RU2678296C2 - Component development system - Google Patents

Abstract

FIELD: astronautics.SUBSTANCE: group of inventions relates to external deployable components of a spacecraft, for example, solar panels or antennae, mounted primarily on small satellites. Apparatus comprises a group of panels folded tightly to each other on a side of a spacecraft, a group of flexible members connected to the panels, and an interface system associated with said groups. Interface system is configured to move the group of panels from the folded configuration to a deployed configuration when the group of flexible members is extended from the spacecraft.EFFECT: technical result is aimed at increasing the compactness of deployable spacecraft components in a folded state.12 cl, 38 dwg

Description

FIELD OF THE INVENTION

The present invention relates generally to the deployment of components, and more particularly to systems for deploying components. In particular, the present invention relates to a method and apparatus for deploying components from a spacecraft.

BACKGROUND

Spacecraft while in space are used for various purposes. A spacecraft, as used in this application, is a vehicle, ship or car, configured to perform many operations in space. A spacecraft may be self-propelled space vehicles intended for short-term space flights, or space vehicles that are configured to be in space for a long time. In other instances, the spacecraft may be a space station, satellite, or some other suitable structure.

Satellites are a type of spacecraft that is gaining popularity and is used to perform many different operations while in space. These satellites are used in a wide variety of areas. For example, without limitation, satellites can be used for navigation, communications, environmental monitoring, weather forecasting, research, broadcasting and other tasks. Most homes, businesses, government organizations, and other users can use satellites every day to watch entertainment, communications, gather information, and more.

Satellites are usually launched into orbit using launch vehicles, and they can be launched from a moving platform or from the ground. The size of the satellite determines the value of the thrust required by the launch vehicle to launch the satellite. To reduce the cost of putting a satellite into orbit, operators reduce the size of the satellite. As the satellite size decreases, the value of the traction force required to launch the satellite also decreases. Reducing the size of the satellite, as well as the type of launch vehicle used, reduces the cost of launching the satellite into orbit. As a result, more and more organizations are reducing the size of satellites used.

Depending on the task of the satellite, components can be deployed from the satellite during various stages of operation. These components may take the form of panels. For example, at least one of a control device, a braking device, an antenna, a solar panel, and some other type of component can be deployed from a satellite.

In some cases, solar panels are used by a satellite to receive energy from the sun. In this illustrative example, a solar panel is a group of photovoltaic modules electrically connected to a supporting structure. These solar panels can be stored next to the satellite’s body during launch and subsequent deployment while the satellite is in orbit. For example, solar panels stored close to the side of the satellite can be expanded. Similarly, other components, such as antennas, are stored next to the satellite’s body during transport from Earth to orbit and during subsequent deployment.

When using small satellites, the satellite capacity may not be large enough. For example, a satellite may need more energy to carry out work operations, which can be obtained by solar panels stored in and deployed from this satellite. In other examples, the size of the antenna, which can be deployed from a satellite, may not provide the necessary amount of communication. As a result, the satellite may not function as needed. Thus, it would be necessary to create a device and method that solve at least some of the problems described above, as well as solve other possible problems.

SUMMARY OF THE INVENTION

In one illustrative embodiment, the device comprises a panel group, a group of flexible elements connected to a group of panels, and an interface system associated with a panel group and a group of flexible elements. The panel group is located in a folded configuration close to the many sides of the spacecraft. The interface system is configured to move a group of panels from a folded configuration to an expanded configuration when this group of flexible elements is pulled out of a spacecraft.

In another illustrative embodiment, a method for deploying a panel group is provided. A group of flexible elements are advanced from a spacecraft. A group of panels is moved from a folded configuration to an expanded configuration using an interface system so that a group of flexible elements extends from the spacecraft.

In another illustrative embodiment, the antenna system comprises a group of flexible rods, an interface system, and a deployment mechanism. A group of flexible rods is connected to a group of panels with an antenna reflector located in a folded configuration close to the many sides of the satellite. The interface system is connected to a group of panels with an antenna reflector and a group of flexible rods. The interface system is configured to move a group of panels with an antenna reflector from a folded configuration to an expanded configuration. The deployment mechanism is configured to extend a group of flexible rods as a group of panels with an antenna reflector moves from a folded configuration to an expanded configuration. In another illustrative embodiment, the control system comprises a platform and a movement system. The platform is located in the satellite body. The movement system is connected to the platform and is configured to change the orientation of the platform for maneuvering a group of components in space after this group of components is deployed from a satellite.

In another illustrative embodiment, the antenna system comprises a reflector for parabolic antennas, a group of flexible rods, an interface system, and a deployment mechanism. The reflector for parabolic antennas is located in a folded configuration close to the many sides of the satellite. A group of flexible rods is connected to a reflector for parabolic antennas. The interface system is connected to a reflector for parabolic antennas and a group of flexible rods. The interface system is configured to move the reflector for parabolic antennas from a folded configuration to an expanded configuration. The deployment mechanism is configured to extend a group of flexible rods as the reflector for parabolic antennas moves from the folded configuration to the deployed configuration.

In another illustrative embodiment, the solar panel comprises a solar battery. The solar battery contains a first set of solar modules and a second set of solar modules. The first set of solar modules is configured to fold along the longitudinal axis, and the second set of solar modules is configured to fold along the diagonal axis to form a folded configuration of the solar battery.

In various embodiments of the present invention, the properties and functions can be provided independently of each other or can be combined in some other embodiments, in which you can see additional detailed information with reference to the following section "The implementation of the invention" and the section "Drawings".

BRIEF DESCRIPTION OF THE DRAWINGS

New features, considered as characteristics of illustrative embodiments, are defined in the attached claims. However, illustrative embodiments, as well as the preferred mode of use and their additional tasks and features will be better understood by reference to the following detailed description of an illustrative embodiment of the present invention when read by reference to the accompanying drawings.

In FIG. 1 shows an image of a spacecraft in accordance with an illustrative embodiment.

In FIG. 2 shows a block diagram of an aerospace vehicle in accordance with an illustrative embodiment.

In FIG. 3 shows a block diagram of an interface system in accordance with an illustrative embodiment.

In FIG. 4 shows a block diagram of a panel in accordance with an illustrative embodiment.

In FIG. 5 shows an image of a deployment system for a satellite in accordance with an illustrative embodiment.

In FIG. 6 is a view of a deployment system for a satellite in accordance with an illustrative embodiment.

In FIG. 7 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment.

In FIG. 8 is an illustration of a deployment system for a satellite in accordance with an illustrative embodiment.

In FIG. 9 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment.

In FIG. 10 is a view of a folded panel configuration for a satellite in accordance with an illustrative embodiment.

In FIG. 11 shows an image of a panel deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 12 is a view of a panel deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 13 shows an image of a panel deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 14 is an image of a panel deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 15 is an image of a panel deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 16 is a view of a panel deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 17 shows an image of a panel deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 18 is a sectional view of a panel in an expanded configuration and an interface system in accordance with an illustrative embodiment.

In FIG. 19 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment.

In FIG. 20 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment.

In FIG. 21 is a cross-sectional view of a satellite deployment system in accordance with an illustrative embodiment.

In FIG. 22 shows an image of a satellite with a group of panels deployable from a satellite, in accordance with an illustrative embodiment.

In FIG. 23 shows an image of a satellite with a group of panels deployable from a satellite, in accordance with an illustrative embodiment.

In FIG. 24 shows an image of a satellite in accordance with an illustrative embodiment.

In FIG. 25 shows an image of a satellite in accordance with an illustrative embodiment.

In FIG. 26 shows an image of a satellite with a group of components deployed from a satellite in accordance with an illustrative embodiment.

In FIG. 27 shows an image of a solar panel in accordance with an illustrative embodiment.

In FIG. 28 is a view of a portion of a solar cell in accordance with an illustrative embodiment.

In FIG. 29 shows an image of a solar module in accordance with an illustrative embodiment.

In FIG. 30 shows an image of a solar module in accordance with an illustrative embodiment.

In FIG. 31 is an image of a control system for a satellite in accordance with an illustrative embodiment.

In FIG. 32 is a cross-sectional view of a satellite control system in accordance with an illustrative embodiment.

In FIG. 33 is a cross-sectional view of a satellite control system in accordance with an illustrative embodiment.

In FIG. 34 is a cross-sectional view of a satellite control system in accordance with an illustrative embodiment.

In FIG. 35 is a flowchart of a panel group deployment process in accordance with an illustrative embodiment.

In FIG. 36 is a flowchart of a process for deploying a group of panels from a spacecraft in accordance with an illustrative embodiment.

In FIG. 37 is a depiction of a method for manufacturing and servicing a spacecraft in the form of a flowchart in accordance with an illustrative embodiment.

In FIG. 38 shows an image of a spacecraft in the form of a block diagram according to which an illustrative embodiment may be implemented.

DETAILED DESCRIPTION OF THE INVENTION

Illustrative implementation options show and take into account one or more different principles. For example, illustrative embodiments show and take into account that it may be necessary to increase the size of components that are deployed from a satellite. For example, illustrative embodiments show and take into account that with large solar panels deployed from a satellite, more energy can be generated by the satellite. This increase in energy may provide additional operational capabilities of the satellite and / or increase the life of the satellite. Similarly, illustrative embodiments show and take into account that increasing the size of an antenna deployed from a satellite improves the communication capabilities of that satellite.

However, illustrative embodiments show and take into account that some currently deployed deployment systems for deploying components from a satellite contain designs that can increase the weight of the satellite more than necessary and / or require more storage space in the satellite than necessary. For example, some systems currently in use contain panels that are stored close to the sides of the satellite. These panels are deployed from a spacecraft after the satellite is put into orbit. Console systems for unfolding panels may not hold the panels as inelastic or flat as necessary, and may add more weight to the satellite than necessary. An increase in satellite weight is undesirable since the cost of launching a satellite increases as the weight of this satellite increases. In addition, the size of the panels, properly stored close to the side of the satellite, may be limited by the size of the satellite.

