device to move or remove artificial satellites

专利摘要:
DEVICE FOR MOVING OR REMOVING ARTIFICIAL SATELLITES The present invention relates to a device (10, 20, 40, 50) for attaching to a space satellite (20 ', 20 ") before the latter is launched for the purpose of changing orbit said satellite and / or return it to Earth.The device comprises: means for controlling the device (10, 20, 40, 50); propulsion means operatively connected to the control means; means for receiving control signals operatively connected control means; means for electrically driving the device (10, 20, 40, 50); means (310, 320, 330, 340 ', 340 ", 350, 360) for mechanically coupling the device (10, 20, 40 , 50) to said satellite (20 ', 20 ") before the last one is launched. The propulsion means are enabled by the control means in the reception of control signals to remove the satellite from orbit (20', 20") and transfer the same to a given orbit.
公开号:BR112014001459B1
申请号:R112014001459-0
申请日:2012-07-18
公开日:2021-01-12
发明作者:Luca Rossettini；Giuseppe Jussef Tussiwand；Renato Panesi；Thomas Panozzo
申请人:D-Orbit S.R.L.；
IPC主号:

专利说明:

[0001] [0001] The present invention relates to a device for removing artificial satellites from space or for moving artificial satellites to a different orbit in space. In particular, the present invention relates to a device for controlled and safe removal of satellites, namely for the orbit withdrawal of artificial satellites or space vehicles, in which the term "withdraw from orbit" means a controlled and rapid return of said artificial satellites, or space vehicles, from low orbits in the Earth's atmosphere to pre-established places on Earth. According to another aspect, the present invention relates to a device capable of effecting the orbit change of artificial satellites, or space vehicles, which is a controlled and rapid transfer of said artificial satellites, or space vehicles, from their orbits. for a parking orbit.
[0002] [0002] More particularly, the present invention relates to a device for moving artificial satellites, or space vehicles, and / or removing them from the orbit of the mission at the end of its useful life or when they become defective.
[0003] [0003] In a further aspect, the present invention relates to a device capable of serving as a reserve propulsion system for changing the orbital position of artificial satellites or space vehicles. In another aspect, the present invention relates to a method for removing a space satellite from its orbit in space or moving said satellite to a different orbit in space. More particularly, the present invention relates to a method for operating the orbit withdrawal or change of orbit of a space satellite, or space vehicle, by an independent movement / removal device provided on said space satellite or space vehicle.
[0004] [0004] The terms "artificial satellite", "space satellite" or "space vehicle" refer, for the purposes of the present invention, to artificial satellites or vehicles capable of orbiting or moving in space from low orbits in the atmosphere of the Earth.
[0005] [0005] The term "mission orbit" refers, for the purposes of the present invention, to the orbit designated for the satellite or space vehicle for the operations required during its stay in space.
[0006] [0006] The term "low orbit" refers, for the purposes of the present invention, to a low Earth orbit (LEO), that is, a circular orbit at an altitude between the atmosphere and the 200-inch Van Allen Belt 2,000 kilometers away from the Earth's surface.
[0007] [0007] The term "high orbit" refers, for the purposes of the present invention, to an average Earth orbit (MEO), or a geosynchronous orbit (GSO) or a geostationary Earth orbit (GEO). A MEO is an orbit that lies between a LEO and a GEO, approximately 10,000 km away from the Earth's surface. A GSO is an orbit around the Earth with an orbital period that is the same as the period of the Earth's sidereal rotation. A GEO is a particular type of GSO and is an equatorial circular orbit at a distance of approximately 36,000 kilometers above the equator.
[0008] [0008] The term "parking orbit" refers, for the purposes of the present invention, to an orbit in which the artificial satellite or space vehicle can remain harmless or occupy space that is useful for other artificial satellites or space vehicles.
[0009] [0009] The first artificial satellite was launched in 1957. Since then, more than 6,000 satellites have been launched, only 800 of which are currently in use. Most satellites that are no longer operational continue to orbit around the Earth and it is impossible to control their trajectories.
[0010] [0010] With the increasing number of satellites being launched, there is a consequent reduction in the space available to position new satellites in orbit. In addition, non-operational and / or uncontrolled satellites have a high probability of colliding with another and exploding. This poses a number of problems because approximately 50% of the trackable objects in orbit are generated by explosions or collisions in space. As a consequence, there are approximately half a million pieces of space debris currently moving at more than 30,000 km per hour around the Earth, polluting the orbital spaces that are most suitable from a scientific, technical and commercial point of view. Each new satellite launched into orbit not only increases the number of objects in space, but also the amount of uncontrolled debris resulting from the continuous collisions and explosions to which abandoned debris is subject.
[0011] [0011] For the low orbit class (LEO), the increasing amount of debris constitutes a threat and a source of pollution that is far from negligible. This orbital region is relatively small and almost saturated, with a considerable risk of collisions between debris and space objects that are still in use. There is a risk of severe damage to artificial satellites, or even their total destruction, with the consequent failure of the corresponding space missions. The uncontrolled growth of debris in this space can lead to what is known as Kessler syndrome, that is, when a certain density of debris in orbit is reached, there may be a chain of collisions that would prevent any spatial activity or use of satellites by generations. Additionally, space objects left in the lower orbits return to Earth's atmosphere and fall on the planet's surface months or years after any available orbital control systems on board the satellite have run out of fuel, or when a satellite may be out of range. control due to a defect, which means that their re-entries are not controlled and therefore put any object or human being at risk. In fact, if this debris is not completely destroyed when it moves through the atmosphere - as may be the case, depending on the materials used in the construction of the spacecraft and the dynamics of its reentry - it can fall at great speed to Earth, becoming a danger to humans, buildings or infrastructure (for example, gas tanks, electric power lines, nuclear reactors, etc.). Even disregarding any direct impact on objects and humans, uncontrolled debris re-entry can be highly dangerous. In fact, some satellites may have radioactive or highly toxic material on board (such as fuels containing hydrazine), which could be dispersed in the atmosphere over densely populated areas. Currently, the number of objects that fall on Earth from space already reaches one per day and no one can predict when or where the impact will occur. In general, the speed of impact varies from approximately 30 km / hour for light objects to more than 300 km / hour for heavier items,
[0012] [0012] Space objects placed in medium or high orbits (MEO and GEO) are not decelerated by the highest layers of the Earth's atmosphere (the exosphere), so they do not fall towards the Earth's surface. Consequently, artificial satellites occupy commercially useful orbital segments for a very long time. At the end of their useful lives, which are typically 15 years old, they become useless and are left uncontrolled, and this prevents the positioning of any other spatial object in its vicinity. Given the importance of the orbits in question and the limited available space, satellites have to be relocated to a graveyard or parking orbit before they reach the end of their useful lives. This is typically a high orbit of no use for the purposes of space missions. Said repositioning of the satellite is carried out using the same propulsion systems as for orbital maneuvers, if they are available on board. A portion of the fuel stored in the tanks can be reserved for this purpose, consequently limiting the satellite's useful life and sacrificing a portion of the fuel loaded on board to allow the parking maneuvers to be completed. Changing an artificial telecommunications satellite's orbit involves its commercial operations for approximately 6 months. For a satellite with an initial mass of approximately 1,000 kilograms and a seven-year mission, this means an economic loss of approximately $ 10,000,000. In the event of any problems, such as a malfunction of the used thrusters, a lack of fuel, the lack of an adequate propulsion system, or a mechanical failure, the satellite remains in its position practically forever, preventing its replacement by new satellites with more advanced technology. Several simulations have estimated that each artificial satellite in the geostationary orbit passes within approximately 50 meters of another space object every year, with the high risk of related impact.
[0013] [0013] The "International Convention on Liability for Damage Caused by Space Objects" establishes that whoever throws an object into space is responsible for any damage caused by the object, both in space and on Earth. The IADC (Inter-Agency Space Debris Coordination Committee) states that a generic space vehicle must be maneuvered at the end of its work mission to prevent it from interfering with other space systems in orbit and, in the case of artificial satellites in orbit low, its re-entry into the atmosphere must be guaranteed within 25 years of the end of its useful life.
[0014] [0014] In general, the cost incurred by an artificial satellite in a GEO orbit to meet IADC requirements represents approximately 10% to 15% of the total cost of the satellite. Therefore space debris poses a growing threat to scientific and commercial developments in orbit. Consequently, it should be useful to produce a device capable of removing from orbit or moving (changing orbit) safely artificial satellites and orbiting space objects in general that have reached the end of their useful lives, in a reliable and controlled manner.
