulled out by the central actuator 142. As the actuator 142, an actuator that performs a rotary motion can be used, and examples thereof include a combination of a motor and a speed reducer, and a combination of a cylinder and a motion conversion device that converts a linear motion into a rotary motion. Further, the connecting portion 102 between the beams 101 has a rotatable joint structure such as a hinge. Further, the beam 101 on both sides of the connecting portion 102 across which the telescopic device 141 is provided has a coil spring 144 at the center thereof. It is provided, and has a structure that can be bent freely at the relevant portion. Wires 103 are stretched crosswise on both sides of the extension device 141.

FIG. 7B shows a state where the module is folded. When the actuator 142 is operated to pull out the telescopic wire 143 from the state in which the module M is folded, the beam 101 is opened by the spring force of the coil spring 144 according to the amount of pulling out of the telescopic wire 143. With module M fully deployed, wire 10
3 is stretched in a cross shape with a constant tension, and functions as a reinforcing member. When folding from the expanded state, the actuator 142 may be driven to wind the telescopic wire 143. By making the module M a folded structure, the space structure 100 can be further compacted. Note that the configuration of the extension device 141 is the same as that of the actuator 1.
42 and the extension wire 143 are not limited. For example, a nut is provided in the connecting portion 102 of the module M, and the connecting portion 10
A configuration may be adopted in which a screw screwed to a nut is passed between the two and the screw is rotated by an actuator (not shown).
In this case, when the screw is rotated by the actuator, the nut moves on the screw, and accordingly, the module M can be folded and unfolded.

FIG. 10 is an explanatory view showing a storage configuration of the space structure. The space structure 100 is folded in the state shown in FIG. 7A and stored in the payload envelope 150 in the rocket fairing. For example, the payload envelope 150 of the H-2 rocket has a diameter d = 4 m and a length L = 12 m (a cylindrical portion is about 7 m). The U.S. Space Shuttle envelope has a cylindrical shape with a storage height of about 15 m and a diameter of about 4 m. The space structure 100 can be folded to obtain a diagonal dimension (D = 4.6 m) of the module M.
Rocket or space shuttle payload envelope 1
50 can also be easily stored.

[Space System Deployment System] FIG.
FIG. 1 is a configuration diagram illustrating a deployment system of a space structure. The expansion system 200 includes a storage device 201 that stores a predetermined expansion program described later, a processing device 204 that loads the expansion program into a memory 202 (Random Access Memory) and executes it by a CPU 203, and performs a module expansion operation. A communication system 206 and a communication antenna 207 for performing communication with the ground by microwaves, an electric circuit 208 for driving the joint J and the coupler C of the module M and the temporary fixing device 125, and a ground command. And a base 209.

The ground command base 209 is provided with the space structure 100
A remote control system 210, a large communication antenna 211 for transmitting or receiving microwaves, and a communication system 212 connected to the remote control system 210. A command signal for driving the joint J or the like is placed on a microwave carrier wave to perform frequency modulation. Send by doing. The execution instruction of the deployment system 200 is:
This is performed using the remote control system 210 of the ground command base 209. The remote control system 210 includes a computer 214 having a remote control program 213, an image display device 215, and an input device 216. As shown in the figure, components other than the ground command base 209 are mounted on the space structure 100 side. However, a storage device 201 storing the deployment program and a processing device 204 executing the program are provided on the remote side of the ground. It is also possible to shift to the operation system 210 and send a drive command directly to the electric circuit 208 installed on the space structure 100 side.

The storage device 201 includes a nonvolatile memory such as a hard disk device, a magneto-optical disk device, and a flash memory, a volatile memory such as a RAM,
Alternatively, it is constituted by a combination of these. Also,
As shown in the figure, the ground command base 209, the communication system 212, and the communication antenna 211 can be installed at a plurality of points on the earth, and these can be connected via the Internet 217 to remotely control the space structure 100. Further, the power supply 205 may be supplied with power from a solar cell panel 218 attached to the module M of the space structure 100.

FIG. 12 is a functional block diagram of the deployment system shown in FIG. The unfolding system 200 is realized by a combination of the unfolding program and various hardware, and includes a temporary fixing releasing unit 231 that releases the temporary fixing of the folded module M, and a joint J that rotatably connects the modules M to each other. Actuator 105
, A coupler operating unit 234 for connecting the coupler C, and a control unit 235 for controlling these. The control unit 23
5 and the temporary release unit 231, the joint operating unit 232, and the coupler operating unit 234, when the control unit 235 is provided on the space structure 100 side, by the transmission line.
When 35 is provided on the ground command base 209 side, communication is performed by microwaves.

FIG. 13 is a flowchart showing the operation of this expansion system. First, at the ground command base 209, a command to deploy the space structure 100 is input from the input device 216 of the remote control system 210. Alternatively, a deployment command is automatically issued for a predetermined time or by a predetermined trigger signal (step S101). Next, this expansion command is pulse-modulated by the communication system 212 and transmitted from the communication antenna 211 to outer space through microwaves. The microwave signal is received by the communication antenna 207 on the space structure 100 side, demodulated into a deployment command by the conversion circuit of the communication system 206, and sent to the processing device 204 (step S102).