Illustrative implementation options further show and take into account that some currently deployed deployment systems for deploying components from a satellite may not be implemented as needed. For example, when solar panels are deployed, some deployment systems may not even deploy solar panels so that these solar panels are laid horizontally. As a result, solar panels may not capture the required amount of light energy from the sun.

Thus, illustrative embodiments provide a method and apparatus for deploying a panel group. In one illustrative embodiment, the device comprises a group of panels located in a folded configuration close to the many sides of the spacecraft, a group of flexible elements connected to a group of panels, and an interface system associated with a group of panels and a group of flexible elements. The interface system is configured to move a group of panels from a folded configuration to an expanded configuration when this group of flexible elements is pulled out of a spacecraft.

In the figures, and in particular in FIG. 1, an image of a spacecraft in accordance with an illustrative embodiment is shown. Spacecraft 100 takes the form of satellite 102.

As shown, the satellite 102 comprises a housing 104 with a plurality of sides 106. The plurality of sides 106 comprises a side 108, a side 110, a side 112 and a side 114. The satellite 102 also includes an upper plate 116 and a lower plate 118.

In this illustrative example, satellite 102 comprises a deployment system 120. The deployment system 120 is associated with the housing 104 of the satellite 102. In this illustrative example, the deployment system 120 is located in the main body 122 of the housing 104. The deployment system 120 is configured to deploy a group of components (not shown in this view) from the satellite 102.

In FIG. 2 shows a block diagram of an aerospace vehicle in accordance with an illustrative embodiment. In this shown example, an aerospace vehicle 200 is a vehicle that is configured to fly in the Earth’s atmosphere and / or outside the Earth’s atmosphere. For example, an aerospace vehicle 200 may be selected from one of the following: an aircraft, a spacecraft, and some other suitable type of aerospace vehicle.

As shown, the aerospace vehicle 200 takes the form of a spacecraft 201. In this shown example, the spacecraft 100 of FIG. 1 is an example of a physical implementation of a spacecraft 201 shown in block form in this figure. In this illustrative example, spacecraft 201 is a satellite 202.

Satellite 202 is an object that can be placed in orbit. In this illustrative example, satellite 202 is an artificial satellite. In other words, satellite 202 is a human-made satellite.

In this shown example, satellite 202 may take various forms. For example, satellite 202 may take the form of at least one of the following: astronomical satellite, research satellite, communications satellite, navigation satellite, reconnaissance satellite, weather satellite, and other suitable types of satellites.

As used in this application, the phrase “at least one of”, when used with a list of objects, means that various combinations of one or more of the listed objects can be used, and only one of these objects in the list may be necessary . An object may be a specific object, thing or category. In other words, the phrase “at least one of” means that any combination of objects or many of these objects can be used from this list, but not all of these objects in this list may be necessary.

For example, “at least one of object A, object B and object C” can mean object A, object A and object B, object B, object A, object B and object C or object B and object C. In some cases, “ at least one of object A, object B and object C "can mean, for example, without limitation, two objects A, one object

B and ten objects C, four objects B and seven objects C, or some other suitable combination.

In this illustrative example, satellite 202 is a small satellite 204. Small satellite 204 is a satellite with a weight of from about 1 gram to about 500 kilograms. In other illustrative examples, satellite 202 may be smaller or larger, depending on the particular implementation. Small satellite 204 can be selected as one of the following: a miniature satellite, a microsatellite, a nanosatellite, a cubic satellite, and other suitable types of small satellites.

In this shown example, satellite 202 comprises a housing 206 with multiple sides 208. As used herein, the phrase “multiple” objects means one or more objects. For example, the plurality of sides 208 represents one or more parties.

As shown, housing 206 forms a satellite structure 202 that provides mechanical support to satellite 202. In this illustrative example, housing 206 is formed from a material selected as one of the following: metal, metal alloy, composite material, other suitable types of materials, and combinations of these materials.

As shown, the deployment system 210 is physically connected to the housing 206 of the satellite 202. The first component, such as the deployment system 210, can be considered as a component physically connected to the second component, such as the housing 206 of the satellite 202, by attaching to this second component, connecting to this second component by means of a binder, fastening to this second component, welding with this second component, bonding to this second component, connecting to this second component in some other suitable way, or x combination. The first component can also be connected to the second component using the third component. In addition, the first component can be considered as a component associated with the second component by forming part of this second component, continuing this second component, or a combination thereof.

In this shown example, the deployment system 210 is located in the housing 206 of the satellite 202. The deployment system 210 comprises a group of components 212, a group of flexible elements 214, an interface system 216, and a deployment mechanism 218. As used in this application, a “group” of objects is one or more objects. For example, component group 212 represents one or more components. Components in component group 212 may contain the same or different types of components.

In this illustrative example, the deployment system 210 is configured to deploy a group of components 212 from satellite 202. For example, the deployment system 210 is configured to move a group of components 212 from inside 220 of satellite 202 to outside 222 of satellite 202.

In other illustrative examples, deployment system 210 moves a group of components 212 stored in the outer portion 222 of satellite 202 to a desired position. For example, deployment system 210 may decompose a group of components 212.

In this shown example, the group of components 212 contains a group of panels 224. The group of panels 224 are structures configured to take the required shape when deployed from satellite 202. In the illustrative example, the group of panels 224 can take the form of flat, rectangular structures when deployed from satellite 202.

As shown, the panel group 224 includes a panel 226. The panel 226 can take many different forms. For example, a panel 226 in a panel group 224 may be selected as one of the following: a solar panel, an antenna, a braking device, and some suitable types of panels.

In this illustrative example, the deployment system 210 is configured to move the panel group 224 between the folded configuration 228 and the expanded configuration 230. The folded configuration 228 is a structure from the panel group 224 in which each of the panel group 224 is bent over itself so that at least as part of each panel group 224 closes another part of each panel group 224. In the folded configuration 228, each of the panel group 224 can be folded multiple times to fit snugly into sensor body 206 of the satellite 202.

The expanded configuration 230 is a structure in which the panel group 224 is extended. In this illustrative example, the expanded configuration 230 is an expanded construction of a panel group 224.

A group of panels 224 in a folded configuration 228 is associated with a plurality of sides 208 of a satellite 202. For example, a panel 226 in a group of panels 224 is folded close to a side 232 in a plurality of sides 208 of a housing 206 of a satellite 202.

In this shown example, a group of flexible elements 214 is connected to a group of panels 224. A group of flexible elements 214 is configured to extend a group of panels 224 from a satellite 202.

In this illustrative example, the group of flexible elements 214 may take various forms. For example, a group of flexible members 214 may take the form of a rod, cable, cord, rope, tape, wire, bundle, line, or some other suitable type of flexible member.

A group of flexible elements 214 may be formed from structures with various types of section shapes. For example, a group of flexible members 214 may be formed from structures with a sectional shape selected from at least one of the following: circular shape, triangular shape, rectangular shape, irregular shape, hexagonal shape, or some other suitable type of sectional shape. The group of flexible members 214 may have the same sectional shape or sectional shape different from the others.

In this illustrative example, a group of flexible members 214 can also be formed from many different types of materials. For example, a group of flexible members 214 may be formed from a material selected from at least one of the following: composite material, carbon fiber, metal, metal alloy, fiberglass, polymer, or some other suitable material. The material selected for the group of flexible elements 214 is a material that allows the group of flexible elements 214 to be bent as needed without breaking while supporting the group of panels 224 and the interface system 216. In this illustrative example, the group of flexible elements 214 may take the form of a group of flexible rods 215.

As shown, the interface system 216 is physically connected to the group of panels 224 and the group of flexible elements 214. The interface system 216 is configured to move the group of panels 224 from the folded configuration 228 to the expanded configuration 230 when this group of flexible elements 214 is pulled out of the satellite 202. For example, the interface system 216 keeps the panel group 224 substantially flat when the panel group 224 is in the expanded configuration 230.

In this illustrative example, the deployment mechanism 218 is coupled to the satellite body 206 202. For example, the deployment mechanism 218 may be located in the satellite body 206 202. The deployment mechanism 218 is a mechanical device configured to extend a group of flexible members 214 from the interior 220 of the satellite 202 to the outer part 222 of the satellite 202 so that the panel group 224 moves from the folded configuration 228 to the expanded configuration 230.

As shown, deployment mechanism 218 takes various forms. For example, the deployment mechanism 218 may take the form of a spring 233, an engine 235, and some suitable types of deployment mechanisms.

When the deployment mechanism 218 takes the form of a spring 233, this spring 233 can be selected from one of the following: compression spring, torsion spring, travel spring, belt spring, flat spring, coil spring and other suitable types of springs. When the deployment mechanism 218 takes the form of an engine 235, a group of flexible members 214 can be deployed from the satellite 202 in a controlled manner. In other words, the deployment of the group of flexible members 214 can be controlled based on the speed of the engine 235. The speed of the engine 235 can be predefined or adjusted during the deployment of the group of flexible members 214.

In other illustrative examples, deployment mechanism 218 may comprise many types of devices in addition to, or instead of, spring 233 and motor 235. In this illustrative example, the deployment mechanism 218 takes the form of a torsion spring 234. Torsion spring 234 is a flexible, resilient device that retains mechanical energy when twisted.

In this illustrative example, a torsion spring 234 is configured to wind it around a drum 236 located in satellite housing 206 202. A group of flexible elements 214 is also configured to wind it onto drum 236. Once released, a torsion spring 234 is unwound from drum 236 to extend the group flexible elements 214. Thus, the group of panels 224 is deployed using the group of flexible elements 214 and torsion spring 234. Alternatively, when an engine 235 is used in deployment mechanism 218, this engine 235 rotates drum 236 to extend a group of flexible members 214.

In this shown example, a group of panels 224 is connected to the housing 206 of the satellite 202 at a plurality of contact points 238 along the housing 206. A plurality of contact locations 238 are located on a plurality of sides 208 of the satellite 206 of the satellite 202. For example, the panel 226 may be connected to the side 232 of the housing 206 in 240 pins.