[0015] [0015] There are other known devices for removing space objects from orbit or moving them to parking orbits. They usually comprise passive devices, that is, they have no control over the time, trajectory and final destination of the object to be moved.
[0016] [0016] The document US 6,655,637 describes a device to launch into space that can grab objects already in orbit and make them go out of orbit towards Earth through an inflatable reentry module. This is typically a spherical object of considerable dimensions to ensure that it is attracted to the Earth as a result of its interaction with the residual layers of atmosphere. This device has the disadvantage, however, of having to carry the object to be removed into the atmosphere without being able to guarantee its control during the re-entry phase or any substantial reduction in the related entry times. Consequently, it cannot even guarantee that the orbit found during the "fall" phase will not put it at risk of colliding with other satellites. In addition, this device is unable to move objects located in high orbits to a parking orbit. An additional drawback is that recovery operations require delicate orbital maneuvers and dedicated launches for each object to be removed, which makes the recovery process expensive and risky.
[0017] [0017] US 5.120.008 describes a device that exploits solar radiation that passes through lenses to dissolve space debris. This device has the disadvantage that it is useful only for small items of debris that would spontaneously re-enter the atmosphere in any way within a relatively short time. In addition, the device must be equipped with a propellant that can be rekindled to reach the various pieces of debris in orbit. Another drawback lies in the need to organize a dedicated launch to reach the orbit (s) where the identified debris has to be dissolved, here again making the cleaning process expensive and risky.
[0018] [0018] US 4,991,799 describes an apparatus consisting of a spacecraft that rotates around its own axis with a plurality of panels that have a wide surface area against which space debris must impact and remain trapped when the sweeping apparatus is moving. Like those previously mentioned, this device has the disadvantage of requiring a propellant that can be rekindled to reach the various pieces of debris in orbit. Another disadvantage is that it is only useful for small items, which would spontaneously re-enter the atmosphere in any way within a relatively short time. An additional disadvantage is the need to prepare a dedicated launch to reach the orbit (s) from which the identified debris has to be eliminated, here again making the cleaning process expensive and dangerous.
[0019] [0019] US 5,082,211 describes a "toothed" device for removing space debris from orbit. This device consists of a long coiled cable that can be unwound at the beginning of the phase out of orbit. The method can be applied only to debris in low orbits and uses the residual atmosphere at these altitudes to gradually slow down the debris and re-enter the atmosphere. The main disadvantage of this solution, however, is that it is impossible to guarantee any control over the trajectory or the impact site on Earth. In addition, the operations to unwind and stabilize the cable are highly complex and expensive.
[0020] [0020] US 7,093 800 describes a method for manipulating a satellite at the end of its life using a portion of the fuel loaded on board for purposes of orbit withdrawal. The main disadvantage of this method resides in the need for the satellite to still be in good working order when it comes time to remove it from orbit. The same is also possible only for removing satellites equipped with propellants that can be restarted by burning liquid fuels. Another disadvantage arises from the need to have an additional fuel mass and volume inside the fuel tanks to use in the orbit withdrawal maneuver, thereby reducing the satellite's useful life. The safety of this system necessarily depends on the safety of the liquid fuel engines and the control system.
[0021] [0021] US 6,629,673 describes a solid hybrid propellant propellant used to move a transfer vehicle with people on board. This device can be rewired and is controlled by the spaceship to which it is connected. This thruster allows an emergency displacement to be carried out, but thruster control requires proper functionality of the vehicle to be moved and its control system.
[0022] [0022] CN 201165916U describes a method for using a group of four solid propellant engines for orbital transfers. The main disadvantage of these devices lies in their lack of autonomy since they have to be controlled by the space vehicle with which they are associated, so their safety depends on the latter.
[0023] [0023] US 6,024,328 describes a method for controlling a satellite by means of a liquid fuel propellant. This system requires tanks for the fuel, a perfectly sealed tank for the pressurized gas used to drive the fuel and oxidizer inside the combustion chamber, and valves and other components. The amount of components in the system contributes to reducing its safety in general and additionally increases the mass that has to be launched, and consequently the costs involved. Another disadvantage is the space occupied within the spacecraft that has to be removed from orbit.
[0024] [0024] Other known devices and methods are described in non-patent literature for removing space objects from orbit.
[0025] [0025] The document by Janosky R. et al., "End-Of-Life De-Orbiting Strategies for Satellites", DGLR Jahrbuch 2002, 1-10, Deutscher Luft und Raumfahrtkongress, Stuttgart, 23-26 Dept. 2002, describes a method for removing satellites from orbit. It provides an overview of different types of propulsion means suitable for removing satellites from orbit. In particular, it specifies that the propulsion systems most suitable for operating the orbit withdrawal are means of solid propulsion propulsion. This document, however, does not describe how the system works or should be configured. In addition, the method is described only to operate the orbit withdrawal of satellites working in a LEO orbit. The means of propulsion are described as a part of a satellite or as the same means of propulsion of the satellite itself, thus there is no description of a device for removing from autonomous orbit with respect to the satellite.
[0026] [0026] Schonenborg R.'s document, "Solid Propellant De-orbiting for Constellation Satellites", Proceedings of the fourth International Spacecraft Propulsion Conference (ESA SP-55), 2-9 June 2004, Chia Laguna (Cagliari) , describes the use of an orbit withdrawal system equipped with solid propulsion means. The document describes the use and positioning of solid propulsion means on Iridium satellites. The most relevant disadvantage of this system refers to the connection with the satellite, which does not allow a control of the process of withdrawal from orbit when the propellant is lit.
[0027] [0027] Therefore, although the known devices adopt solutions that comprise propulsion systems that can be used for orbit withdrawal purposes, they are not able to guarantee an economical and really safe operation or are not able to control the operations withdrawal from orbit. orbit / orbit change when they are started. In fact, the devices and methods described do not really involve an autonomous device for orbit withdrawal / orbit change and, if so, do not reveal the interaction between said device and the satellite to change orbit / withdraw from orbit. In addition, these devices require launches for the specific purpose of identifying objects to be removed from orbit, which increases the cost involved in these operations. In addition, they comprise a large number of components with complex interactions between them, which means that the real security of the system is drastically reduced.
[0028] [0028] Therefore, there is a need for a device and a method for the removal of objects orbiting in space that is autonomous with respect to said objects, but coupled to them.
[0029] [0029] There is also a need for a device and a method capable of ensuring that the operations involved in said removal are economical.
[0030] [0030] It would be desirable that said device and method were able to preserve a high degree of security for the entire duration of the mission of the space object to be removed.
[0031] [0031] It would also be desirable for said device and method to be able to operate independently, without the need to interact with the space object to be removed.
[0032] [0032] It would also be desirable that said device and method were able to guarantee the removal of the object from its orbit in a safe and controlled manner.
[0033] [0033] Within the context of the above technical target, an objective of the present invention is to provide an autonomous device capable of minimizing or eliminating the inconveniences that arise from objects orbiting in space, such as artificial satellites, which are no longer in use. Another objective of the present invention is to provide a standalone device capable of guaranteeing a high degree of its entire life cycle to thereby ensure its proper operation at any time,
[0034] [0034] Another objective, which is part of the target mentioned above, is to provide an autonomous device that comprises a minimum number of parts, and consequently implying limited production and assembly costs. In addition, another objective of the present invention is to provide a device capable of completing the procedure for removing the object from its orbit in a safe and controlled manner.
[0035] [0035] The term "autonomous" refers, for the purposes of the present invention, to a device that is coupled to the satellite to be moved or removed, but that works autonomously without using any satellite system for its orbit withdrawal operations / orbit change.
[0036] [0036] In addition, another objective of the present invention relates to a method to remove a space satellite from its orbit in space or to move said satellite to a different orbit in space remotely by an independent movement / removal device coupled to the satellite itself.
[0037] [0037] The above objectives are achieved by a device, according to claim 1. In particular, the above objectives are achieved by a device to be coupled to a space satellite before launch to remove said satellite from space or move the said satellite to a different orbit in space by means of remote control or by said device, characterized by the fact that said device is independent with respect to said satellite and said means of remote control, in which said device comprises:
[0038] [0038] - means of control on board said device;
[0039] [0039] - means to receive control signals from said remote control means or means to emit control signals to activate the movement / removal sequence, operationally connected to the control means on board;
[0040] [0040] - propulsion means operatively connected to the control means on board, wherein said means of propulsion are activated by said control means on board as a result of receiving said control signals to remove said satellite from space for a predefined area on Earth or move said satellite to a different predefined orbit in space;
[0041] [0041] - means of supplying electrical energy for said device, to make said device independent of said satellite;
[0042] [0042] - means for pre-launching mechanical coupling of said device to said satellite;
[0043] [0043] - means to mitigate the misalignment of the impulse vector, operationally connected to the said means of control on board.