The processing device 204 reads out the development program of the space structure 100 from the storage device 201 and loads it into the memory 202 (step S103). The expansion program includes programs for releasing the temporary fixation, the expansion order and expansion speed of the module J described later, the operation timing of the coupler C, and the like. The control unit 235 controls the temporary fixing releasing unit 231, the joint operating unit 232, and the coupler operating unit 234 according to the deployment program.

Subsequently, the temporary fixing releasing section 231 releases the temporary fixing of the module M by the temporary fixing device 125 according to the expansion program (step S104). For example, FIG.
As shown in (1), when the explosion bolt 127 is used as a temporary fixing device, the electric circuit 208 is controlled to supply a current to the explosion bolt 127 to cause an ignition explosion. FIG.
As shown in (2), when a notch pin type is used as the temporary fixing device 125, the actuator 131 is operated by controlling the electric circuit 208, and the latch 129 is operated.
Release the temporary fixation by.

Next, the joint operating section 232 operates the actuator 105 of the joint J, and deploys the module M whose temporary fix has been released (step S10).
5). Here, when the module M is expanded, if it is not necessary to connect and fix the module M which has already been expanded (step S106), the temporary fixing of the next module M is further released, and the released module M is expanded. (Steps S104 to S105). On the other hand, when it is necessary to connect and fix the module M already deployed, the coupler operating unit 23
In step S107, the corresponding coupler C is operated to connect and fix the modules (step S107).

The operation of the coupler C is necessary when the coupler C has a gripping mechanism 121 as shown in FIG. 6C, and the fixed latch 113 shown in FIGS. When using 118, coupler C
Does not need to be operated by a deployment program. Then, the above operation is repeated until all the modules M are expanded (step S108). In addition, the process of the deployment operation is described in the communication system 2.
It is also possible to monitor on the ground by using 06, 212.
Further, the unfolding operation may be performed after all the temporary fixing is released.

[Order of Deploying Module of Space Structure] Next, the order of deploying the module M of the space structure 100 will be described. When the modules are deployed in the order of the joint numbers in the planar development view shown in FIG. 2, the modules M may interfere with each other and cannot be deployed. So, module M
Is tentatively determined, and a simulation of whether or not interference occurs by expanding the module M in accordance with the expanded structure (a simple expanded simulation is sufficient, and the moment of inertia, attractive force, and other conditions of the module are unnecessary). ) Was done. As a result, it was found that interference can be avoided by expanding the modules M in the order shown in the following table.

[0058]

[Table 1]

In the same table, the left column shows the deployment order, and the right column shows the joint numbers developed in the order.
When the joint numbers are described in parallel in the right column, it means that the joints J are simultaneously developed. Since the order of deployment prevents interference between modules, the modules M can be reliably deployed in space. The expansion order can be obtained by simulation other than the expansion order, but this expansion order is excellent in that the movement of the space structure 100 during expansion can be minimized.

FIGS. 14 to 28 are explanatory diagrams showing the process of deploying the space structure. In these figures, the modules of the space structure are represented in plane for easy explanation. The space structure 100 is folded and held in a hexagonal prism shape as shown in FIG. The space structure 100 is released from the envelope, and starts deployment in response to a deployment command from the ground command base 209. First, in step 1, as shown in FIG. 15, the joint J1a is operated to deploy the module M1 by 138 ° (hereinafter, the module with “a” added to the joint number is deployed by 138 °). . Here, in the zero-gravity state, the other modules rotate to a considerable extent due to the reaction of expanding the module M1.

In step 2, as shown in FIG. 16, the module M2 is connected to the module M2 by operating the joint J2b.
The joint is developed by 2 ° (hereinafter, the joint number with “b” is developed by 222 °). Subsequently, in step 3, as shown in FIG.
Is activated to deploy the module M3. Also,
In step 4, as shown in FIG.
The module M5 is changed to the module M4
Expand with. Here, the reason why the joint J4b is not operated is to prevent interference between the modules as described above.

At step 5, as shown in FIG. 19, the module M4 is deployed by operating the joint J4b. Here, the expansion of the module M4 causes the modules M5 and M1 to connect to the couplers C1a and C1a.
It is connected and fixed by b. Next, in step 6, as shown in FIG.
Is operated to deploy the module M7 together with the module M6. In step 7, as shown in FIG. 21, the joint J6b is operated to deploy the module M6.

In step 8, as shown in FIG. 22, the joints J8b and J9a are simultaneously operated, and the modules M8 and M9 are simultaneously deployed. Here, as the module M8 is expanded, the module M8 and the module M1 are connected and fixed by the couplers C2a and C2b. Subsequently, in step 9, FIG.
As shown in FIG.
11a are operated simultaneously, and the modules M10 and M11 are simultaneously deployed. By the expansion of the module M10, the module M10 and the module M2 are connected and fixed by the couplers C3a and C3b.