Each of the group of panels 224 can be connected to the housing 206 in a plurality of contact points 238 in several ways. For example, each of a group of panels 224 can be connected to a housing 206 at a plurality of contact points 238 using an adhesive. In one illustrative example, each of a group of panels 224 can be connected to a housing 206 at a plurality of contact points 238 using fasteners. These fasteners can be selected as one of the following: staples, pins, screws, couplers and other suitable types of fasteners.

As shown, the panel group 224 is adjacent to the plurality of sides 208 of the satellite 202 body 206 when this panel group 224 is located in the folded configuration 228. The multiple doors 242 are associated with the panel group 224 adjacent to the plurality of sides 208 of the satellite 206.

In this illustrative example, a plurality of doors 242 are configured to close a group of panels 224 located adjacent to a plurality of sides 208 of a housing 206 in a folded configuration 228, and to be opened so that a group of panels 224 moves from a folded configuration 228 to a deployed configuration 230. For example, a door 244 in a plurality of doors 242 is configured to open to move a panel 226 located close to side 232 of the housing 206 from a folded configuration 228 to a deployed configuration 230.

As shown, the housing 206 of satellite 202 also includes a control system 246. The control system 246 is configured to maneuver a group of components 212 in space after this group of components 212 has been deployed using deployment system 210. For example, control system 246 moves a group of panels 224 so that this group of panels 224 faces at least one of the following: the sun, an intended target, and some other suitable design in space. The control system 246 moves a group of panels 224 while simultaneously moving the satellite body 202 toward a structure, away from that structure, toward another object of interest, or to which combination thereof.

In one illustrative example, control system 246 necessarily moves a group of panels 224 in such a way that a flight target can be achieved. For example, the control system 246 can tilt the panel group 224 so that the panel group 224 receives the required number of electromagnetic signals.

In this illustrative example, the control system 246 comprises a platform 248 and a movement system 249. In other illustrative examples, control system 246 contains other components and combinations of these components. The components in the control system 246 are configured to change the orientation of the platform 248. In particular, the components in the control system 246 are configured to change the normal plane of the plane of the deployed platform 248 so that the group of panels 224 can indicate any number of areas of interest.

As shown, the platform 248 is located in the housing 206 of the satellite 202. In some illustrative examples, the platform 248 is located in the space between the group of panels 224 in the satellite 202. In this case, the platform 248 is attached to the group of panels 224 so that the movement of this platform 248 changes the configuration of the group panels 224.

In this shown example, the movement system 249 is configured to tilt the normal vector of platform 248 to its desired location. The movement system 249 may comprise, for example, without limitation, at least one gear-rack motor or some other suitable type of movement system.

In some illustrative examples, the control system 246 comprises at least one pivot point around which the platform 248 is tilted using the movement system 249. In yet another example, the control system 246 may include springs in addition to the movement system 249, which are extendable when the movement system 249 pulls and contracts when it is extended. In this case, the springs are designed to balance the actions of the movement system 249 and / or to reduce the torque from the side of this movement system 249. As a result, the size of the movement system 249 can be reduced.

When using the deployment system 210 for satellite 202, a group of panels 224 can be deployed from this satellite 202 as necessary. As a result, large panels can be stored behind multiple doors 242 in a folded configuration 228 compared to some of the systems currently in use. The group of flexible elements 214 and the interface system 216 provide the ability to extend the group of panels 224 as necessary so that each of the group of panels 224 is expanded.

In FIG. 3 shows a block diagram of an interface system in accordance with an illustrative embodiment. In this shown example, components in the interface system 216 of FIG. 2.

As shown, the interface system 216 comprises connectors 300 and a group of interaction designs 302. Connectors 300 are coupled to a panel group 224 and a group of flexible members 214 of FIG. 2. In this illustrative example, the connectors 300 are adapted to be attached to the panel group 224. For example, the connector 304 in the connectors 300 is configured to be attached to the panel 226 of FIG. 2 in panel group 224.

In this illustrative example, each one of the connectors 300 is attached to one of a group of panels 224 at various locations on the panel. For example, when three connectors 300 are connected to a panel 226, and the panel 226 is a rectangular structure, each of these connectors is attached to the corner of the panel 226. Thus, one corner of the panel 226 is mechanically and electrically attached to the housing 206 of the satellite 202 at 240 contact point FIG. 2, and each of the three connectors 300 is attached to the panel 226 in one of the three corners of the panel 226.

In this illustrative example, the connectors 300 are movable as the panel 226 is deployed from the satellite 202. For example, the connectors 300 move as the panel 226 is moved from the folded configuration 228 to the deployed configuration 230.

In this shown example, connectors 300 comprise channels 306. Each of connectors 300 contains one of channels 306. For example, connector 304 comprises channel 308. Channels 306 are configured to receive a group of interaction designs 302.

In this illustrative example, the channels 306 may be channels with the same size or different sizes. As an example, one of the channels 306 may be larger than the other channel from the channels 306. When three or more channels 306 are made, each channel may have a different size.

In this shown example, a group of interaction structures 302 is associated with a group of flexible elements 214. For example, a group of interaction structures 302 can be attached to a group of flexible elements 214 at various places along this group of flexible elements 214. In this illustrative example, each structure from a group of interaction structures 302 has a different size.

Channels 306 correspond to a group of interaction structures 302. In particular, one of the channels 306 corresponds to one of a group of interaction structures 302. For example, channel 308 in connector 304 corresponds to interaction structure 310 in a group of interaction structures 302.

The interaction structure 310 in the group of interaction structures 302 is configured to interact with the connector 304 in the connectors 300 to move the panel 226 in the panel group 224 from the folded configuration 228 to the expanded configuration 230. In this case, the channel 308 has such a shape that the interaction structure 310 can be received by channel 308 as necessary. As the interaction structure 310 is received by the channel 308, these two components help maintain the deployed configuration 230 of the panel 226.

In this illustrative example, the interaction structure 310 takes the form of a sphere 312. The sphere 312 interacts with the channel 308 in the connector 304. The channel 308 has such a shape that the sphere 312 is received by the channel 308. For example, the channel 308 may be in the shape of a cone.

A group of flexible members 214 extends through channels 306 of connectors 300. Thus, as a group of panels 224 are advanced, each of a group of interaction structures 302 interacts with a respective one of connectors 300 depending on the size and shape of each channel 306 in connectors 300 As a result, the interaction of the group of interaction structures 302 with the channels 306 in the connectors 300 moves each of the panel group 224 from the folded configuration 228 to the expanded configuration 230.

In FIG. 4 is a block diagram of a panel in accordance with an illustrative embodiment. In this shown example, components are shown that can form the panel 226 of FIG. 2.

As shown, panel 226 may be selected from a solar panel 400, antenna 402, or braking device 404. In other illustrative examples, other types of panels may be used for panel 226 depending on a particular implementation.

In this shown example, the solar panel 400 is configured to supply electrical energy to the satellite 202 of FIG. 2. The solar panel 400 creates electrical energy from the sun.

The solar panel 400 comprises a solar battery 406. In this illustrative example, the solar battery 406 is a structure of solar modules 408. The solar modules 408 comprise cells configured to convert solar radiation to direct current energy. The solar battery 406 comprises a first set of solar modules 410 and a second set of solar modules 412. The solar panel 400 with the solar battery 406 is configured to be folded and stored behind a plurality of doors 242 in FIG. 2.

In this illustrative example, the first set of solar modules 410 is configured to fold along the longitudinal axis 414 of the solar battery 406. The second set of solar modules 412 is configured to be folded along the diagonal axis 416 of the solar battery 406. The first set of solar modules 410 and the second set of solar modules 412 form folded configuration of the solar battery 406.

As shown, antenna 402 is a device that sends and receives signals from satellite 202. In some illustrative examples, antenna 402 may take the form of a parabolic antenna. A parabolic antenna is an antenna with a curved surface with a sectional shape in the form of a parabola, made with the possibility of directing radio waves. In other illustrative examples, the antenna 402 has a flat, rectangular structure similar to that of the solar panel 400.

Antenna 402 may be located adjacent to side 232 of housing 206 of satellite 202 of FIG. 2 during launch and subsequent deployment to perform satellite communications 202. These communications operations may take the form of telephone communications, television communications, radio communications, Internet operations, military communications, and other types of communications operations using different frequencies from the electromagnetic spectrum.

In this shown example, the braking device 404 is a structure used to slow down the movement of the satellite 202. The braking device 404 can be deployed during various stages of the satellite 202 to create braking.

The depiction of the aerospace vehicle 200 in FIG. 2 and the components in the aerospace vehicle 200 does not imply the introduction of physical or architectural limitations in the manner in which an illustrative embodiment can be implemented. In addition to or instead of the components shown, other components may be used. Some components are optional. In addition, blocks are presented to illustrate some functional components. One or more of these blocks may be combined, divided, or combined and divided into different blocks when implemented in an illustrative embodiment.

For example, although illustrative examples of an illustrative embodiment are described with respect to a spacecraft, the illustrative embodiment may be applied to other types of platforms. The platform may be, for example, a movable platform, a stationary platform, a ground structure, a water structure and a space structure. In particular, the platform may be a surface vessel, tank, personal transport, train, aircraft, submarine, automobile, powerplant, bridge, dam, house, wind power installation, production facility, building and other suitable platforms.

In some illustrative examples, the aerospace vehicle 200 takes the form of an aircraft. In an illustrative example, aerospace vehicle 200 takes the form of an unmanned aerial vehicle. In this case, deployment system 210 can be used to deploy a group of components 212 from an unmanned aerial vehicle. For example, a braking device may be deployed from an unmanned aerial vehicle to create braking to slow down this unmanned aerial vehicle.

In other illustrative examples, satellite 202 is a different type of satellite than the small satellite 204. In still other illustrative examples, control system 246 is not present in satellite 202.

In yet another illustrative example, solar modules 408 in solar cell 406 in FIG. 4 contain only one set of solar modules of the same type. In this case, all of the solar modules 408 can be configured to fold along the entire longitudinal axis 414 or diagonal axis 416 of the solar battery 406.

In one illustrative example, panel 226 may comprise two or more types of arrays. For example, half of the panel 226 may be a solar panel, and the other half of the panel 226 may be an antenna. In this example, panel 226 provides various functions of satellite 202. In yet another illustrative example, a group of panels 224 can be deployed at different points in time with respect to each other.