[0044] [0044] The term "remote control means" refers, for the purposes of the present invention, to software and / or hardware means capable of sending control signals to the device. Said software and hardware must be part of another spacecraft or a space station orbiting, or a control station located on Earth.
[0045] [0045] With the characteristics described above, the device according to the invention can effect the withdrawal of orbit or change the orbit of the satellite with which it is independently associated with said satellite and its state of life. Said orbit withdrawal or orbit change can also be controlled remotely by means of remote control transmitting from another spacecraft or space station orbiting, or from a control station located on Earth. The control means on board the device carry out the procedure of withdrawing orbit or changing the orbit in a few direct steps.
[0046] [0046] The simultaneous displacement of the device and satellite allows the device to properly discard the satellite with which it is coupled at the end of the latter's mission. In particular, the said disposal process is achieved by moving the satellite from its operational orbit, or by removing the satellite from orbit towards Earth. When the necessary control signals are received, the device according to the present invention is able to clear the space of the orbit previously occupied by the satellite with which it is associated, and thus prevent it from interfering with other satellites or spacecraft. orbiting in the neighborhood. In particular, the re-entry procedure is carried out in a safe and controlled manner, avoiding any damage due to collision and impact with people or property on Earth, or with other space objects orbiting nearby, identifying a predefined safe area on Earth in which the satellite will collide.
[0047] [0047] The means to mitigate the misalignment of the impulse vector allow the control of the trajectory of orbit withdrawal / orbit change when the propellant is turned on, without the use of systems that belong to the satellite. In particular, by mitigating misalignment, the device can be turned on only once without reinflaming the propellant to control the trajectory of the device and the satellite.
[0048] [0048] Preferably, the device comprises means to detect and / or control the behavior of the satellite, operationally connected to the means to mitigate misalignment. More preferably, the means for mitigating misalignment comprises means for active and / or passive control of the drive vector which effect the alignment of said drive vector.
[0049] [0049] In this way, the device controls the behavior and / or the impulse vector and / or the misalignment of the satellite during the propulsion of the device, allowing the satellite to move to the predefined trajectory on Earth or to the new orbit as calculated . In particular, said control and alignment allow to minimize the fuel needed.
[0050] [0050] Preferably, the means for active and / or passive control of the impulse vector make an adjustable and / or mobile connection between the satellite and the device to effect the alignment of the impulse vector. In this way, the control of misalignment does not involve the means of propulsion and the alignment of the impulse vector can be done in a simple and economical way.
[0051] [0051] Preferably, the means to detect and / or control the behavior of the satellite are operationally connected to the control means on board and carry out the directional control of the device and the satellite when the propulsion means are activated. In this way, the device has active control over the trajectory allowing the device to complete orbit withdrawal / orbit change operations safely and in a few steps.
[0052] [0052] Preferably, the device comprises safe ignition means operatively connected to the control means on board to receive the ignition control signal and to operate the ignition of the propulsion means. In this way, ignition of the device is possible only if a safety signal is received by the device, thus preventing spontaneous ignition or unauthorized intrusion into the device's systems.
[0053] [0053] Preferably, the device comprises telecommunications means, operationally connected to the control means on board, to receive and / or send data from the device to the remote control means; telecommunication means comprising at least means for receiving control signals and / or at least means for transmitting data.
[0054] [0054] Preferably, the onboard control means comprise means for emitting pre-programmed control signals and / or means for calculating said control signals to be sent to the means for receiving control signals. In this way, the device can be placed in an autonomous operation by starting the ignition sequence by itself.
[0055] [0055] Preferably, the device comprises first sensing means, operationally connected to the control means on board, for the detection of other objects within a defined security zone around the device and / or the satellite. More preferably, the device comprises second sensing means, operationally connected to the onboard control means, for detecting impact damage to the device and / or the satellite. More preferably, the device comprises third sensor means, operationally connected to the control means on board, for the detection of satellite failures.
[0056] [0056] More preferably, the device comprises four sensing means, operationally connected to the control means on board, to detect the orientation over time of the device and / or to detect the orbit of the device and / or to detect the position in space the device.
[0057] [0057] Preferably, the device comprises means of communication with the satellite, operationally connected to the control means on board; the means of communication report faults on the satellite and / or communicate control signals to the device.
[0058] [0058] More preferably, the device comprises means for monitoring the state of the device itself, operationally connected to the control means on board, and for sending said state to the remote control means. In this way, the device is completed with means and sensors capable of detecting any anomalies in the operation of the device itself or of the artificial satellite with which it is associated, which advantageously allows the device to discard malfunctioning satellites for which re-entry it would no longer be possible.
[0059] [0059] In the case of a satellite severely damaged during its lifetime, the device according to the present invention allows the damaged satellite to be moved into a parking orbit, or induced to re-enter the Earth's atmosphere, thereby greatly reducing the risk of collision with other satellites.
[0060] [0060] Proximity sensors prevent collisions with any other (even previously unknown) objects, if they enter a given distance safely away from the satellite. The device according to the present invention therefore eliminates the risk of the satellite accidentally colliding with other unknown objects during its orbiting life cycle.
[0061] [0061] Preferably, the means of supplying electrical energy comprise at least one rechargeable energy source. In this way, the device has power independence from the satellite even in the event of failure of the satellite or the electrical source of the device's power supply means.
[0062] [0062] Preferably, the coupling means make an adjustable connection between the satellite and the device. More preferably, the coupling means comprise the active and / or passive control of the drive vector to effect said adjustable and / or mobile connection between the satellite and the device for aligning the drive vector. In this way, misalignment can also be done with passive control means without using any auxiliary motors to control and align the device during movement.
[0063] [0063] Preferably, the propulsion means comprises at least one solid propellant engine and at least one igniter for the solid propellant engine; wherein said igniter is operatively connected to the control means on board to receive the instant of ignition. More preferably, the propulsion means comprise one or more independent cartridges or charges filled with solid propellant.
[0064] [0064] Alternatively, the propulsion means comprise hybrid propulsion means or gel fuel propulsion means.
[0065] [0065] The objectives listed above are also achieved by a method, according to claim 22. In particular, the above objectives are achieved by a method to remove a space satellite, provided with an independent movement / removal device, from its orbit in space or move said satellite to a different orbit in space by means of remote control or by said device, characterized by the fact that it comprises at least the following steps:
[0066] [0066] - calculate, through said means of remote control or by said device, a new desired orbit in space or a desired trajectory that leads to an impact area on Earth, respectively;
[0067] [0067] - calculate, by means of remote control means or by said device, the time to activate the device to move / remove to reach the new desired orbit or trajectory based on the current orbit of the satellite;
[0068] [0068] - send from said remote control means to said device to move / remove a control signal or start the orbit withdrawal / orbit change procedure by said device to move the satellite to the new orbit or trajectory calculated;
[0069] [0069] - ignite the means of propulsion of the device to move / remove at the calculated time when it receives the control signal,
[0070] [0070] With the characteristics described above, the method according to the invention allows to implement the action of removing from orbit / changing a satellite by an associated autonomous device. This method allows a simple and controlled satellite orbit withdrawal / orbit change, calculating "a priori" a safe impact area on Earth or a new orbit in space.
[0071] [0071] Preferably, the method also comprises the step of sending a safety signal from said remote control means to said means for receiving control signals from said device to move / remove to activate an operational state before sending the said control signals to move said satellite to said new orbit or calculated path.
[0072] [0072] In this way, the start of orbit withdrawal / orbit change operations are possible only if a safety signal is sent to the device, thus preventing spontaneous ignition or unauthorized intrusion into the device.