In step 10, as shown in FIG.
The joints J12b and J13a are simultaneously operated, and the modules M12 and M13 are simultaneously deployed. By the expansion of the module M12, the modules M12 and M3 are connected to the couplers C4a and C4a.
It is connected and fixed by 4b. In Step 11, FIG.
As shown in FIG.
15a are operated simultaneously, and the modules M14 and M15 are simultaneously deployed. By the expansion of the module M14, the module M14 and the module M4 are connected and fixed by the couplers C5a and C5b. Further, by expanding the module M15, the module M15 and the module M6 are connected and fixed by the couplers C6a and C6b. Note that detailed module numbers are omitted from this figure.

In step 12, as shown in FIG.
The joints J16b and J17a are simultaneously operated, and the modules M16 and M17 are simultaneously deployed. By the expansion of the module M17, the modules M17 and M7 are connected to the couplers C7a and C7a.
7b. In step 13, FIG.
By operating the joint J19b, the module M19 is deployed together with the module M18 as shown in FIG.
Further, by expanding the module M19, the module M1
9 and module M11 are connected and fixed by couplers C9a and C9b.

Next, in step 14, as shown in FIG. 28, the module M18 is deployed by operating the joint J18. By the expansion of the module M18, the module M18 and the module M9 are connected and fixed by the couplers C8a and C8b. Further, the module M16 and the module M20 are provided with a coupler C11a,
The module M20 and the module M13 are provided by C11b.
Are connected and fixed by the couplers C10a and C10b. Through the above expanding operation, the substantially spherical space structure 100 as shown in FIG. 28 is assembled.

According to the above-described development order, when the inventor actually simulated, it was found that the movement before and after the development was minimized. Also, very low levels of acceleration and angular acceleration during deployment were also observed. Further, it has been found that the angular acceleration gradually decreases because the difference between the moment of the deployment module M and the moment of the other module M appears as the deployment step proceeds.

In this simulation,
In order to confirm the basic operation of deployment and the effect of gravity gradient torque, commercially available motion analysis software "AMAM
Using "S", a deployment simulation in a zero-gravity state and a simulation in consideration of gravity gradient torque were performed.
First, as an analysis condition, an initial velocity was given so that the structure in the housed state orbits in the earth direction. The joint deployment operation used the STEP5 function implemented in the software to smooth the operation. Other detailed conditions are shown in Table 2 below.

[0069]

[Table 2]

The stacking direction of the modules was set to the Y axis, and the module surface was set to the XZ axis. The initial speed is V
_{o i = (G × M /} r i) 1/2 (i = 1,2,3, ... 20)
Given in. As a result of this simulation, in the zero gravity state, the angular velocity converged and showed a value close to zero, but when the gravity gradient torque was considered, the angular velocity occurred in the X-axis and Z-axis directions. However, it was found that the influence of the gravitational gradient torque was almost negligible because the rotation speed was very low.

[Embodiment of Space Structure] The space structure 100, which is a basic embodiment, includes a module M having a Young's modulus E
= 7300 kg / mm ² and an aluminum alloy having a Poisson's ratio ν = 0.33 was used. In addition, a tube-shaped beam having a closed cross section is used, and its length (one side of a hexagon) is 2000 m.
m (diameter of the spherical shell after development is 9.91 m), outer diameter D = 80 m
m and the plate thickness t = 2 mm. Further, a steel wire was used as the reinforcing member 103, and was stretched in a star shape. The beams 101 are rigidly connected to each other to improve bending rigidity.

[Application Example of Space Structure] FIG. 29 is an explanatory diagram showing an application example of the space structure. Space structure 10
The solar cell panel 301 is attached to each module M of No. 0 to form a power generation module 300. By providing the solar cell panel 301 in each module M, the solar rays surely hit half of the sphere. Further, since the power generation module 300 has a substantially spherical shape, the power generation module 3
There is no need to control the attitude of 00. Therefore, the probability of occurrence of a failure during power generation is extremely low. In addition, it is not necessary to provide the solar cell panel 301 in every module M,
Any number can be used as needed. Furthermore, if the surface of the solar cell panel 301 is formed into a spherical shape, the entire shape approaches a spherical shape, so that sunlight can be efficiently received (not shown). In addition, the wiring derived from the solar cell panel 301 is collected at a predetermined location through the inside of the tubular beam 101 (not shown).

FIG. 30 is a configuration diagram showing a solar power generation satellite. The solar power satellite 350 is composed of a structural material such as an aluminum alloy or CFRP, and has a main structure 352 having a plurality of arms 351 and a main structure 35.
Power generation module 300 provided at the tip of second arm 351
And a power transmission device 353 and a power transmission antenna 354 provided at one end of the main structure 352. Wiring for power collection is provided in the main structure 352, and electrically connects the power generation module 300 and the power transmission device 353. Further, the main structure 352 may have a tether chain structure.