Although in one example, the panel 226 in the panel group 224 has been described as a flat, rectangular structure, this panel 226 may take other forms. For example, panel 226 may include a parabolic structure, an annular structure, a sphere, a triangular structure, a hexagonal structure, or some other suitable type of structure. When the panel 226 is not a rectangular structure, more or fewer connectors and interaction structures can be used to deploy this panel 226.

In FIG. 5 shows an image of a deployment system for a satellite in accordance with an illustrative embodiment. In this shown example, the deployment system 120 is shown from the side of the main body 122 of FIG. 1. Other components in satellite 102 have been removed in this image in order to show deployment system 120 in more detail.

As shown, a plurality of doors 500 are associated with a plurality of sides 106 of the housing 104 of the satellite 102. In this illustrative example, a group of components 502 is stored behind a plurality of doors 500. The plurality of doors 500 are configured to open to deploy a group of components 502 from satellite 102.

In this shown example, a door 504, a door 506, a door 508, and a door 510 are associated respectively with a side 108, a side 110, a side 112, and a side 114 of the housing 104 of the satellite 102. At least one of the door 504, door 506, door 508, and door 510 configured to open for deployment a group of components 502 from satellite 102.

For example, a command may be sent to a device in deployment system 120 to execute a door open command. This device may be a trigger release (not shown in this view). In one illustrative example, the trigger lock includes an electrically controlled heating element located adjacent to the polymer wire. After actuation, a plurality of doors 500 are opened using, for example, without limitation, torsion springs. In other illustrative examples, the release latch may include an electrically controlled pull wire fuse or some other suitable component, depending on the particular implementation.

As shown, deployment system 120 is used to deploy a group of components 502 from satellite 102 after a plurality of doors 500 have been opened. In some examples, part of deployment system 120 may be open to the environment surrounding the main body 122 of satellite 104 of housing 102. In this illustrative example, a space 512 is made between a group of components 502.

In FIG. 6 is a view of a deployment system for a satellite in accordance with an illustrative embodiment. In this example shown, satellite 102 of FIG. 1 shows a plurality of doors 500 in an open position to open a group of components 502 stored behind this plurality of doors 500.

In this illustrative example, the group of components 502 contains a group of panels 600. In this illustrative example, the group of panels 600 is a solar panel. Door 504, door 506, door 508, and door 510 were opened to open panels 602, panels 604, panels 606, and panels 608, respectively. Panel panel 600 can now be deployed from satellite 102 using deployment system 120.

In FIG. 7 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment. In this example shown, a cross-sectional view of deployment system 120 taken along lines 7-7 of FIG. 6. This illustrative example shows in more detail the components in the deployment system 120 in the main body 122 of the housing 104 of the satellite 102.

As shown, the deployment system 120 includes a drum 700, a chamber 701, a spring compartment 702, a plurality of rod compartments 704, a bearing 706, and a trigger lock 708. A spring 710 is wound around the drum 700 and stored in the spring compartment 702.

The camera 701 of the deployment system 120 accommodates a spring compartment 702, a plurality of rod compartments 704, a spring 710, and a group of flexible rods 712. In this illustrative example, a group of flexible rods 712 is an example of one implementation of a type of group of flexible elements that can be deployed from satellite 102 .

In this shown example, a group of flexible rods 712 are wound around a drum 700 and stored in a plurality of rod compartments 704. In this illustrative example, each flexible rod in the group of flexible rods 712 is stored in one of a plurality of rod compartments 704.

When the spring 710 stored in the spring compartment 702 is released, this spring 710 is unwound. The bearing 706 attached to the drum 700 causes this drum 700 to rotate as the spring 710 unwinds. This unwinding and movement of the drum 700 using the bearing 706 causes that a group of flexible rods 712 is pushed out of the plurality of rod compartments 704. A plurality of rod compartments 704 guide each of the group of flexible rods 712 to extend as necessary. In this illustrative example, the bearing 706 takes the form of a reliable bearing. In other illustrative examples, bearing 706 may take other forms to rotate drum 700.

As shown, the trigger lock 708 activates the unwinding of the spring 710 from the drum 700. The trigger lock 708 also causes the opening of multiple doors 500 using, for example, torsion springs in each door canopy.

In FIG. 8 is an illustration of a deployment system for a satellite in accordance with an illustrative embodiment. In this shown example, a view of the deployment system 120 is shown in the direction of lines 8-8 in FIG. 6.

As shown, the chamber 701 of the deployment system 120 has an opening 800. A group of flexible rods 712 in a plurality of rod compartments 704 is visible through the opening 800. The flexible rod 802 extends from the opening 800 in the deployment system 120 as the spring 710 unwinds from the drum 700. shown in FIG. 7. Flexible rod 802 is configured to deploy panel 604 as this flexible rod 802 extends.

In this illustrative example, the first end 804 of the flexible shaft 802 is attached to the cylindrical nut 806. The cylindrical nut 806 is a type of nut that contains threads perpendicular to the length of this nut. A second end of the flexible rod 802 (not shown) is attached to the drum 700.

In some illustrative examples, the cylindrical nut 806 is rotatable as the flexible shaft 802 extends from the chamber 701. For example, the cylindrical nut 806 may rotate in the direction of the arrow 808 as the flexible shaft 802 extends from the chamber 701. Each the rod in the group of flexible rods 712 extends in a similar manner from the other side of the satellite 102.

Although the first end 804 of the flexible rod 802 is shown and described as the end attached to the cylindrical nut 806, this first end 804 may be attached in the deployment system 120 in some other way. For example, the first end 804 of the flexible shaft 802 may be secured using a different type of part or combination of parts in addition to, or instead of, a cylindrical nut 806.

In FIG. 9 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment. In this example, shown is a sectional view of the deployment system 120 for satellite 102 taken along lines 9-9 of FIG. 6.

As shown, the spring 710 is shown wound around the drum 700. The first end 900 of the spring 710 is attached to the structure in the chamber 701 of the deployment system 120 at the contact point 902. The second end 904 is connected to the drum 700. The second end 904 of the spring 710 is connected to the drum 700 at the contact point 906. In this shown example, the spring 710 is unwound in the direction of arrow 908.

In FIG. 10-18 illustrate satellite panel deployment using a deployment system. In particular, FIGS. 10-18 show a deployment system 120 as the panel 604 moves from the folded configuration to the deployed configuration from side 110 of satellite 102. Although FIG. 10-18 show the movement of one panel 604, panel 602, panel 606, and panel 608 (not shown in these figures) are deployed from satellite 102 in a similar manner.

In FIG. 10 shows a folded configuration of a satellite panel in accordance with an illustrative embodiment. In this illustrative example, the door 506 may be open, and the panel 604 may be ready for deployment. The panel 604 is located in a folded configuration 1000. In this illustrative example, the panel 604 is arranged to unfold in a flat, rectangular structure.

In FIG. 11 shows an image of a panel deployable from a satellite in accordance with an illustrative embodiment. In this shown example, the first movement of the panel 604 is for folding in half as the door 506 opens.

The spring 710 (not shown in this view) this time did not start to unwind from the drum 700. When the spring 710 begins to unwind and move the drum 700, the flexible rod 802 in the group of flexible rods 712 moves in the direction of the arrow 1100.

12 is a view of a panel deployed from a satellite in accordance with an illustrative embodiment. In this shown example, the spring 710 starts to unwind from the reel 700. As the spring 710 unwinds from the reel 700, the flexible rod 802 in the group of flexible rods 712 moves out of the satellite 102, which causes additional folding of the panel 604.

In this illustrative example, the interface system 1200 is associated with a group of flexible rods 712 and a panel 604. The interface system 1200 includes connectors 1202 and interaction structures 1204. Connectors 1202 are connected to panel 604, and interaction designs 1204 are connected to flexible rod 802.

In this illustrative example, connectors 1202 for panel 604 comprise a connector 1207, connector 1208, and connector 1210. Each one of connector 1207, connector 1208, and connector 1210 is attached to a different corner of panel 604. The fourth corner of panel 604 is attached to satellite body 104 at contact point 104. (not shown in this view) on the housing 104.

In this shown example, each of the interaction structures 1204 is configured to interact with one of the connectors 1202. In particular, each of the interaction structures 1204 is received by a channel (not shown in this view) in one of the connectors 1202.

In this illustrative example, the channels in connectors 1202 have different sizes, so that each of connectors 1202 interacts with interaction structures 1204 with a different size. Thus, some interaction designs 1204 will advance through the channels in connectors 1202 before interacting with the corresponding connector.

This view shows the interaction structure 1212 in the interaction structures 1204. The interaction structure 1212 is connected to the flexible rod 802 and moves as the flexible rod 802 is advanced by the deployment system 120. The interaction structure 1212 is configured to be received by the channel in the connector 1207. The interaction structure 1212 moves through the channel in the connector 1210 and the channel in the connector 1208 before interacting with the connector 1207.

In FIG. 13 shows an image of a panel deployed from a satellite in accordance with an illustrative embodiment. In this view, the interaction structure 1212 was adopted by the channel in the connector 1207. In this illustrative example, the connector 1207 is attached to the corner 1300 of the panel 604.

The interaction of the interaction structure 1212 with the connector 1207 facilitates the unfolding of the panel 604 as necessary. In particular, the interaction of the interaction structure 1212 with the connector 1207 helps to straighten the panel 604.

In this shown example, the interaction structure 1302 in the interaction structures 1204 is shown. The interaction structure 1302 is attached to the flexible rod 802 and moves as the flexible rod 802 extends. The interaction structure 1302 is configured to be received by the channel in the connector 1208.

As shown, the interaction structure 1302 advances through the channel in the connector 1210 until it interacts with the channel in the connector 1208. In this view, the interaction structure 1302 moves from the camera 701 shown in FIG. 7.

In FIG. 14 is an image of a panel deployed from a satellite in accordance with an illustrative embodiment. This view shows the interaction structure 1400 in the interaction structures 1204. The interaction structure 1400 is connected to the flexible rod 802 and moves as this flexible rod 802 is advanced by the deployment system 120.