[0073] [0073] Additional features and advantages of the present invention will become clear from the description of preferred modalities below, non-limiting examples of which are illustrated in the attached Figures, in which:
[0074] [0074] Figure 1 is a schematic representation of the types of orbit used for satellites and the methods of orbit withdrawal adopted with the device according to the present invention;
[0075] [0075] Figure 2 is a schematic cross-sectional view of a first embodiment of the device according to the present invention;
[0076] [0076] Figure 3 is a schematic cross-sectional view of a second embodiment of the device according to the invention comprising adjustable mechanical coupling means;
[0077] [0077] Figure 4 is a schematic cross-sectional view of the adjustable mechanical coupling means of the device in Figure 3;
[0078] [0078] Figure 5 is a schematic view of an active control version for the adjustable mechanical interface according to another modality of the device according to the present invention;
[0079] [0079] Figure 6 is a schematic view of an active control version for the adjustable mechanical interface according to another embodiment of the device according to the present invention;
[0080] [0080] Figure 7 is a schematic view of an active control version for the adjustable mechanical interface according to another modality of the device according to the present invention;
[0081] [0081] Figure 8 is a schematic view of an active control version for the adjustable mechanical interface according to another embodiment of the device according to the present invention;
[0082] [0082] Figure 9 is a schematic view of an active control version for the adjustable mechanical interface according to another embodiment of the device according to the present invention;
[0083] [0083] Figure 10 is a schematic view of a detail used in the adjustable mechanical interface of Figure 9;
[0084] [0084] Figure 11 is a schematic view of a valve used in the means shown in Figure 10 when, in the closed position, it prevents any flow of the fluid;
[0085] [0085] Figure 12 is a schematic view of a valve used in the means shown in Figure 10 when, in the open position, it allows the lateral flow of the fluid;
[0086] [0086] Figure 13 is a schematic cross-sectional view of an additional embodiment of the device according to the present invention, complete with means for mitigating any misalignment of the drive vector;
[0087] [0087] Figure 14 is a schematic top view of the device in Figure 13;
[0088] [0088] Figure 15 is a schematic cross-sectional view of an additional embodiment of the device according to the present invention, complete with means for mitigating any misalignment of the impulse vector;
[0089] [0089] Figure 16 is a schematic representation of the exchange of control and / or information signals between a station on Earth and the device for its remote control in accordance with the present invention;
[0090] [0090] Figure 17 shows a flow chart of the control for the basic operation of the device according to the present invention;
[0091] [0091] Figure 18 shows an extended control flowchart with respect to the representation in Figure 17;
[0092] [0092] Figure 19 shows an extended control flowchart with respect to the representation in Figure 18;
[0093] [0093] Figure 20 shows an extended control flowchart with respect to the representation in Figure 19;
[0094] [0094] Figure 21 shows a hybrid propulsion engine;
[0095] [0095] Figure 22 shows a gel fuel engine.
[0096] [0096] A device according to the invention is described in detail in the description below.
[0097] [0097] Figure 1 schematically shows a device according to the present invention associated with two different types of satellite 20 ', 20 ", orbiting around a celestial body 1, such as Earth. In particular, satellite 20' operates in a low orbit (LEO) 2, and has to be taken out of orbit towards the Earth's surface within a given space and time window of orbit. During the orbit withdrawal process, the position and orientation of the system satellite device has to ensure a safe and controlled reentry of the system when the device according to the present invention is enabled. A second 20 "satellite operates, instead, in a high orbit (MEO or GEO) 3. This is the case, for example, of satellites with high added value, such as a geostationary telecommunications satellite, or a scientific satellite. This satellite will instead be taken out of orbit towards a parking orbit 4 that is located further away than its orbit from the mission, and this will be done in such a way as to avoid interfering with other satellites or spacecraft.
[0098] [0098] A first embodiment of the device 10 according to the present invention is described below with reference to Figure 2 mentioned above.
[0099] [0099] The device 10 comprises a container body 110, preferably of cylindrical shape and made of a metallic or composite material. The metallic materials used for this purpose are preferably aluminum or steel, but other metallic materials may also be suitable. Body 110 is preferably made up of three parts consisting of a convex or semi-spherical head part 112, a cylindrical central part 114, and a flat end part 116, which can also be made in a convex shape. The parts can be made separately by milling, turning and / or rotary drawing processes, then joined, preferably by means of welding, for example, laser circumferential welding or electronic beam welding. It is also possible to join the three parts by means of mechanical joints, such as bolted flanges, bolted joints or adhesive joints, or any other method of coupling. If the fuel charge for the propulsion means is adhered to the body 110, the fuel poured into the central part 114 solidifies, thereby joining the head part 112 and the end part 16 to the central part 114, as explained below. The watertightness of the various segments is preferably guaranteed by means of the same welding joints or with the aid of elastomeric coatings. Alternatively, the head part and the center part! can be made in the form of a single spherical container (not shown). The use of composite materials allows a lighter body 110 to be prepared, but these materials have a shorter service life under vacuum conditions and in space in general, due to radiation, for example, so they have to be protected.
[0100] [00100] The device 10 also comprises means of propulsion in the form of one or more solid propellant engines, each preferably equipped with one or more independent cartridges supplied with solid propellant. These engines comprise at least one suitably solid propellant charge. 212, at least one combustion chamber 214, and at least one igniter 216, for said solid propellant. These means of propulsion also comprise at least one exhaust nozzle 218 for discharging the flue gases, preferably installed to face the opposite direction to the orbital velocity vector. The components described above of the propulsion means are sealed in a generally cylindrical container, which also contains the fuel charge 212, separated from them by a coating layer, as explained in more detail below. In the described embodiment, said container coincides with the body 110 of the device itself. Alternatively, the means of propulsion of the device may be in the form of one or more hybrid propulsion means as shown in Figure 21.
[0101] [00101] The hybrid propulsion means comprise a rocket engine that uses fuel mixtures in different forms, such as solid and gaseous form, solid and liquid form or solid and gel form. The engine comprises a compartment 213 'containing gaseous or liquid fuel or gel connected to a separate combustion chamber 214' containing the solid propellant and equipped with an igniter 216 '. The connection between compartment 213 'and combustion chamber 214' is made by one or more valves 111 'to control the flow of fuel from injector 112' from compartment 213 'to combustion chamber 214'. In addition, the engine comprises an exhaust nozzle 218 ', in a convergent-divergent manner, to discharge the flue gases, preferably installed to face the opposite direction to the orbital velocity vector, to convert said hot gases into propulsion. The non-solid fuel compartment 213 'can be pressurized to facilitate flow movement.
[0102] [00102] Compared to the solid propulsion engine, the hybrid engine has superior performance, it is safer because the fuel and the oxidizer are stored in different compartments and the same can be turned on after stopping more times, in order to modulate the propulsion . With this engine, the device must minimize propulsion misalignment or trajectory errors by performing one or more orbital maneuvers. Alternatively, the device's propulsion means may be in the form of one or more gel fuel propulsion means, as shown in Figure 22.
[0103] [00103] The gel fuel propulsion means comprise a rocket engine that uses a biofuel liquid in which the fuel and oxidant are in the form of a gel. The engine comprises a first compartment 212 "containing the fuel in the form of a gel and a second compartment 213" containing the oxidant in the form of a gel. As shown in Figure 22, the second compartment 213 "is housed in a central part of the device and the first compartment 212" surrounds the second compartment 213 ". It is also possible to store fuel and oxidant together as a gel-shaped mixture in a single The engine comprises a separate combustion chamber 214 "connected to compartments 212", 213 "by one or more valves 110", 111 "that control the flow of gel fuel from the injector 112" from compartments 212 ", 213" to the combustion chamber 214 ". In addition, the engine comprises a 218 "convergent-divergent exhaust nozzle for discharging the flue gases, preferably installed to face the opposite direction to the orbital velocity vector, to convert said hot gases into propulsion. ", 213" for the gel fuel can be pressurized to facilitate the flow movement.
[0104] [00104] Regarding the solid propulsion engine, the hybrid engine can be restarted after stopping more times, in order to modulate the propulsion. With this engine, the device must minimize propulsion misalignment or trajectory errors by performing one or more orbital maneuvers. The device may additionally comprise safe ignition means operatively connected to the control means on board to receive the ignition control signal and to operate the ignition of the propulsion means. In particular, said safe ignition means may form part of the propulsion means as a safety and ignition unit (SIU) integrated in the igniter.