In the power transmission device 353, the power generation module 30
The DC current transmitted from 0 is converted to a high-frequency microwave of 2.45 GHz. The power transmitting antenna 354 has a configuration in which a plurality of power transmitting antenna elements are arranged in an array, and each power transmitting antenna element is provided with a power amplifier. The microwave generated by the power transmitting device 353 is
And transmitted from the power transmitting antenna 354 to a receiving antenna (not shown). As a method of accurately transmitting a microwave to a receiving antenna, Japanese Patent Application Laid-Open No. 6-32
The technology disclosed in Japanese Patent No. 7172 is known.

In addition, the storage device 201 and the processing device 204 for executing the development program of the space structure 100 described above include the control device 3 installed in the power transmission device 353.
55. Further, the control device 355 houses a communication system 206 and an electric circuit 208 that communicate with the ground and transmit microwaves to the ground. The power supply 205 is powered by power obtained from the solar cell panel 301.

According to the solar power generation satellite 350, uniform power generation can be performed irrespective of the direction of the sun's rays, so that there is no need for attitude control as compared with a general panel-shaped solar power generation satellite. For this reason, the actuator for attitude control can be omitted or simplified, and the occurrence of a failure in outer space where maintenance is difficult can be greatly suppressed. Further, the arrangement of the power generation modules 300 can be freely set.

For example, as shown in FIG. 31, the power generation modules 300 of the solar power generation satellite 360 can be arranged in a square plane. The power generation module 300 is used as a connection point of the truss structure constituted by the beam 361 having the nanotube structure. At this time, the connection is a pin connection, but a rigid connection ramen structure may be used to transmit a bending moment. With such a configuration, the solar power generation satellite 360 can be used as a relatively large structure.

Further, when the solar power generation satellite 360 is in a geostationary orbit in a specific attitude, a reflecting plate 362 may be provided on the back side of this square plane (indicated by a dotted line in the figure). By doing so, the power generation module 300
Can be irradiated with sunlight. In addition, even if the reflecting plate 362 is slightly inclined with respect to the sun rays, it is possible to secure sufficient reflected light by using the slightly larger reflecting plate 362.

Further, as shown in FIG. 32, an arm 372 extends radially from a main structure 371 having a tether chain structure, and a power generation module 300 is attached to the tip of the arm 372.
May be provided. At this time, it is preferable to change the extending direction of the arm 372 between the upper and lower stages. This is because when the vehicle is in a geostationary satellite orbit in a specific attitude, it is difficult for other power generation modules 300 to block sunlight. In addition, a power transmission device 373 and a power transmission antenna 374 are provided at one end of the main structure 371. The power transmission method is the same as that of the solar power satellite shown in FIG.

FIG. 33 shows the configuration of another solar power generation satellite. The solar power satellite includes a plurality of power generation modules 30.
0 is coupled to a beam 381 having a nanotube structure by a tether 382. In the above-mentioned solar power generation satellite, the number of power generation modules 300 is not limited to the number shown in the drawings, and can be appropriately determined according to required power.

FIG. 34 is a configuration diagram showing a modification of the power generation module. FIG. 35 is an explanatory diagram illustrating a light-collecting device of the power generation module illustrated in FIG. 34. This power generation module 400 has a configuration in which a space structure 100S is further installed inside the space structure 100L (FIG. 34 is partially cut away). In the module M of the outer space structure 100L, a light collecting device 401 that collects sunlight is provided. As shown in FIG. 35A, the condensing device 401 has a configuration in which a plurality of rectangular mirrors 402 are provided on the inner surface of the module M, and its rotation is controlled by a control servomotor 403 installed on the beam 101. I have. On the other hand, the solar cell panel 404 is installed in the module M of the inner space structure 100S.

The power generation module 400 is installed on an arm 405 of the main structure. The wiring from the solar cell panel 404 passes through the arm 405 and the main structure, and is connected to the power transmission antenna (not shown).
The method of deploying the space structure is as described above. However, in the power generation module 400, while the outer space structure 100L is being deployed, the inner space structure 100
S may be inserted therein, or both space structures 100 may be deployed in the same manner, and in the process, the outer space structure 100L may be accommodated in the outer space structure 100L. .

First, as shown in FIG. 35 (b), the sunlight K is reflected by the rectangular mirror 402 and condensed on the solar cell panel 404. Solar panel 404
Then, the condensed sunlight K is photoelectrically converted. According to such a configuration, even without using a large solar cell panel,
The current collection efficiency can be improved. The mirror control of the condensing device 401 is performed by a control device 406 provided at the tip of the arm. The power generation module 40
At 0, the small space structure 100S having the solar cell panel 404 is arranged in the large space structure 100L, but the space structure 100S may be omitted and only the solar cell panel may be arranged ( Not shown).

As shown in FIG. 36, the mirror 4
02 is a rectangular mirror frame 407 and this mirror frame 407
A plurality of mirrors 408 installed in the
8 may be constituted by a servomotor 409 for controlling the rotation. The servo motor 409 is connected to the control device 40
6 is controlled so that the sunlight can be focused on the solar panel 404. In this manner, the solar cell panel 404 can be appropriately irradiated with the solar light in a portion of the power generation module 400 where the solar light obliquely strikes, so that the light collection efficiency can be further improved. Note that the light condensing device 401 shown in FIGS.
Is preferably about 8 to 10 mirrors, but other numbers may be used.