In this shown example, the interaction structure 1400 is configured to interact with a channel in the connector 1210. In this view, the interaction structure 1400 moves from the camera 701 shown in FIG. 7.

In FIG. 15 is an image of a panel deployed from a satellite in accordance with an illustrative embodiment. In this shown example, the panel 604 is connected to the housing 104 of the satellite 102 at a contact location 1500 on the housing 104. In this illustrative example, the angle 1502 of the panel 604 is connected to the housing 104 at the contact location 1500.

As shown, the interaction structure 1302 was adopted by the channel in the connector 1208. In this illustrative example, the connector 1208 is connected to an angle 1504 of the panel 604. The interaction of the interaction structure 1302 with the connector 1208 further straightens the panel 604.

Then, in FIG. 16 is a view of a panel deployed from a satellite in accordance with an illustrative embodiment. As more flexible rod 802 is deployed from satellite 102, panel 604 continues to unfold and straighten.

In FIG. 17 shows an image of a panel deployable from a satellite in accordance with an illustrative embodiment. In this view, the interaction structure 1400 was received by the channel in the connector 1210. In this illustrative example, the connector 1210 is attached to the corner 1700 of the panel 604. When the interaction structure 1400 is brought into contact with the connector 1210, the panel 604 is straightened.

The panel 604 is now located in the expanded configuration 1702. The forces in the direction of the arrow 1704, the arrows 1706, and the arrows 1708 continue to straighten the panel 604 and hold this panel 604 in the expanded configuration 1702.

In FIG. 18 is a sectional view of a panel in an expanded configuration and an interface system in accordance with an illustrative embodiment. In this example, shown is a sectional view of a panel 604 in an expanded configuration 1702 with an interface system 1200 along lines 18-18 of FIG. 17.

In this view, channels 1800 are made in connectors 1202. Channels 1800 comprise channel 1802, channel 1804, and channel 1806. Channel 1802 corresponds to connector 1207, channel 1804 corresponds to connector 1208, and channel 1806 corresponds to connector 1210.

As shown, channel 1802, channel 1804, and channel 1806 are of different sizes. In this illustrative example, channel 1802 represents the smallest channel, channel 1804 represents the next largest channel, and channel 1806 represents the largest channel. The dimensions of the channels 1800 are selected in such a way that only one of the interaction structures 1204 is adopted by each of the channels 1800. In this illustrative example, the interaction structure 1212 moves through channel 1806 and channel 1804 before being received by channel 1802. Similarly, interaction structure 1302 moves through channel 1806 to received by channel 1804. In this illustrative example, the interaction design 1400 is received by channel 1806.

Despite the fact that in FIG. 10-18 illustrate the deployment of a panel 604; other panels in a panel group 600 may be expanded in a similar manner. In some examples, one or more of the group of panels 600 may be deployed at the same point in time or at different points in time.

In addition, although the interface system 1200 is shown with three connectors 1202 and three interaction structures 1204 1204, a different number of connectors 1202 and corresponding interaction structures 1204 can be used. For example, two connectors, four connectors, five connectors, twelve connectors, or some other number of pairs of connectors 1202 with corresponding interaction structures 1204 may be used to deploy the panel 604.

In FIG. 19-21 show the deployment of a group of flexible rods from a satellite using a deployment system. In particular, in FIG. 19-21 are a sectional view of a satellite 102 with a deployment system 120 formed along lines 19-19 of FIG. 6 as the group of rods 612 is pushed out of the deployment system 120. The group of flexible rods 712 is shown without a group of panels 600 to illustrate the movement of the group of these flexible rods 712. Each of the group of panels 600 is expanded in a similar manner to the panel 604 shown in FIG. 10-18.

In FIG. 19 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment. In this shown example, a group of flexible rods 712 began to extend from the chamber 701 of the deployment system 120.

The group of flexible rods 712 includes a flexible rod 1900, a flexible rod 1902, and a flexible rod 1904 in addition to the flexible rod 802. In this illustrative example, the flexible rod 1900 is attached to the cylindrical nut 1906, the flexible rod 1902 is attached to the cylindrical nut 1908, and the flexible rod 1904 is attached to the cylindrical nut 1910. The cylindrical nut 1906, the cylindrical nut 1908, and the cylindrical nut 1910 are rotatable as each respective flexible shaft extends, as described with respect to the flexible the shaft 802 and the cylindrical nut 806 in FIG. 8.

In FIG. 20 is a cross-sectional view of a deployment system for a satellite in accordance with an illustrative embodiment. In this shown example, the spring 710 began to unwind from the drum 700, as described in FIG. 9, which causes the group of flexible rods 712 to slide out of the chamber 701 of the deployment system 120.

The interaction structures 2000 also unfold from the chamber 701 as the group of flexible rods 712 extends. In this illustrative example, the interaction structures 2000 are attached to the flexible rod 1900, the interaction structures 2002 are attached to the flexible rod 1902, and the interaction structures 2004 are attached to the flexible rod 1904 according to this illustrative for example.

In FIG. 21 is a cross-sectional view of a satellite deployment system in accordance with an illustrative embodiment. In this shown example, the group of flexible rods 712 is fully extended.

In FIG. 22 shows an image of a satellite with a panel of panels deployed from a satellite, in accordance with an illustrative embodiment. In this illustrated example, satellite 102 is shown after a group of panels 600 has been deployed from satellite 102 using deployment system 120, which is described in FIG. 10-21. In this view, all four panels in panel group 600 were deployed by deployment system 120. Panel group 600 can be used to create electrical energy for satellite 102.

In FIG. 23 shows an image of a satellite with a panel of panels deployed from a satellite in accordance with an illustrative embodiment. In this illustrated example, satellite 2300 is an example of another physical implementation of satellite 202, shown in block form in FIG. 2.

In this illustrative example, satellite 2300 is shown after a group of panels 2302 has been deployed from satellite 2300 using deployment system 2304. The deployment system 2304 functions in a similar manner to the deployment system 120, which is described in accordance with FIG. 1 and FIG. 5-21.

As shown, the panel group 2302 is a panel group 2303 with an antenna reflector. The panel group 2303 with an antenna reflector comprises a panel 2306, a panel 2308, a panel 2310 and a panel 2312. In this illustrative example, a panel 2306, a panel 2308, a panel 2310 and a panel 2312 are arrays of reflective antennas.

Panel 2306, panel 2308, panel 2310 and panel 2312 are part of the antenna system 2314. A group of panels 2303 with an antenna reflector in the antenna system 2314 is configured to reflect radio waves in the desired direction. In this illustrative example, the satellite body 2300 is used as part of the design of the antenna system 2314. The size of each panel from the panel group 2303 with an antenna reflector is not limited to the size of the satellite 2300.

In FIG. 24 shows an image of a satellite in accordance with an illustrative embodiment. In this illustrated example, satellite 2400 is an example of another physical implementation of satellite 202, shown in block form in FIG. 2.

As shown, satellite 2400 comprises a housing 2402 with a plurality of sides 2404. The plurality of sides 2404 includes side 2406, side 2408, side 2410, and side 2412. Satellite 2400 also includes an upper plate 2414 and a lower plate 2416.

Satellite 2400 comprises a group of solar panels 2417. In this illustrative example, a group of solar panels 2417 comprises a solar panel 2418, a solar panel 2419, a solar panel 2420 and a solar panel 2421. A solar panel 2418, a solar panel 2419, a solar panel 2420 and a solar panel 2421 are connected respectively with side 2406, side 2408, side 2410 and side 2412 of satellite 2400.

In this illustrative example, satellite 2400 also includes a deployment system 2422. Deployment system 2422 is coupled to satellite body 2400 of satellite 2400. Deployment system 2422 is configured to deploy a group of components (not shown in this view) from satellite 2400 after solar panel 2418, solar panel 2419, solar panel 2420, and solar panel

2421 were decomposed from satellite 2400.

In FIG. 25 shows an image of a satellite in accordance with an illustrative embodiment. In this shown example, the satellite 2400 of FIG. 24 is shown with a solar panel 2418, a solar panel 2419, a solar panel 2420, and a solar panel 2421 in the unfolded position. A parabolic antenna reflector 2500 is located in a folded configuration 2501 around a plurality of sides 2404 of a satellite 2400. Thus, a parabolic antenna reflector 2500 is stored in a folded configuration 2501 between a group of solar panels 2417 and a plurality of sides 2404 of a satellite 2400 until deployment.

Reflector 2500 for parabolic antennas is configured to be deployed using deployment system 2422. In this illustrative example, deployment system 2422 is located at the top of satellite 2400. The system

Deployment 2422 may include a flexible rod group, an interface system, and a deployment mechanism. The deployment mechanism extends a group of flexible rods from the deployment system 2422 to deploy the reflector 2500 for parabolic antennas, as described above.

A group of structural rods 2502 is attached to a reflector 2500 for parabolic antennas. A group of structural rods 2502 is configured to be deployed with a reflector 2500 for parabolic antennas so that this reflector 2500 for parabolic antennas has the desired shape.

In this illustrated example, satellite 2400 also includes a plate 2504. Plate 2504 is deployed from the top of satellite 2400 and used to focus electromagnetic signals to satellite 2400.

In FIG. 26 shows an image of a satellite with a group of components deployed from that satellite in accordance with an illustrative embodiment. In this shown example, the reflector 2500 for parabolic antennas is in a deployed configuration 2601.

The deployment is performed using a group of flexible rods 2600 associated with a reflector 2500 for parabolic antennas, and a deployment mechanism (not shown) configured to extend the group of flexible rods 2600 as the reflector 2500 for parabolic antennas moves from the folded configuration 2501 to expanded configuration 2601.

A group of structural rods 2502 deploys a reflector 2500 for parabolic antennas with the necessary parabolic shape. The satellite 2400 is surrounded by a reflector 2500 for parabolic antennas when this reflector 2500 for parabolic antennas is in a deployed configuration 2601.