[0105] [00105] In this mode, the ignition device 216, and consequently any safety and ignition unit, are operationally connected to the onboard control means described later, to receive the ignition signal. Said signals can be in an encrypted mode to secure the transmission and prevent unauthorized intrusion and ignition. The igniter 216 contains a charge of energetic material in a suitable container (preferably cylindrical). In this embodiment, which comprises a solid propellant propulsion means, this container has an opening towards the main solid propellant charge. In the mode described in this document, igniter 216 is inserted into a cylindrical hole in fuel charge 212. When igniter 216 is enabled, this opening allows combustion products to make contact with, and consequently ignite, the charge main fuel 212. The opening is preferably round, sealed by means of a sheet of normal metal, and designed to explode as soon as the igniter 216 has been enabled. The charge within the igniter 216 can consist of different types of mixtures known to generate hot gases and a large amount of glowing metal oxide particles, which in turn can ignite the main charge of solid propellant 212 when they come in contact with the latter. Mixtures of boron potassium nitrate or potassium nitrate, or pyrotechnic fillers can generally be used, with the optional addition of a conventional solid propellant filler as a filler. This charge can be cylindrical with a central perforation, or simply a cylindrical fuel piece, or it can consist of a plurality of small fuel pieces so that its resulting larger combustion surface area reduces the total combustion time of the device ignition. If the igniter 216 contains only a pyrotechnic mixture, it is all contained within the igniter. If, on the other hand, there is also a reserve charge, this charge and the pyrotechnic mixture are kept separated within the igniter 216 by a thin membrane. This membrane is broken when the pyrotechnic mixture is ignited, causing ignition of the reserve charge and rupture of the leaf covering the opening of the igniter 216. This releases the combustion products from the igniter 216 towards the main charge of fuel 212. The pyrotechnic mixture can be in grains or small tablets, depending on the dimensions of the device 10. The container for the igniter 216 is aligned with the axis of the main charge and designed so that it cannot explode when the ignition device 216 is enabled. The igniter is preferably coupled with the head part 112 of the body 110 by means of a joint, such as a screw connection, which is sealed by means of an elastomeric sealing ring type coating, or with the aid of an adhesive, for example. Alternatively, the igniter 216, according to the description above, can also be inserted directly into the central hole in the load 212.
[0106] [00106] The safety and ignition unit (not shown) is enabled by electrical signals generated by the control means on board with which it is operationally connected, and it is used to enable the ignition device 216. The SIU unit it can include a high voltage inline igniter, a standalone (offline) low voltage igniter, or a low voltage inline igniter. In the SIU fitted with a high voltage in-line igniter, the ignition pulse is transmitted directly to the ignition device charge 216 by means of an explosive foil trigger or detonator with a striker integrated in or positioned close to the pyrotechnic charge in igniter 216. These mechanisms are well known in the literature. The unit with a stand-alone low-voltage igniter consists of an electrical cable immersed in a small amount of pyrotechnic mixture identical or similar to that of the igniter 216. This mixture is encapsulated and sealed. The ignition device is made immune to any electromagnetic interference by means of an electric filter. The igniter electronics may also include an integrated test circuit to test the state of the explosive foil trigger or low voltage igniter. Alternatively, an unintended activation state can be detected by the autonomous low voltage version if the piston it contains moves and consequently changes the state of a circuit, for example, breaking a limit switch, or by pressing a button. The SIU can be made even more secure by adding one or more transistors. In particular, the activation and ignition signals sent by the on-board control means are actually transmitted to the active components of the ignition device 216 (the electrical charge of the capacitor or the explosive foil trigger, or the low voltage igniter ) only if the transistors were enabled by separate signals generated by electronic boards not electrically coupled and totally independent.
[0107] [00107] As previously described, the fuel load 212 in the first embodiment (shown in Figure 2) is preferably attached to the body 110. Alternatively, the propulsion means comprise one or more independent cartridges supplied with solid propellant. Said separate cartridges are preferably attached again to the body 110. Said fuel charge 212 can also be divided into two segments separated by a membrane to thereby produce a double pulse. The gases generated by the first charge, therefore, are released directly through the nozzle 218, while the gases generated by the second charge reach the nozzle 218 through a duct (not shown), which is protected with a layer of ablative material, such as a phenolic resin reinforced with silica. The combustion volumes of the two charges are separated by a membrane that prevents their simultaneous ignition. Therefore the second charge is ignited by an additional igniter (not shown). The formulation of the solid propellant and the shape of the grain must have characteristics such as to ensure that it fills the available volume and maximizes the specific impulse, while being insensitive to pressure. In particular, the shape of the cargo and the formulation of the fuel used must maximize its performance and minimize its mass and volume, while maximizing the necessary safety requirements. The solid propellant charge 212 can also have regressive combustion characteristics, consuming most of the fuel immediately after ignition and as little as possible thereafter. The combustion rate must be as high as possible, subject to the maximum propulsion demands that derive from the structural requirements of the 20 ', 20 "satellite and the mechanical coupling interface between the device 10 and the 20', 20" satellite, described posteriorly. If the load is inserted in a cartridge, the fuel is poured into a rubber container that in turn is attached to the body 110. If the fuel is poured directly into the body 110, as described for the first modality in Figure 2, then the fuel filler 212 is attached to body 110 by means of a coating layer, i.e. an elastomer that can be made from the same binder as fuel 212 and contains a filler, such as carbon black. The liner is connected to the fuel load 212 by means of an excess of recirculating agent and coupling agent. The length-to-diameter ratio of the fuel load 212 is usually too small to minimize the total length of the device 10.
[0108] [00108] The combustion products are released through an exhaust nozzle 218, which is an integral part of the body 110, in order to obtain necessary propulsion, as previously explained. In the first embodiment illustrated in Figure 2, this nozzle has a convergent-divergent shape and is made of a metallic or ceramic material, depending on the material used to make the body 110 to which it is attached. The nozzle 218 can be at least partially embedded in the body 110 to reduce the dimensions of the device 10. A nozzle made of a metallic material (for example, aluminum, steel or the like) can be protected from the hot gases generated by the combustion process by by means of a suitable layer of ablative material, such as phenolic resin reinforced with silica or other known compositions. This protection may be unnecessary if the combustion time is limited. A ceramic nozzle can be made of a monolithic ceramic material, or a ceramic material reinforced with a composite material containing long or short fibers. The same is preferably made of a "carbon-carbon" ceramic material (carbon fibers in a carbonic matrix) or a C-SiC (carbon fibers in a silicon carbide matrix), or even a CC / SiC (carbon- carbon / silicon carbide). The diameter of the passage in the nozzle 218 is large enough to generate a strong propulsion and to expel a considerable gas flow, minimizing the total combustion time. In the first embodiment, the nozzle 218 is attached to the end part 116 of the body 110 by means of screws 218 ', 218 ". Alternatively, the nozzle 218 can be installed inside a specific container, which is screwed or attached in some way to the body of the propulsion means 110. Or it can be directly attached to the body 110, or installed by means of a flange integrated to the shape of the nozzle 218. in said embodiment, the convergent-divergent nozzle 218 is connected to the body 110 which serves as a combustion chamber.
[0109] [00109] Nozzle 218 is also completed with an environmental seal (not shown), which consists of a membrane that covers the nozzle, separating the solid propellant charge 212 from the external environment, this prevents any contamination of the solid propellant in transit from the site of production to the location of the launch by humidity, or any loss of volatile species such as the plasticizer of the fuel. The membrane is preferably made of metal or plastic and welded to a metal ring with a screw or glued to it. The ring is sealed by means of an elastomeric coating, such as a seal ring, or welded or glued to the nozzle 218. The seal has a thermal protection layer to prevent any excessive transfer of thermal energy between the external environment and the interior of the device 10. The membrane is preferably engraved with a cross-shaped motif so that it can open when the device 10 is lit. This allows the membrane to be opened in four sections of equal size at the same time, however it remains attached to the ring, and consequently to the nozzle 218, in order to avoid any further contamination of the space with additional debris. Alternatively, the membrane can be weakened in the vicinity of the circumference of the ring and attached by means of a chain or metal wire to the outside of the nozzle 218. In both compartments, the risk that parts of the environmental seal are released in Space is avoided or minimized, thereby preventing further orbital pollution.
[0110] [00110] The propulsion means are operationally connected to the control means on board (not shown in the Figures) serving device 10, preferably comprising electronic controllers designed to be immune to electromagnetic interference or radiation that occur in space, in the described mode in this document, the control means on board specifically consist of an electronic board fitted with microcontrollers and an electronic and / or electrical interface for connecting to the additional means comprising the device 10 operationally connected to them. In particular, the onboard control means send instructions to and / or receive information from said means on device 10. These on-board control means also allow device 10 to be independent of the satellite 20 ', 20 "to be taken out of orbit / changed orbit Another function of the onboard control means in the present modality is to manage and process controls and signals exchanged with a receiving / transmitting device at a station on Earth or other space vehicles. On-board control means activate and ignite device 10, thus enabling the activation of the SIU and consequently also the propulsion means or, more specifically, the igniter 216.
[0111] [00111] An additional purpose of the onboard control means is to send data regarding the operating status of the device 10 and / or the satellite 20 ', 20 "to the Earth or to another space vehicle, or to the satellite 20', 20 "with which device 10 is coupled, by a telecommunications unit, which recorded said data by appropriate sensors, as explained in the detail below.