FIG. 37 is an explanatory diagram showing a space base using a space structure. The space base 500 uses the space structure 100 on the lunar surface B. The space structure 100 includes a pressurizing block 501 having an inflatable film structure. This pressurized block 501
After being expanded inside the space structure 100, the structure can be provided inside the space structure 100 or combined with a module in an expanded state (not shown) to be used as the space base 500. Further, a metal protective plate 5 is provided on the surface of the module M.
By providing 02, it is possible to prevent the pressurized block from being damaged due to the fall of space debris or meteorite. Also,
By providing a shielding plate for shielding heat and radiation, the influence of solar heat and radiation in outer space can be reduced.

Further, by providing a solar cell panel as described above in a specific module M, it is possible to procure electric power in the base. Thus, the space structure 100
The space base 500 which can be easily installed and has high strength can be provided. In addition, when installing on the lunar surface B, it is preferable to fix it to the ground by simple fixing means 503.

FIG. 38 is a configuration diagram showing a case where the space structure is used for a planetary exploration ship. When used for a planetary exploration ship, the power generation module 300 is used for an electric propulsion device or for supplying power onboard. The mother ship 601 has a tether 6
02, a plurality of power generation modules 300 are connected. An arm 603 extends from the side surface of the mother ship 601, and an electric propulsion device 604 is provided at an end of the arm 603. Examples of the electric propulsion device 604 include a plasma jet, an ion engine, a colloid engine, an MPD arc jet, and a microwave accelerator.

The electric power generated by the power generation module 300 in response to the sunlight is sent to the mother ship 601 and the electric propulsion device 60
4 for power use. It is also used as electric power for electric equipment in the mother ship 601. According to such a configuration, regardless of the traveling direction of the mother ship 601, the solar radiation is always applied to the power generation module 300. Further, since it is not necessary to control the attitude of the power generation module 300, it is unlikely to break down during space navigation.

[Form of Space Structure] FIG. 39 is an explanatory view showing another form of space structure. This space structure 700
Has a deployment structure different from that of the space structure 100. As shown in FIG.
Is divided into a north side (reference number "N") and a south side (reference number "S") of the hemisphere, and a reference code "A" is given to a module M serving as a center. Are connected to each other, and a module M having a code "C" on two sides of the module MB having a code "B"
C are connected, and the outermost modules MC are partially connected to each other. Further, the space structure 700 is stacked and stored in a hexagonal prism state, as in the above. Table 3 shows the stacking order. In the same table, the lamination order is shown in the left column, and the corresponding module M is shown in the middle column. The order of development is shown in the right column.

[0090]

[Table 3]

In the space structure 700, the modules are connected to each other by the same joint J as described above, and the respective sides of the modules MA to MC to be connected during the unfolding operation are similar to the above. A coupler C is provided (all are not shown). By expanding the modules M in such an expansion order, each module can be expanded without interference. In addition,
The position movement before and after the deployment as described above is not so small as in the above-mentioned space structure, but not so much as to hinder the work.

FIG. 40 is a perspective view showing a space structure having 120 surfaces. This space structure 710 is a C120
It is constructed using a carbon allotrope as a model.
This is a configuration in which the zero-type space structure 100 is joined by a central beam 711. Other configurations are similar to those of the above-mentioned space structure 100.
The description is omitted here. This space structure 710 can be used as a solar power generation satellite by installing a solar cell panel 712 on the 120 surface, or can be used as a space base. Alternatively, the solar cell panel 712 may be installed in one sphere, and a power transmission antenna for transmitting power in the form of a high-frequency microwave may be installed inside the other sphere (not shown). In order to deploy the space structure 710, first, both spherical parts are separately deployed (FIGS. 14 to 14).
After that, they may be moved to predetermined positions, and then connected by the center beam 711.

FIG. 41 is a perspective view showing a donut-shaped space structure. Space structure 720 shown in FIG.
Is a configuration in which the hexagonal module M is combined in a donut shape. As described above, this space structure 720
The solar cell panel 721 may be installed on the module M of the above to be used as a solar power generation satellite,
It may be used as a space station. Furthermore, it can be used in combination with the space structure 100 of the C60 type. For example, as shown in FIG.
The solar cell panel 30 is attached to the space structure 100 of the C60 type.
1 and installed on both sides of the C200 type space structure 720, and a partition of the module of the C200 type space structure 720 is provided with a partition 722 to be a manned satellite.