In this illustrative example, the satellite body 2400 and the reflector 2500 for parabolic antennas form a system of 2602 antennas. The 2400 satellite hull is used as part of the design of the 2602 antenna system. Plate 2504 is used to focus electromagnetic signals onto a system of 2602 antennas.

Images of a deployment system 120, a deployment system 2304, and a deployment system 2422 and a component in a deployment system 120, a deployment system 2304, and a deployment system 2422 of FIG. 1 and FIG. 5-26 do not imply the introduction of physical or architectural restrictions in the manner in which an illustrative embodiment may be implemented. In addition to or instead of the components shown, other components may be used. Some components may be optional.

For example, although solar panels and antenna reflector panels are shown deployed from satellite 2400, deployment system 2422 may also deploy a braking device. In other illustrative examples, the solar panel group 2417 may be absent. Instead, parabolic antenna reflector 2500 can be deployed in response to opening multiple doors in deployment system 2422 to open a group of flexible rods 2600.

In still other illustrative examples, satellite 102 may be smaller than that shown in FIG. 1. For example, satellite 102 may contain only the space between the group of panels 600 shown in FIG. 6. In this example, electronic means may be embedded in the space between the panel group 600. In other words, the entire spacecraft is limited by the maximum linear dimensions of the deployment system 120 shown in FIG. 5.

In FIG. 27 shows an image of a solar panel in accordance with an illustrative embodiment. In this shown example, the solar panel 2700 is an example of the physical implementation of the solar panel 400 shown in block form in FIG. 4. The solar panel 2700 contains a solar battery 2702.

In this illustrative example, solar modules 2704 contain a first set of solar modules 2706 and a second set of solar modules 2708. The first set of solar modules 2706 contains cells that can be stacked along the longitudinal axis 2710 of the solar battery 2702, and the second set of solar modules 2708 contains cells that can be folded along the diagonal axis 2712 or the diagonal axis 2714 of the solar battery 2702. In this illustrative example, part 2716 of the solar battery 2702 contains two modules from the first set solar modules in 2706 and two modules of a second set of solar modules in 2708.

In FIG. 28 shows a solar cell section in accordance with an illustrative embodiment. In this example shown, section 2716 of the solar battery 2702 of FIG. 27.

As shown, section 2716 of solar battery 2702 contains a solar module 2800, a solar module 2802, a solar module 2804 and a solar module 2806. The solar module 2800 and solar module 2804 are part of the first set of solar modules 2706 and are configured to fold along diagonal axis 2712 The solar module 2806 and the solar module 2804 are part of a second set of solar modules 2708 and are configured to fold along a longitudinal axis, such as the longitudinal axis 2710 of FIG. 27.

In this illustrative example, the solar module 2800, the solar module 2802, the solar module 2804, and the solar module 2806 are attached to each other at the junction 2808. Connection points 2808 may include mechanical connections and / or electrical connections. In this illustrative example, solar module 2800 is connected to solar module 2806 at junction 2810 and to solar module 2802 at junction 2812. Similarly, solar module 2804 is connected to solar module 2802 at junction 2814 and to solar module 2806 at junction 2816.

In this shown example, at the junction 2808, the solar module 2800, the solar module 2802, the solar module 2804 and the solar module 2806 are connected together and then electrically connected. In an illustrative example, the connection can be made using resistance welding. The bond used to connect each solar module may contain an adhesive that contains a material selected from at least one of the following: silicon, acrylic, or some other suitable type of material.

In FIG. 29 shows an image of a solar module in accordance with an illustrative embodiment. In this shown example, the solar module 2802 of FIG. 28.

As shown, solar module 2802 contains a plurality of solar cells 2900. In this illustrative example, solar cells 2900 contain nine cells. In other illustrative examples, more or fewer cells may be made in solar module 2802, depending on the particular implementation.

In this shown example, buses 2901 surround a solar module 2802. Tires 2901 are conductive structures that surround a solar module 2802.

In this illustrative example, buses 2901 comprise a negative conductivity bus 2902, a negative conductivity bus 2904, a positive conductivity bus 2906, and a positive conductivity bus 2908. Bus 2901 is configured to allow current to flow to other solar modules.

In particular, current flows in the direction of arrow 2910 using interconnects 2912 electrically connecting the solar cells 2900 to each other. In other words, arrow 2910 shows the energy path of the solar module 2802. Mutual connections 2912 contain a conductive material selected from at least one of the following: copper, gold, silver, and some other suitable type of material.

In other illustrative examples, tires 2901 are arranged in various configurations as shown in this figure. For example, the position of the negative conductivity bus 2902 and the negative conductivity bus 2904 can be switched to the position of the positive conductivity bus 2906 and the positive conductivity bus 2908. As a result, current will flow to other solar modules in various ways.

In FIG. 30 shows an image of a solar module in accordance with an illustrative embodiment. In this example shown, the solar module 2800 of FIG. 28.

As shown, in this illustrative example, solar module 2800 contains solar cells 3000. Solar cells 3000 contain nine cells. In other illustrative examples, depending on the specific implementation, in the solar module 2800 can be performed more or less cells.

In this example shown, buses 3001 surround a solar module 2800. Buses 3001 comprise a negative conductivity bus 3002, a positive conductivity bus 3004, a positive conductivity bus 3006 and a positive conductivity bus 3008. Buses 3001 are configured to allow current to flow between the solar modules. In particular, current flows in the direction of the arrow of the ZOE using reciprocal connections 3012 electrically connecting the solar cells 3000 to each other. In other words, the ZOI arrow shows the energy path of the solar module 2800. Mutual connections 3012 comprise a conductive material selected from at least one of copper, gold, silver, and / or some other suitable type of material.

The configuration of the solar modules 2704 in the solar battery 2702 in FIG. 27-30 allows folding of the 2700 solar panel into a folded configuration. In particular, the first set of solar modules 2706 and the second set of solar modules 2708 are configured to fold along the axis so that the risk of damage to one or more of the solar cells in the solar panel 2702 can be reduced.

Instead of folding the solar modules 2704 through one or more solar cells in the modules, each of the solar modules 2704 can be folded at the junction. For example, in one illustrative example, the solar modules 2704 are stacked at the junctions between the modules, the interconnections between the solar cells and / or in the module. In addition, the position of the busbars 2901 and the busbars 3001 in the solar module 2802 and the solar module 2800, respectively, provide vertical and horizontal connections during the integration of the solar module 2802 and the solar module 2800 with other solar modules.

In FIG. 31 is an image of a control system for a satellite in accordance with an illustrative embodiment. In this shown example, the control system 3100 is an example of a physical implementation of the control system 246 shown in block form in FIG. 2. In other illustrative examples, the components in the control system 3100 can be implemented in various ways as shown in this figure, depending on the inherent functionality.

In this shown example, the control system 3100 is located in the satellite compartment 3101. For example, a compartment 3101 may be located in space 512 between the panel group 600 of the deployment system 120 shown in FIG. 6. In one illustrative example, the control system 3100 may be located in a compartment above the deployment system in or below the satellite housing.

As shown, the control system 3100 comprises a movement system 3102 and a movement system 3104. The movement system 3102 and the movement system 3104 are configured to move the platform 3106 around the pivot point 3108.

In this illustrative example, the movement system 3102 includes a drive device 3110 and a spring 3112. The movement system 3104 includes a drive device 3114 and a spring 3116. In this illustrative example, the drive device 3110 and the drive device 3114 may be drive devices driven by gear rack and pinion. In other illustrative examples, the actuator 3110 and the actuator 3114 may take other forms, depending on the particular implementation.

As shown, the drive device 3110 and the drive device 3114 move in the direction of the arrow 3118. For example, the drive device 3110 and the drive device 3114 advance and retract in the direction of the arrow 3118.

In this shown example, a spring 3112 and a spring 3116 are attached to the satellite body. In particular, the spring 3112 and the spring 3116 are attached to the plate 3120. In this illustrative example, the plate 3120 can form the upper part of the deployment mechanism housing in the deployment system.

Platform 3106 is movable in response to a movement of actuator 3110 and / or actuator 3114. In this illustrative example, platform 3106 moves around pivot point 3108. Thus, platform 3106 is an articulated platform. Deployable components (not shown) are attached to the platform 3106 so that these deployable components tilt as the platform 3106 moves.

A control system 3100 with a motion system 3102 and a motion system 3104 provides control of components deployed from a satellite during the duration of a satellite flight program. By using the moving system 3102 and the moving system 3104, the components to be deployed can be moved in the desired orientation. For example, platform 3106 can be moved so that essentially all of the solar panels attached to the satellite are tilted so that these solar panels receive solar radiation.

In FIG. 32 is a cross-sectional view of a satellite control system in accordance with an illustrative embodiment. In this shown example, a cross-sectional view of a control system 3100 with a moving system 3102 is shown along lines 32-32 of FIG. 31. In this illustrative example, the actuator 3114 moves in the direction of the arrow 3200.

In FIG. 33 is a cross-sectional view of a satellite control system in accordance with an illustrative embodiment. In this illustrative example, the actuator 3114 has been moved in the direction of arrow 3200 shown in FIG. 32, resulting in the displacement of the platform 3106. Then, the actuator 3114 moves in the direction of the arrow 3300.

On Fig shows a view in section of a control system for a satellite in accordance with an illustrative embodiment. In this illustrative example, the actuator 3114 has moved in the direction of arrow 3300 shown in FIG. 33, resulting in the displacement of platform 3106. In one illustrative example, a control system 3100 may be implemented for an orthogonal axis to provide complete control over the plane of components to be deployed.

Images of solar modules 2704 in solar battery 2702 in FIG. 27-30 and control system 3100 in FIG. 31-34 do not imply the introduction of physical or architectural restrictions in the manner in which an illustrative embodiment may be implemented. In addition to or instead of the components shown, other components may be used. Some components may be optional.

In addition, the various components shown in FIG. 1 and FIG. 5-34 may be illustrative examples of how the components shown in block form in FIG. 2-4, can be implemented in the form of physical structures. In addition, some of the components in FIG. 1 and FIGS. 5-34 may be combined with the components of FIG. 2-4 are used with the components in FIG. 2-4 or a combination thereof.