[0112] [00112] The means for receiving control signals (not shown) comprise one or more low gain or high gain antennas and an electrical and / or electronic interface between the antennas and the onboard control means to which they are connected operationally. These means for receiving control signals are preferably part of the telecommunications means (not shown) operationally connected to the control means on board. In particular, in said mode the telecommunications means also comprise additional means for transmitting data, including at least one transmitting antenna for sending signals and devices suitable for receiving the same. The resulting communications are preferably exchanged directly with a receiving station on Earth or with another vehicle located in space, using an adequate communication bandwidth. This communication enables the status of device 10 and / or associated satellites 20 ', 20 "to be verified, as explained below. The communication channel also allows confirmation signals to be sent, related to effecting activation or ignition of the device For example, signals sent to and from telecommunications media, whether from Earth or from space, have transmission characteristics designed to minimize their mass and volume.The telecommunications media are also designed to be immune to electromagnetic interference and radiation in space.
[0113] [00113] The electrical power supply means (not shown) for the device 10 preferably comprise one or more primary rechargeable or non-rechargeable batteries, connected together to provide sufficient energy to ignite the device 10 when necessary. The released power must also be able to sustain the other components of the device 10 throughout its useful life. If the batteries are of the rechargeable type, they can be recharged directly by the satellite 20 ', 20 "with which the device 10 is connected, as long as there is an electrical connection between the two. In the preferred mode, these batteries can be recharged using any form of energy generation suitable for operation in space, such as photovoltaic technology, which can also be recharged by exploring a planet's magnetic field, or an energy collection device based, for example, on the temperature difference between two points on device 10 (for example, between a point exposed to the sun and one in the shade). These electrical power supply means are also designed to be immune to electromagnetic interference and radiation in space. Electric power supply means are operationally connected to, and controlled by on-board control means, for which they also deliver an energy supply gia.
[0114] [00114] The mechanical coupling means 310 couple said device 10 to the satellite 20 ', 20 "before it is launched. In this embodiment, said coupling to the satellite is achieved by means of a mechanical interface platform. This mechanical coupling is completed before launch, namely, before the satellite is in service. Therefore the mechanical coupling means 310 allow the simultaneous displacement of the satellite 20 ', 20 "and device 10, when the latter is activated for the purpose of withdrawing from orbit / change the device-satellite system's orbit.
[0115] [00115] In the first embodiment shown in Figure 2, the mechanical coupling means 310 comprise a single joint fixed between the device 10 and the satellite 20 ', 20 ", positioned centrally with respect to the head portion 112 of the device 10.
[0116] [00116] The device 10 is also equipped with means of thermal protection, in the described mode, these include passive thermal insulation to limit the temperature change in the most crucial components of the device 10 and to guarantee low temperature gradients. This is because a high temperature, for example, accelerates the chemical aging of fuel and coatings. Low temperature thermal cycles, on the other hand, cause tension and distension in the fuel, negatively influencing the volume it occupies and consequently also its performance. Limiting the minimum temperature that the fuel load can reach in orbit therefore leads to an improvement in its performance. For fuel, the minimum allowable temperature roughly coincides with its glass transition temperature (for finished hydroxy polybutadiene fuels [HTPB], the minimum temperature should not fall below -60 ° or -80 ° C, depending on the cooling rate). The maximum permissible temperature is that at which the fuel starts to deteriorate chemically very quickly or even to ignite (maximum temperatures should not exceed 70 ° C or 80 ° C). The other components that also require thermal insulation are the electronic units and the elements that comprise the means of supplying electrical energy. A thermal insulation system preferably consists of multilayer insulators (MLI), that is, layers of insulation material contained within a sheet of metal, special colored paint, or other passive or active systems. Passive thermal insulation is essential to ensure the same temperature in the fuel load and any small non-reusable auxiliary rockets, as described later. Even more preferably, the insulation can be done by inserting a layer of highly conductive material under several insulating layers. in the present modality, additional active thermal insulation is used preferably to protect the electronics of the device 10 and its means of supplying electrical energy when the temperature differences in these components are not adequately controlled by passive insulation. Alternatively, it should be possible to use this active system in isolation if the passive system becomes unnecessary for the thermal protection of the device 10 components. Additionally, if the 20 ', 20 "satellite is equipped with thermal protection, it becomes possible to implement actions and synergies with device 10 to contain temperature changes in the latter.
[0117] [00117] During the assembly of the device 10 on the satellite 20 ', 20 ", small errors can occur in the alignment of the impulse vector in relation to the main axes of inertia that pass through the center of mass of the device-satellite system. Systems severity can also vary uncontrollably over the lifetime of the 20 ', 20 "satellite (for example, due to a defect in the latter), giving rise to an excessively large impulse vector error for the efficient operation of the device 10. To prevent this, the device is equipped with means to mitigate the misalignment of the impulse vector, operationally connected to the control means on board. In particular, said means for mitigating misalignment may comprise means for active and / or passive control of the drive vector which effects the drive vector alignment, as described below in an additional embodiment.
[0118] [00118] The device may additionally comprise means for detecting and / or controlling the behavior of the satellite with which it is coupled, operationally connected to the means to mitigate misalignment. In addition, said means for detecting and / or controlling the behavior of the satellite are operationally connected to the control means on board and perform directional control of the device and the satellite when the propulsion means are activated.
[0119] [00119] Means for active and / or passive control of the impulse vector make an adjustable and / or mobile connection between the satellite and the device are used in the modality described below. In this mode, a passive or actively adjustable mechanical interface aligns the drive vector.
[0120] [00120] In the second embodiment, shown in Figure 3, the device 20 is as described above for the first embodiment, but it also comprises adjustable means 320 for mechanical coupling to the satellite 20 ', 20 ". These mechanical coupling means 320 perform a adjustable and / or movable connection between satellite 20 ', 20 "and device 20. This allows an adjustment of the orientation of device 20, during assembly and / or passive control of the system, to align the propulsion direction with that of the center of mass of the satellite 20 ', 20 ", possibly with a blocking action at the end of said alignment.
[0121] [00121] The adjustable mechanical coupling means 320 according to the description given for the modality of the device 20 in Figure 3, are illustrated in more detail in Figure 4. They are realized with an adjustable passive mechanical interface comprising a first part 322 in contact with device 20, and a second part 324 in contact with satellite 20 ', 20 ". The adjustable mechanical coupling means 320 also comprise a ball joint in which two flanges 326', 326", with a semi-concavity - spherical on one side and a flat interface fixed on the other (in contact with the first part), contains a hollow sphere 325. Therefore, device 20 can rotate in relation to the central axis of satellite 20 ', 20 "by means of relative movement of the first part 322 of ball 325. Fine adjustment is achieved by rotating the ball joint after it has been installed on flanges 326 ', 326 ". To reduce the impulse vector error to zero, the device can be rotated until the nozzle axis 228 in the propulsion means passes through the center of gravity of the system-satellite device. The joint is then locked in the required position by means of an appropriate number of screws 328 ', 328 ". If ball 325 is made of a softer material than screws 328', 328", then the screws are screwed into position until they pierce the sphere, thus locking it in place. Otherwise, adjusting a sufficient number of screws or large screws would interfere with the rotation of the ball 325, due to friction, for example. An alternative solution is to use a sphere of magnetic material inserted in a cavity of non-magnetic material. Screwing the magnetic poles until they are close enough to the sphere should allow the magnetic force to prevent its rotation. To avoid any loosening of the screws, the chosen material cannot be subject to deformation as a result of the temperature differences that occur when the satellite is launched or is in orbit. The screws are also preferably locked with the aid of a thread-locking adhesive.
[0122] [00122] Another modality (not shown) involves the nozzle 218 being connected to the body 110 by means of a flexible joint. This solution is useful in special applications, such as satellites operating in a geostationary orbit. When this solution is adopted, if the interface for the mechanical coupling (described later) is adjustable, it can be simplified, requiring adjustment only before the satellite is launched, during the assembly stage. The adoption of this last coupling solution with a flexible joint makes it unnecessary to adopt a system to mitigate any misalignment of the impulse vector, as described in more detail later.