[0094]

As described above, the space structure of the present invention (Claim 1) has a plurality of hexagonal modules each having a circular section in cross section and constituting a hollow shell-like structure. A rotation connection means for rotatably connecting the sides of the module, and a connection fixing means for connecting and fixing the sides when the sides of the module are in contact with each other according to the expansion of the module. When the module is unfolded and deformed into a hollow shell-like structure from the folded state, the rotational connection means and the connection fixing means are combined and attached to the module so that the modules do not interfere with each other. For this reason, it is possible to reliably deploy the module in outer space, and when launching into outer space, it can be compactly stored by folding with the rotating connection means, while on the other hand, when launching into outer space, the module can be deployed. Then, by fixing the modules together by the connecting and fixing means, a space structure having a hollow shell structure as a whole can be formed. The space structure thus formed is structurally stable and has high strength, and can be used for various applications.

The space structure of the present invention (Claim 2)
Since the module has a frame structure composed of six beams and is provided with wires and other reinforcing members for connecting the nodes of the frame structure, the rigidity of the space structure can be improved.

The space structure of the present invention (Claim 3)
Since the joints are provided at the nodes of the frame so as to be foldable in the plane, the space structure can be folded more compactly.

The space structure of the present invention (Claim 4)
In the frame structure composed of six beams, a center joint is provided at the center of two pairs of opposing beams, and an expansion / contraction means that expands and contracts by an actuator between joint portions shared by the two pairs of beams is provided. As a result, the space structure can be folded more compactly.

Further, the space structure of the present invention (Claim 5)
Then, in deploying the module, the force balance at the time of deployment under zero gravity is determined by simulation, and based on the deployment pattern in which the positional displacement of the space before and after deployment is reduced, the module has the rotational connection means and the fixed connection means. , The movement of the module during deployment can be extremely small.

The space structure of the present invention (claim 6)
Then, among the twenty hexagonal modules that can form a spherical shell shape, five modules (M1 to M5) are connected in a C-shape with each side, and the next module (M6)
Are connected to the outermost outer sides forming the C-shape of the module (M5) located at the rear end of the C-shape, and eight modules (M6-M15) including the module (M6) are connected to each side. Zigzag with and is connected in a substantially straight line,
The next five modules (M16 to M20) are connected in an inverted C-shape with each side, and the module (M15) positioned at the end of the linearly connected modules (M6 to M15) is A plurality of rotary connecting means for connecting to the outermost outer side constituting the inverted C-shape of the module (M16) located at the tip of the C-shape; And a plurality of connection fixing means for connecting the two. For this reason, from the state in which the first module (M1) to the last module (M20) are sequentially stacked, for example, the rotation connecting means between the modules is changed to J1, J2,.
When represented by J19, first, J1, J2, J3, J5,
J4, J7, J6 one by one, then J8 and J9, J10 and J11, J12 and J13, J1
4 and J15, two each in the order of J16 and J17,
Then, by expanding the modules one by one in the order of J18 and J19, the modules can be expanded without causing interference. According to the exemplary deployment method, it is possible to further suppress the movement of the space structure before and after deployment.

In the space structure deployment system according to the present invention (claim 7), the space structure and the temporary holding means for holding the module in a folded state and releasing the holding based on a release command. Deployment means for deploying the module around the rotation connection means based on the deployment command, and control means installed on the space structure and issuing a release command and a deployment command to the holding means and the deployment means. Therefore, the space structure can be deployed in outer space.

In the space structure deployment system according to the present invention (claim 8), the communication means for performing communication between the space structure and the ground, and the control means provided on the ground and via the communication means. And a command means for instructing the space structure to issue a release command or a deployment command, so that the space structure can be deployed remotely from the ground.

The space structure of the present invention (claim 9)
Then, by installing a pressurized block of an inflatable membrane structure with an available internal space in the space structure,
The space structure can be used as a protection structure for the pressurized block. For example, by providing a protective wall on the module, it is possible to protect the pressurized block from collisions such as space debris. Further, by providing the radiation shielding plate, it becomes possible to suppress the exposure in the pressurized block.

In the solar power module of the present invention (claim 10), since a solar cell panel is installed on the module of the space structure, the solar power can be generated without controlling the attitude.

In the solar power generation module according to the present invention, the space structure module is provided with a plurality of reflecting mirrors, and a servo motor for controlling the angles of these reflecting mirrors is provided. Since the solar cell panel is installed inside the device, light can be focused on the solar cell panel using a larger space structure.

Further, in the solar power module of the present invention (claim 12), since the solar cell panel is installed on a small space structure module installed inside the space structure, the light is condensed by a reflector. The collected sunlight is photoelectrically converted by a solar cell panel installed in a small space structure, so that the light collection efficiency is improved.

The solar power generation satellite according to the present invention (Claim 1)
3) includes a plurality of solar power generation modules, a main structure supporting the solar power generation satellite, and power transmission means for converting power generated by the solar power generation satellite into high-frequency microwaves and transmitting power from an antenna. The power generated by the solar power generation module can be transmitted to the ground in the form of high-frequency microwaves.

The spacecraft of the present invention (claim 14)
Is mounted on the mother ship, and the solar power module mounted on the mother ship for space navigation and the solar power module mounted on the mother ship,
Since the electric power propulsion unit is provided with a propulsion force by the electric power generated by the solar power generation module, it is possible to prevent the energy supply from being disabled due to failure.