In FIG. 35 is a flowchart of a panel group deployment process in accordance with an illustrative embodiment. The process described in FIG. 35 may be implemented using the deployment system 210 in satellite 202 in FIG. 2.

The process begins with activating the deployment system (operation 3500). For example, a release latch in a deployment system may actuate a deployment system.

Then, the process advances a group of flexible elements from the spacecraft (operation 3502). In this illustrative example, a spacecraft is a satellite.

The panel group is then moved from the folded configuration to the expanded configuration using the interface system so that this group of flexible elements extends from the spacecraft (operation 3504), respectively, followed by the completion of the process.

In FIG. 36 is a flowchart of a process for deploying a group of panels from a spacecraft in accordance with an illustrative embodiment. The process described in FIG. 36 may be implemented using deployment system 210 in satellite 202 of FIG. 2.

The process begins by opening a plurality of doors covering a panel of panels located adjacent to a plurality of sides of the spacecraft body (operation 3600). A group of panels can be located in a folded configuration close to the many sides of the spacecraft.

Then the deployment mechanism is activated (operation 3602). This deployment mechanism may contain

a torsion spring wound around the drum, or an engine in the deployment system housing.

Accordingly, a group of flexible elements are unwound from the drum (operation 3604). A group of flexible elements is pushed out of the chamber in the deployment system when the deployment mechanism is actuated.

The process then enters into interaction a group of interaction structures associated with a group of flexible elements with connectors connected to a panel group to move this panel group from a folded configuration to an expanded configuration (operation 3606) and then terminating the process.

The flowcharts and block diagrams in the various illustrated embodiments illustrate the architecture, functionality, and operation of some possible implementations of devices and methods in an illustrative embodiment. Moreover, each block in the block diagrams and structural diagrams may reflect at least one of a module, segment, function, part or combination thereof, as well as from an operation or step.

In some alternative implementations of the illustrative embodiment, the function or functions indicated in the blocks may be performed in an order different from the order indicated in the figures. For example, in some cases, two blocks shown in series can be executed essentially simultaneously, or these blocks can be performed in the reverse order, depending on the inherent functionality. Also, other blocks may be added to the blocks shown in the block diagram or block diagram.

Illustrative embodiments of the present invention may be described in the context of a method 3700 for manufacturing and servicing a spacecraft, which is shown in FIG. 37, and the spacecraft 3800, which is shown in FIG. 38. In FIG. 37 shows a method for manufacturing and servicing a spacecraft in the form of a flowchart in accordance with an illustrative embodiment. During preparation, the method 3700 for the manufacture and maintenance of a spacecraft may include the development and design of 3702 spacecraft 3800 of FIG. 38 and material support 3704. During manufacture, the production of 3706 components and subunits and the integration of 3708 spacecraft systems 3800 in FIG. 38. Accordingly, the spacecraft 3800 of FIG. 38 may go through 3710 certification to be placed on 3712 service. 3710 certification may include alignment with customer requirements, industry requirements, government requirements, and some combination thereof. While being serviced by the customer, for the spacecraft 3800 in FIG. 38 make up a schedule of planned maintenance and service 3714, which may include modification, reconfiguration, restoration, and other technical support or maintenance.

Each of the processes of method 3700 for manufacturing and servicing a spacecraft can be performed or implemented by at least one of the following: a system integrator, a third party, and an operator. In these examples, the operator may represent the customer. For the purposes of this description, a system integrator may contain, without limitation, any number of manufacturers of spacecraft and subcontractors for major systems, a third party may contain, without limitation, any number of sellers, subcontractors and suppliers, and the operator may be a company, a military organization serving organization, etc.

In FIG. 38 shows an image of a spacecraft in the form of a block diagram in which an illustrative embodiment can be implemented. In this illustrative example, the spacecraft 3800 is fabricated by the method 3700 for manufacturing and servicing the spacecraft in FIG. 37. The 3800 spacecraft may comprise a hull 3802 with multiple systems 3804 and an interior 3806.

Examples of a plurality of systems 3804 comprise one or more of a propulsion system 3808, an electrical system 3810, a hydraulic system 3812, a climate system 3814, and a thermal system 3816. Although the example is shown for the aerospace industry, various illustrative embodiments may be applied to other industries such as the aviation industry, the automotive industry, the shipbuilding industry, or other suitable industries.

The device and methods implemented in this application can be applied during at least one of the steps of the method 3700 for the manufacture and maintenance of the spacecraft in FIG. 37. In particular, the deployment system 210 of FIG. 2 may be implemented during any one of the steps of method 3700 for manufacturing and servicing a spacecraft. For example, without limitation, the deployment system 210 may be installed during at least one of the following stages: production of 3706 components and subunits, integration of 3708 systems, technical support and maintenance of 3714, or some other stage of the method 3700 for the manufacture and maintenance of a spacecraft. In one illustrative example, deployment system 210 is used to deploy components while in service 3712.

In one illustrative example, components and subunits manufactured during the production of 3706 components and subunits in FIG. 37 may be industrially manufactured or manufactured in a manner similar to that of industrial production or production of components and subunits fabricated while the spacecraft 3800 is in service 3712 of FIG. 37.

As another example, a plurality of device implementations, method implementations, or a combination thereof can be used during manufacturing steps, such as manufacturing 3706 components and subunits and integrating 3708 systems in FIG. 37. A variety of device implementations, method implementations, or a combination thereof can be used while the spacecraft 3800 is in service 3712 and / or during maintenance and service 371 of FIG. 37. The use of many different illustrative implementation options can significantly accelerate the assembly of the 3800 spacecraft, reduce the cost of the 3800 spacecraft, or a combination thereof.

Thus, illustrative embodiments provide a method and apparatus for deploying a group of panels 224. In one illustrative embodiment, the device comprises a group of panels 224 located in a folded configuration 228 adjacent to a plurality of sides 208 of a spacecraft 201, a group of flexible elements 214 connected to a group panels 224, and an interface system 216 associated with a group of panels 224 and a group of flexible elements 214. The interface system 216 is configured to move a group of panels 224 of the folded configuration 228 in expanded configuration 230 when a group of flexible elements 214 is released from the spacecraft 201.

When using the illustrative implementation option, large components can be stored in the satellite and deployed from it, unlike some currently used systems. For example, when the panels are folded and stored in a satellite, the surface area of the panels may be larger than the surface area of the satellite. Compared to console solutions in which the size of the panels is limited by the size of the satellite, illustrative embodiments provide a method and system configured to deploy larger panels than before. In other words, with the exception of some currently used systems, the size of panels deployed from a satellite is not limited to the size of this satellite.

Illustrative embodiments also provide the ability to fold large panels in a more efficient manner while reducing the risk of damage to the components in these panels. For example, when using a solar panel 400 with a first set of solar modules 410 and a second set of solar modules 412, this solar panel 400 can be folded and stored in a satellite in a manner that reduces the risk of damage to one of the solar modules.

In addition, a deployment system 210 with a group of flexible members 214 and an interface system 216 provide a neutral path for deploying a group of panels 224 as needed. As a result, the size of the group of panels 224 stored in the deployment system 210 can be increased in comparison with some currently used systems that have panels deployed on the satellite side. Large panels enable more efficient satellite operations. The group of flexible elements 214 and the interface system 216 enables the necessary extension of the group of panels 224 so that each of the group of panels 224 is expanded.

In other examples, when it is necessary for the panel group 224 to have an alternative configuration, the deployment system 210 may be located at various positions in the satellite to deploy the panel group 224 as needed. For example, when deploying a parabolic antenna, the deployment system 210 is located on the top of the satellite and deploys components stacked around the satellite.

In addition, deployment system 210 provides a compact, lightweight deployment system that does not increase satellite weight more than necessary. The deployment system 210 can be implemented in various sizes of satellites without significantly increasing the weight of the satellite and, in turn, increasing the cost of launching the satellite.

Using the control system 246, a group of panels 224 deployed from satellite 202 can be maneuvered to track the sun, track objects of interest, increase drag during satellite operations 202, or a combination thereof. As a result, a group of panels 224 can receive the required amount of solar radiation used to create energy to perform satellite operations 202.

In addition, the present invention contains embodiments according to the following points:

1. A device comprising:

a group of panels in a folded configuration, located close to the many sides of the spacecraft,

a group of flexible elements connected to the panel group, and an interface system associated with the panel group and the group of flexible elements, the interface system configured to move the panel group from the folded configuration to the expanded configuration when the group of flexible elements is pulled out of the spacecraft.

2. The device according to paragraph 1, further comprising: a deployment mechanism configured to

extending the group of flexible elements from the inside of the spacecraft to the outside of the spacecraft so that the panel group moves from the folded configuration to the deployed configuration.

3. The device according to claim 2, wherein the deployment mechanism comprises a torsion spring and / or engine in the spacecraft body.

4. The device according to paragraph 1, further comprising: a drum in the spacecraft body, and

a group of flexible elements is arranged to be wound onto a drum.

5. The device according to claim 1, wherein the interface system comprises:

connectors associated with a panel group and a group of flexible elements, wherein each connector in these connectors contains a channel and is configured to attach to a panel in a panel group, and

a group of interaction structures associated with a group of flexible elements, and the interaction structure in the group of interaction structures is configured to interact with the corresponding connector in the connectors to move the panel in the panel group from the folded configuration to the expanded configuration.

6. The device according to claim 1, wherein the group of panels is connected to the spacecraft through a plurality of contact points.

7. The device according to claim 1, wherein the panel in the panel group is selected from one of the following: a solar battery, an antenna, and a brake device.

8. The device according to paragraph 1, further comprising: a control system configured to

control panel panels in space.

9. The apparatus of claim 1, wherein the interface system keeps the panel group substantially flat when the panel group is in an expanded configuration.

10. The device according to paragraph 1, in which the panel group contains a solar battery containing:

the first set of solar modules made with the possibility of folding along the longitudinal axis, and

the second set of solar modules made with the possibility of folding along the diagonal axis to form a folded configuration of the solar battery.