[0123] [00123] In some compartments, however, active control over the alignment of the drive vector via the adjustable interface may be the best way to explore the full potential of the present invention. This active system is particularly effective in compartments where the satellite's center of gravity 20 ', 20 "changes during its mission and it is impossible to calculate its position in advance, before enabling device 10 (due, for example, to a satellite defect 20 ', 20 "). Active alignment control is also useful in compartments where a small 20 ', 20 "satellite, with relatively small moments of inertia, tilt and rotation, does not have a system to mitigate any propulsion misalignment (described below), and the device 10 it takes a relatively long time to produce the necessary thrust An active method for adjusting the propulsion direction, and the related active mechanical coupling means, can use hydraulic, electric or gas driven pistons located between the second part 324 and the 326 ', 326 "flanges. The adjustment depends on the position of the pistons and can be easily controlled using an inertial platform and a standard proportional control system. A piston is required to control one axis, two pistons for two axes and so on, that is, the number of pistons increases the more axes to be controlled.
[0124] [00124] An active control version for the adjustable mechanical interface 330 is shown in an additional modality in Figure 5. The position relative to the angle of rotation around an axis, such as the inclination axis, is controlled by means of a pair of articulated cylinders 332 ', 332 "that come into contact off center with respect to the main joint 334. When the main joint to control the second axis 336 is locked and the cylinders 332', 332" are operated, it generates a rotation 335 of the second part 324 of the mobile coupling means 330 and consequently of the satellite 20 ', 20 "coupled to them. This allows the mutual rotation of the device and satellite around a first axis, for example, the tilt axis. 337 at the joint (which consists of a bar inside a hole, for example) prevents any moment of bending.The control over the position, for example, of the rotation angle in relation to the other (for example, rotation) axis is identical to that le just described and is exercised by means of an identical mechanism located below or above the tilt control, rotated by 90 °. A more compact version (not shown) of the mobile mechanical coupling means can be made using a cross-shaped element. This element comprises two bars separated from each other and joined at its center by means of a cylindrical element. This cylindrical element can be milled or made from turned cylindrical bars which are then screwed, welded or glued. The upper bar allows the rotation of the flange facing the satellite while the lower bar allows the rotation of the flanges in contact with the device.
[0125] [00125] Another modality, which differs in the actively controllable adjustable mechanical coupling means 340 ', 340 "is illustrated in Figures 6 and 7. Position control in tilt or rotation, or both, is achieved by means of two cylinders with a joint spherical top 341, 34,342, 342 ', installed in line with the second part 344', 344 "in contact with satellite 20 ', 20". In particular, this part is shown with a dome shape in the modality in Figure 6 , and a flat shape in the modality in Figure 7, but it can also be of any other form. said second part 344, 344 'there is a ball joint 346, 346', which allows rotation and does not need to be displaced axially. The two cylinders described above 341, 341 ', 342, 342' and the ball joint 346, 346 'also are positioned at an angle of approximately 120 ° to each other on the surface of the second part 344, 344 '. If one of the cylinders 341, 341 ', 342, 342' is moved, the surface of the second part 344, 344 'tilts with respect to an axis connecting the other cylinder to the ball joint 346, 346'. Controlling the pistons of the two cylinders 341, 341 ', 342, 342' consequently allows any necessary rotation of the thrust vector. These pistons can be controlled electrically, hydraulically, or pneumatically.
[0126] [00126] Another modality of the actively controllable adjustable mechanical coupling means 350 is illustrated in Figure 8. This modality, like those described previously related to Figures 6 and 7, comprises two cylinders 352, 352 ', but they are designed differently of the previous cylinders. The base part 351 of these adjustable mechanical coupling means 350 can be coupled either with the satellite 20 ', 20 "or with the device according to the present invention. The means comprise first cylinders 352, 352' consisting of a first piston 353, 353 'which can be operated electrically, hydraulically, or pneumatically. The piston 353, 353' can rotate around its own axis and be fixed to the adjustable mechanical coupling means 350 by means of a hinge 354, 354 '. second piston 355, 355 'is coupled to joint 354, 354' so that it can rotate in relation to cylinder 352, 352 'and therefore also in relation to first piston 353, 353'. on the opposite side the joints 354, 354 ', said second piston 355, 355' is directly coupled to a second cylinder 356, 356 'which is directly coupled to the base part 351, and can also be controlled electrically, hydraulically, or pneumatically. The base part 351 can then rotate (axis of rotation not shown) and m with respect to piston shaft 355, 355 '. If one or both of the first cylinders 352, 352 'have been operated, the base part 351 can be oriented in any required direction. The base part 351 also comprises a centrally located ball joint 357.
[0127] [00127] The same directional control is achieved using three or four bellows-type pistons, as in the embodiment in Figure 9. This Figure shows 360 actively controllable adjustable mechanical coupling means in which pistons 362, 362 ', 362 "are electrically controlled, hydraulically, or pneumatically, for example, by means of a pressure tap (not shown), and preferably directly by the device's engine. Each piston 362, 362 ', 362 "is connected to the engine via a three-way valve ( not shown), in which one track is connected to the device's motor, one communicates with piston 362, 362 ', 362 "to be controlled, and the third is a side vent that is normally closed. When the motor is started, the pistons 362, 362 ', 362 "are under pressure and the valve closes. To adjust the direction of the device in relation to the satellite 20 ', 20 ", orienting the impulse vector in the required direction, the third valve of one or more pistons 362, 362', 362" opens for a given time, releasing a little consequently reducing the pressure in the pistons. This allows pistons 362, 362 ', 362 "to recede and as a consequence the device rotates. As shown in Figure 9, the three pistons 362, 362', 362" are of the pneumatic type and coupled to the base part 361, which can stay in contact with the device according to the present invention, or with the satellite 20 ', 20 ". The pistons 362, 362', 362" can be positioned close to the edges of the base part 361 or in any other suitable configuration. The coupling with the base part that makes the interface 361 is carried out by means of the ball joint 363 that allows a relative rotation between the axis of pistons 362, 362 ', 362 "and the platform. As illustrated in Figure 10, pistons 362 , 362 ', 362 "are governed by a gas under pressure and have a flexible conduit 365', 365" that allows them to behave like a spring, extending in the direction Δ of the double arrows shown in Figure 10. These flexible conduits 365 ', 365 "are well known in the literature, in the present embodiment they are charged either by a separate gas generator or directly by the means of propulsion of the device according to the present invention. When it is loaded, the conduit 365 ', 365 "stretches under the effect of pressure, to control the relative position of the satellite and the device according to the invention, it is only necessary to release a small pressure from one of the pistons 362, 362 ', 362 "with the aid of a bypass valve 368 (in Figures 11 and 12). The gas can be released laterally through four vents 366, 366 ', 366 ", placed at 90 ° angles to each other to avoid any lateral impulses. Alternatively, the gas can be vented longitudinally in the direction and propulsion of the propulsion means, at the same time increasing their performance.This controlled gas ventilation is done through one of the three bypass valves 368, as shown in Figures 11 and 12. The ventilation duct is normally closed (Figure 11), while the duct between the gas source and the flexible duct is normally closed (Figure 11). Valve 368 comprises a piston 367 inserted in a pipe 369 that connects the gas duct 365 ', 365 "to the base part 361. The course of the pistons 362, 362 ', 362 "is limited by two obstacles 370', 370". The gas generator (which coincides, in the mode described in this document, with the combustion chamber of the engine of the device according to the present invention; alternatively, it can be an external gas generator), is normally connected to the gas conduit 365 ', 365 ". Rubber seals 371, 371', 371", 372, 372 ', 372 "around the side ventilation ducts prevent any gas leaks. When controlled gas release is required, the main pipe between the gas generator and gas conduit 365 ', 365 "is closed (Figure 12) by means of a valve (not shown), such as a solenoid valve. The piston 367 then begins to move, electromagnetically or hydraulically controlled, towards the bottom stop stop identified by the obstacle 370 ", thereby discovering the gas vent hole 373. This allows gas to be released from the conduit gas', 365 "and dissipated in space. An active directional control of this type could require an electronic "data recording and processing" system to allow a feedback control to be implemented, as in well-known proportional derivative control systems. The sensors used are preferably accelerometers to record the angular accelerations around the two axes that control the drive vector. The electronic components described above, for example, the sensors and control system, could be the same as the means of detecting behavior and control on board described later, and / or be part of an inertial platform. These electronic components are operationally connected to the on-board control means for the device according to the invention, which allows greater precision in the final positioning of the device-satellite system during the orbit withdrawal made by said device. These means also allow for a reduction in losses due to any misalignment of the drive vector when the propulsion means of the device according to the invention are working. In an additional embodiment, means to mitigate any misalignment of the impulse vector are used if the position of the device in relation to the satellite 20 ', 20 "to which it is attached is fixed prior to launch and, for the time it is decided that satellite 20 ', 20 "has to be taken out of orbit, or placed in a parking orbit, the losses of useful impulse due to misalignment of the impulse vector have become excessive. These losses depend on the average size and density of the satellite 20 ', 20 ", as well as the possibility of obtaining a constant or variable center of mass, as in the case of moving parts, such as detachable solar panels, or sufficient fuel consumption. to modify its mass. Losses due to misalignment of the thrust vector are reduced a priori by designing the device so that it has a very fast combustion time, a high combustion rate of the solid propellant, and a large neck diameter of the nozzle, as already explained, the means to mitigate can, for example, carry out a stabilizing rotation around the axis of rotation with the aid of small auxiliary rockets, or use similar non-reusable small auxiliary rockets installed off-center with respect to the nozzle at the rear of the device, once lit, these rockets generate torque along the tilt or rotation axis, depending on their positions.