[Brief description of the drawings]

FIG. 1 is a perspective view showing a space structure according to a first embodiment of the present invention.

FIG. 2 is a developed plan view showing the space structure shown in FIG. 1;

FIG. 3 is an explanatory diagram illustrating an example of a module illustrated in FIG. 1;

FIG. 4 is an explanatory view showing a modified example of a mounting form of a reinforcing member.

FIG. 5 is an explanatory view showing the joint structure shown in FIG. 1;

FIG. 6 is a perspective view showing the structure of the coupler shown in FIG.

FIG. 7 is an explanatory diagram showing a state in which the module is folded.

FIG. 8 is an explanatory view showing a notch pin type temporary fixing device.

FIG. 9 is an explanatory diagram showing a modification of the module structure.

FIG. 10 is an explanatory diagram showing a storage configuration of the space structure.

FIG. 11 is a configuration diagram showing a space structure deployment system.

12 is a functional block diagram of the deployment system shown in FIG.

FIG. 13 is a flowchart showing the operation of the deployment system.

FIG. 14 is an explanatory diagram showing a deployment process (before deployment) of the space structure.

FIG. 15 is an explanatory diagram showing a deployment process (step 1) of the space structure.

FIG. 16 is an explanatory diagram showing a deployment process (step 2) of the space structure.

FIG. 17 is an explanatory diagram showing a deployment process (step 3) of the space structure.

FIG. 18 is an explanatory diagram showing a deployment process (step 4) of the space structure.

FIG. 19 is an explanatory diagram showing a deployment process (step 5) of the space structure.

FIG. 20 is an explanatory diagram showing a deployment process (step 6) of the space structure.

FIG. 21 is an explanatory diagram showing a deployment process (step 7) of the space structure.

FIG. 22 is an explanatory diagram showing a deployment process (step 8) of the space structure.

FIG. 23 is an explanatory diagram showing a deployment process (step 9) of the space structure.

FIG. 24 is an explanatory diagram showing a deployment process (step 10) of the space structure.

FIG. 25 is an explanatory diagram showing a deployment process (step 11) of the space structure.

FIG. 26 is an explanatory diagram showing a deployment process (step 12) of the space structure.

FIG. 27 is an explanatory diagram showing a deployment process (step 13) of the space structure.

FIG. 28 is an explanatory diagram showing a deployment process (step 14, after deployment) of the space structure.

FIG. 29 is an explanatory diagram showing an application example of the space structure.

FIG. 30 is a configuration diagram showing a solar power generation satellite.

FIG. 31 is a perspective view showing a modification of the solar power generation satellite.

FIG. 32 is a perspective view showing another modification of the solar power generation satellite.

FIG. 33 is a perspective view showing another modification of the solar power generation satellite.

FIG. 34 is a configuration diagram showing a modification of the power generation module.

FIG. 35 is an explanatory view showing a light collecting device of the power generation module shown in FIG. 34.

FIG. 36 is a perspective view showing a modification of the power generation module shown in FIG.

FIG. 37 is an explanatory diagram showing a space base using a space structure.

FIG. 38 is a configuration diagram showing a case where the space structure is used for a planetary exploration ship.

FIG. 39 is an explanatory view showing another form of the space structure.

FIG. 40 is a perspective view showing a space structure having 120 surfaces.

FIG. 41 is a perspective view showing a donut-shaped space structure.

FIG. 42 is a configuration diagram illustrating an example of a conventional solar power generation satellite.

43 is a configuration diagram showing the power generation module shown in FIG. 42, (a) is a front view, and (b) is a cross-sectional view.

[Explanation of symbols]

 REFERENCE SIGNS LIST 100 space structure 101 beam 102 coupling portion 103 wire M1 to M20 module J1 to J19 joint C1 to C11 coupler 200 deployment system 201 storage device 204 processing device 205 power supply 206 communication system 207 communication antenna 208 electric circuit 209 ground command base 210 remote control System 211 Communication antenna 212 Communication system 213 Remote control program 214 Computer 217 Internet 218 Solar panel 300 Power generation module 301 Solar panel 350 Solar power satellite 351 Arm 352 Main structure 353 Power transmission device 354 Power transmission antenna

 ────────────────────────────────────────────────── ─── Continuing from the front page (72) Inventor Tomohiro Akagi 2-1-1, Shinhama, Arai-machi, Takasago City, Hyogo Prefecture Inside the Takasago Research Laboratory, Mitsubishi Heavy Industries, Ltd. (72) Inventor Satoshi Tahara 1-8-1, Koura, Kanazawa-ku, Yokohama-shi (72) Inventor Katsumi Kito 10 Oemachi, Minato-ku, Nagoya City Mitsubishi Heavy Industries, Ltd.10 Nagoya Aerospace Systems Works (72) Inventor Masayuki Niitsu 10 Oecho, Minato-ku, Nagoya City Nagoya Aerospace Systems Works, Mitsubishi Heavy Industries, Ltd. (72) Inventor Koji Yamawaki 2-1-1 Sengen, Tsukuba, Ibaraki Pref. Advanced Technology Research Division, Space Development Agency of Japan F-term (reference) 5F051 BA02 JA08 JA09 JA10 JA14