11. The device according to paragraph 1, further comprising: a spacecraft body, and a group

panels is located close to the many sides of the housing when this group of panels is located in a folded configuration, and

many doors made with the possibility of closing a group of panels located close to the many sides of the body in a folded configuration, and with the possibility of opening so that a group of panels is moved from the folded configuration to the expanded configuration.

12. The device according to claim 1, in which the group of flexible elements is formed from structures with a sectional shape selected from at least one of the following: round shape, triangular shape, rectangular shape, irregular shape and hexagonal shape.

13. A method for deploying a panel group, according to which:

pushing a group of flexible elements from a spacecraft and

move the panel group from the folded configuration to the expanded configuration using the interface system so that the group of flexible elements extends from the spacecraft.

14. The method of claim 13, further comprising: positioning the panel group in folded

configuration close to the many sides of the spacecraft.

15. The method according to p. 13, further comprising: winding a group of flexible elements around the drum in

deployment mechanism enclosure

bringing the deployment mechanism into action and

unwinding a group of flexible elements from the drum when the deployment mechanism is actuated.

16. The method according to p. 13, according to which the connectors are connected to a group of panels, each connector in the connectors has a channel, and the method further includes:

Bringing a group of interaction designs associated with a group of flexible elements to interact with connectors to move a group of panels from a folded configuration to an expanded configuration.

17. The method according to p. 13, according to which the panel group comprises a solar battery containing a first set of solar modules and a second set of solar modules, said method further comprising:

folding the first set of solar modules along the longitudinal axis and

folding the second set of solar modules along the diagonal axis to form a folded configuration of the solar battery.

18. An antenna system comprising:

a group of flexible rods associated with a group of panels with an antenna reflector, which is located in a folded configuration close to the many sides of the satellite,

an interface system associated with a group of panels with an antenna reflector and a group of flexible rods, the interface system configured to move a group of panels with an antenna reflector from a folded configuration to an expanded configuration, and

a deployment mechanism configured to extend a group of flexible rods as a group of panels with an antenna reflector moves from a folded configuration to an expanded configuration.

19. The antenna system of claim 18, wherein the satellite comprises a body configured to be used as part of the antenna system design.

20. The antenna system according to paragraph 19, further comprising: a plate made with the possibility of deployment from the top

parts of the satellite and with the possibility of using to focus electromagnetic signals on the antenna system.

21. The antenna system according to paragraph 18, in which each group of panels with a reflector antenna contains a matrix of reflective antennas, configured to reflect radio waves in the desired direction.

22. A control system, comprising: a platform located in the satellite housing, and

a moving system associated with the platform and configured to change the orientation of the platform for maneuvering a group of components in space after the group of components is deployed from the satellite.

23. The control system according to paragraph 22, in which the movement system is configured to tilt a group of components by changing the orientation of the platform so that the group of components has the desired orientation.

24. The control system according to paragraph 23, in which the movement system comprises a drive device configured to extend and retract so that the platform moves around the pivot point.

25. The control system according to paragraph 24, in which the group of components contains a group of panels attached to the platform, and the moving system is configured to tilt the group of panels to receive the required number of electromagnetic signals.

26. The control system according to paragraph 22, in which the platform is located in the satellite body.

27. The control system according to paragraph 22, further comprising:

a spring configured to reduce the rotational moment required by the movement system to change the orientation of the platform.

28. An antenna system comprising:

a reflector for parabolic antennas located in a folded configuration close to the many sides of the satellite,

a group of flexible rods associated with a reflector for parabolic antennas,

an interface system associated with a reflector for parabolic antennas and a group of flexible rods, the interface system configured to move the reflector for parabolic antennas from a folded configuration to an expanded configuration, and

a deployment mechanism configured to extend a group of flexible rods as the parabolic antenna reflector moves from the folded configuration to the deployed configuration.

29. The antenna system of claim 28, wherein the satellite is surrounded by a parabolic antenna reflector when the parabolic antenna reflector is in an expanded configuration.

30. The antenna system of claim 28, further comprising: a group of solar panels located adjacent to

a plurality of satellite sides, the reflector for parabolic antennas being stored in a folded configuration between a group of solar panels and a plurality of satellite sides.

31. The antenna system according to paragraph 30, in which the group of solar panels is arranged to unfold before moving the reflector for parabolic antennas in a deployed configuration.

32. The antenna system of paragraph 28, further comprising: a group of structural rods attached to

reflector for parabolic antennas and configured to deploy this reflector for parabolic antennas with the necessary shape.

33. A solar panel containing:

a solar battery containing the first set of solar modules made with the possibility of folding along the longitudinal axis, and the second set of solar modules made with the possibility of folding along the diagonal axis to form a folded configuration of the solar battery.

34. The solar panel according to paragraph 33, wherein the solar panel is connected to a satellite and configured to be deployed from this satellite using a deployment system.

35. The solar panel according to paragraph 33, wherein the first set of solar modules is folded along the longitudinal axis and the second set of solar modules is folded along the diagonal axis so that the risk of damage to one or more solar cells in the solar battery is reduced.

A description of various illustrative embodiments has been provided for purposes of illustration and explanation, and should not be considered exhaustive or limited to the disclosed embodiments. Many modifications and changes will be apparent to those skilled in the art. In addition, various illustrative embodiments may provide various features as compared to other suitable embodiments. The selected implementation option or the selected implementation options are selected and described to best explain the principles of the implementation options, practical application, as well as to provide an opportunity for specialists in the field of technology to understand the essence of the various implementation options with various modifications that are suitable for the specific intended use.

Claims (30)

1. A device for deploying a group of panels (224), comprising: a group of panels (224) located in a folded configuration (228) close to the many sides (208) of a spacecraft (201), a group of flexible elements (214) connected to a group of panels (224), and an interface system (216) associated with a group of panels (224) and a group of flexible elements (214), the interface system (216) configured to move a group of panels (224) from a folded configuration (228) to an expanded configuration (230) when the group of flexible elements (214) is advanced from the spacecraft (201), and the interface system (216) further comprises: connectors (300) associated with the panel group (224) and the flexible element group (214), each of the connectors (300) comprising a channel (308) and configured to attach to the panel (226) from the panel group (224), and a group of interaction structures (302) associated with a group of flexible elements (214), and the interaction structure (310) from the group of interaction structures (302) is configured to interact with the corresponding connector from these connectors (300) by adopting the interaction structure (310) in the channel (308) of the indicated connector for moving the panel (226) from the group of panels (224) from the folded configuration (228) to the expanded configuration (230). 2. The device according to claim 1, further comprising: a deployment mechanism (218) configured to extend a group of flexible elements (214) from the inside (220) of the spacecraft (201) to the outside (222) of the spacecraft (201) so that the panel group (224) makes moving from the folded configuration (228) to the deployed configuration (230), wherein the deployment mechanism (218) comprises a torsion spring (234) and / or an engine (235) in the housing (206) of the spacecraft (201). 3. The device according to any one of paragraphs. 1-2, additionally containing: a drum (236) in the casing (206) of the spacecraft (201), the group of flexible elements (214) being configured to wind it onto this drum (236). 4. The device according to any one of paragraphs. 1-2, in which a group of panels (224) is connected to a spacecraft (201) via a plurality of contact points (238), wherein the panel (226) in the group of panels (224) is selected as one of the following: a solar battery (406), an antenna (402) and a brake device (404). 5. The device according to any one of paragraphs. 1-2, additionally containing: a control system (246) configured to control a group of panels (224) in space. 6. The device according to any one of paragraphs. 1-2, in which the interface system (216) keeps the panel group (224) substantially flat when the panel group (224) is located in the expanded configuration (230). 7. The device according to any one of paragraphs. 1-2, in which the panel group (224) contains a solar battery (406), comprising: the first set of solar modules (410), made with the possibility of folding along the longitudinal axis (414), and the second set of solar modules (412), made with the possibility of folding along the diagonal axis (416) to form a folded configuration (228) of the solar battery (406). 8. The device according to any one of paragraphs. 1-2, additionally containing: a casing (206) of a spacecraft (201), the panel group (224) being adjacent to a plurality of sides (208) of the casing (206) when this panel group (224) is in a folded configuration (228), and a plurality of doors (242) configured to close a group of panels (224) located close to a plurality of sides (208) of the housing (206) in a folded configuration (228), and with the possibility of opening in such a way that the group of panels (224) is moved from folded configuration (228) to expanded configuration (230). 9. The device according to any one of paragraphs. 1-2, in which the group of flexible elements (214) is formed from structures with a sectional shape selected at least one of the following: round shape, triangular shape, rectangular shape, irregular shape and hexagonal shape. 10. The method of deployment of a group of panels (224), according to which: pushing (3502) from the spacecraft (201) a group of flexible elements (214) connected to a group of panels (224) located in a folded configuration (228) close to the many sides (208) of the spacecraft (201), and move (3504) the panel group (224) from the folded configuration (228) to the expanded configuration (230) using an interface system (216) associated with the panel group (224) and the flexible element group (214) when the flexible element group (214) ) is advanced from the spacecraft (201); moreover, the interface system (216) further comprises connectors (300) associated with a panel group (224) and a group of flexible elements (214), and a group of interaction structures (302) associated with a group of flexible elements (214); moreover, the interaction structure (310) from the group of interaction structures (302) interacts with the corresponding connector from these connectors (300), each of which contains a channel (308) and is configured to attach to the panel (226) from the group of panels, due to the adoption of the design (310) interaction in the channel (308) of the indicated connector to move the panel group (224) from the folded configuration (228) to the expanded configuration (230). 11. The method according to p. 10, further comprising: winding a group of flexible elements (214) around the drum (236) in the housing (206) of the deployment mechanism (218), bringing the deployment mechanism (218) into operation and unwinding a group of flexible elements (214) from the drum (236) when the deployment mechanism (218) is activated. 12. The method according to any one of paragraphs. 10-11, according to which the panel group (224) comprises a solar battery (406) comprising a first set of solar modules (410) and a second set of solar modules (412), said method further comprising: folding the first set of solar modules (410) along the longitudinal axis (414) and folding the second set of solar modules (412) along the diagonal axis (416) to form a folded configuration (228) of the solar battery (406).

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