Claims (15)

[Claims] 1. A plurality of hexagonal modules each having a circular section in cross section and being a component of a hollow shell structure, rotation connection means for rotatably connecting sides of the modules, and And connecting and fixing means for connecting and fixing the sides when the sides of the module are in contact with each other. The module is expanded and deformed into a hollow shell-like structure from a state in which the module is folded by the rotary connecting means. A space structure wherein the rotational connection means and the connection fixing means are combined and attached to the module so that the modules do not interfere with each other. 2. The module according to claim 1, wherein the module has a frame structure composed of six beams, and a wire or other reinforcing member for connecting nodes of the frame structure is provided. Space structure. 3. The space structure according to claim 2, wherein a joint is provided at a node of the frame so as to be foldable in a plane. 4. A frame structure comprising six beams, wherein a central joint is provided at the center between two pairs of opposing beams, and an expansion / contraction is performed by an actuator between joint portions shared by the two pairs of beams. 3. The method according to claim 2, wherein the means is passed.
Or the space structure according to any one of 3. 5. When deploying the module, a force balance at the time of deployment under zero gravity is obtained by simulation, and based on a deployment pattern in which the positional displacement of the space before and after deployment is reduced, the module is connected to the rotational connection means. The space structure according to any one of claims 1 to 4, wherein the space structure is mounted in combination with a fixed connection means. 6. It has 20 hexagonal modules capable of forming a spherical shell shape, and 5 of the modules (M1 to M5) are connected to each other in a C-shape with one side thereof. M6) is connected to the outermost side of the C-shaped module (M5) located at the rear end of the C-shape, and includes eight modules (M6) including the module (M6).
6 to M15) are connected in a substantially linear manner in a zigzag with each side, and the next five modules (M16 to M20) are connected in an inverted C-shape with each side, and the modules (linearly connected) M6 to M15), a plurality of rotation connection means for connecting the module (M15) located at the end with the outermost outer side constituting the inverted C-shape of the module (M16) located at the tip of the inverted C-shape; A plurality of connecting and fixing means for connecting the modules which are in contact with each other when each module is unfolded by the rotary connecting means.
6. The space structure according to any one of 5. 7. A space structure according to claim 1, wherein said module holds a module in a folded state, and temporarily holds said module based on a release command. A space structure comprising: deployment means for deploying a module around a rotation connection means; and control means installed on the space structure and issuing release and deployment instructions to the holding means and the deployment means. Body deployment system. 8. Communication means for communicating between the space structure and the ground, and command means installed on the ground for instructing the control means to issue a release command or a deployment command via the communication means. The space system deployment system according to claim 7, comprising: 9. The space structure according to claim 1, further comprising a pressurizing block of an inflatable film structure having an available internal space inside the space structure. body. 10. A solar power module comprising a space structure module according to claim 1 and a solar cell panel installed thereon. 11. A space structure module according to claim 1, further comprising a plurality of reflectors, a servo motor for controlling the angles of the reflectors, and a solar cell inside the space structure. A solar power module comprising a panel. 12. The solar power module according to claim 11, wherein the solar cell panel is installed in a small space structure module installed inside the space structure. 13. A plurality of solar power generation modules according to claim 10; a main structure supporting said solar power generation satellite; and a power generated by said solar power generation satellite converted into high-frequency microwaves and transmitted from an antenna. What is claimed is: 1. A solar power generation satellite comprising: a power transmission means for transmitting power. 14. A mother ship for space navigation, a solar power module according to claim 10 attached to the mother ship, and electricity that is similarly mounted on the mother ship and obtains propulsion by electric power generated by the solar power module. A spacecraft comprising a propulsion device. 15. Twenty hexagonal modules are connected to each other by a plurality of rotational connecting means, and five of the modules (M
1 to M5) are connected in a C-shape with each side, and the next module (M6) is connected to the outermost outer side of the C-shaped module (M5) located at the rear end of the C-shape. Eight modules (M6) including the module (M6)
To M15) in a zigzag with one side connected in a substantially straight line, and the next five modules (M16 to M20)
Are connected in an inverted C-shape with each side, and the module (M15) positioned at the end of the modules (M6 to M15) connected linearly is replaced with the module (M16) positioned at the tip of the inverted C-shape. From the state in which the first module (M1) to the last module (M20) are sequentially stacked and connected to the outermost side forming the inverted C-shape, the rotational connection means between the modules is changed to J1, J2,.
18, J19, one by one in the order of J1, J2, J3, J5, J4, J7, J6, then J8 and J9, J10 and J11, J12
And J13, J14 and J15, J16 and J1
7. A method for deploying a space structure, comprising deploying the modules two by two in the order of 7, and one by one in the order of J18 and J19.