[0128] [00128] Figures 13 and 14 illustrate an embodiment of the device according to invention 40, complete with means to mitigate any impulse vector misalignment. This modality mitigates the misalignment of the impulse vector by stabilizing the satellite by means of a rotation around the axis of rotation obtainable with the aid of two small non-reusable auxiliary rockets 410 ', 410' which are enabled by the SIU (not shown) on the device 40. The 410 ', 410 "rockets are small in size and, when they are enabled, they generate movement around the axis of rotation with a rotation speed that depends on the combustion time and propulsion capacity of the rockets . Rockets 410 ', 410 "are preferably enabled before device 40 is switched on, via a cable or radio signal sent by SIU on device 40. Nozzles 411', 411" of said rockets 410 ', 410 "are positioned at an angle of 90 ° to the flight direction in order to produce a torque to propel around the axis of rotation of the satellite 20 ', 20 ". When they are lit, the rockets produce a torque that makes the device 40, and tell also the satellite 20 ', 20 ", rotate around the axis of rotation. Then device 40 is lit. The rotation of the 20 ', 20 "satellite mitigates the effects of any propulsion misalignment due to the propulsion component perpendicular to the direction of flight inducing a precession movement. If the precession period is longer than the total operating time of the withdrawal from orbit 40, the net effect of misalignment will be small. If the angular momentum induced by rotation is considerable, that is, the angular velocity induced by rockets 410 ', 410 "is high enough, the angular error of misalignment of the misalignment propulsion will be kept within an acceptable level. Another modality of the means to mitigate the misalignment of the thrust vector is illustrated in Figure 15. Device 50 is fitted with small non-reusable auxiliary rockets 510 'similar to those previously described 410', 410 "rockets. Additionally, there are also 512 rockets, 512 ', 512 ", similar to rockets 510', 510" and positioned off center with respect to nozzle 518 of device 50, that is, behind the nozzle. When these rockets 512, 512 ', 512 "are lit, the same generate rotation or inclination torque, depending on their positions.
[0129] [00129] The rockets described in Figures 13, 14 and 15 are in the form of a solid propellant charge. In particular, they are non-reusable single impulse powered by solid propellant. They can be in greater number and are preferably arranged in pairs on each axis. The previously described rockets 410 ', 410 ", 510', 510", 512, 512 ', 512 "are simple to manufacture and of limited dimensions, and they can be operated under the direct control of the onboard control means for the device 40, 50 according to the present invention, being positioned well away from the nozzle axis 418, 518 of device 40, 50, they produce a low level of propulsion, however this propulsion generates a greater torque around the axles of tilt and rotation of what is generated by the misalignment of the propulsion. The rockets 410 ', 410 ", 510', 510", 512, 512 ', 512 "can be attached to the outer walls of the device 40, 50 or moved further away from the nozzles 418, 518 by means of a bar or beam (not shown). If device 40, 50 identifies angular accelerations around the tilt or rotation axes (using instruments to detect accelerations such as those described below), the corresponding rocket is ignited to produce propulsion in the opposite direction. The system is designed so that the rocket has a "balancing" effect and is ignited with a delay that takes into account angular acceleration due to misalignment and the maximum potential for ignition delay of the rocket. The thrust produced as a result induces sufficient acceleration to restore the direction of the thrust vector to the position required initially. Since the total operating time of the device 40, 50 is very limited, the operating and propulsion times of the non-reusable auxiliary balancing rockets are calculated to avoid overreaction. In the event of an error, the rocket opposite the one used for the correction can be ignited to produce an additional counter propulsion. Therefore, the correction of misalignment of the impulse vector can be performed using rockets both placed in contact with the body to obtain a stabilization rotation and in the vicinity of the nozzle for balancing purposes, or to adopt only one of the two solutions.
[0130] [00130] In an additional modality (not shown), defined as 'independent', the device according to the present invention generates the activation and ignition signals alone, without any assistance from a ground station or other stations in the space. This modality can be performed starting from one of the modalities described previously and / or with parts of them. However the same in the different functions of the control means on board the device due to the means for receiving control signals do not receive the ignition signal, but these are generated independently by the device at a predefined time or under predefined conditions.
[0131] [00131] The device according to the present invention works independently of the satellite 20 ', 20 ", and the purpose of these means for receiving signals is not to receive signals related to the activation and / or ignition of the device of stations on Earth or in others space vehicles, but only receive signals for the purpose of stopping the device independent operation sequence, as described in detail below. Provision can be made for said means to receive control signals to be able to receive signals from external emitters to the purpose of interrupting the ignition procedure and enabling its subsequent remote control.In this mode, the on-board control means include means for emitting control signals that have been pre-programmed and / or calculated by the on-board control means and for sending the same for the means to receive control signals. The latter, also being associated with the means of propulsion, then effect the activation and ignition of device 10, therefore enabling the igniter independently.
[0132] [00132] The means to emit the control signals are in the form of electronic controllers and preferably comprise an ignition enable timer adjusted to suit the life of the satellite 20 ', 20 "or the duration of its mission. Alternatively, the time of the orbit withdrawal procedure can be calculated by the control means on board the reception of warning signals from sensors optionally adjusted in the device according to the present invention, as described later.
[0133] [00133] In another mode (not shown), called semi-independent, the device according to the present invention receives the activation signal from a ground station or from other vehicles located in space. This modality consequently has a structural configuration and components similar to the device 10 described in the first modality, except that the ignition signal is generated independently by the device according to the present invention. The modality described in this document also has the means to emit control signals as described and implemented in the previous modality. These means for emitting control signals send to the means for receiving control signals an ignition signal generated by the onboard control means after receiving the activation signal from the ground station or another vehicle in space.
[0134] [00134] All the illustrated and described modalities can also comprise optional means to improve the functional characteristics, or add new characteristics to the device according to the present invention and the system that derives from its coupling to the 20 ', 20 "satellite. An implementation of the characteristics of the device according to the present invention in one or more of the modalities described previously includes telecommunications means that also allow the exchange of commands and signals with the satellite 20 ', 20 "with which the device is coupled, although the latter remains entirely independent of the former.
[0135] [00135] In one or more of the modalities described in this document, the device according to the present invention also preferably comprises first detection means, operationally connected to the control means on board, to detect other objects arriving within a security zone defined around the device and / or the satellite 20 ', 20 "with which it is coupled. Even more preferably, the device according to the present invention includes means for communicating with the satellite 20', with which the it is coupled with the purpose of detecting any impact damage to the satellite. The device preferably also includes third party detection means, operationally connected to the control means on board, to detect any defects in the satellite 20 ', 20 "with the aid of the means communication.
[0136] [00136] The device according to the present invention also comprises second detection means, operationally connected to the control means on board, to detect impact damage to the device itself.
[0137] [00137] In one or more of the modalities illustrated and / or described, and in combination with one or more of the technical implementations mentioned above, the device according to the present invention may additionally include four detection means, operatively connected to the control means on board, designed to detect the time orientation and / or orbit of the device according to the present invention at any time, or at predetermined times, or to detect the space position of the device according to the present invention. These four detection means are preferably operationally connected to an orientation and position monitoring unit also used to independently establish the position and orientation of the device and the 20 ', 20 "satellite with which it is coupled. This unit is used to increase the accuracy of calculating the orientation and position of the system in order to reduce any errors in calculating the reentry path. Another purpose of this unit is to send instructions for the means to mitigate the misalignment of the drive vector and / or for the means to control the impulse vector actively in the adjustable mechanical coupling means. The unit is designed to be immune to electromagnetic interference and radiation occurring in space. These sensors are preferably also operationally connected to the status monitoring means. They collect test results electrical devices arriving from the means installed in the device according to the invention as well as the signals coming from the monitoring sensors previously described. The collected data are sent, upon request or at regular intervals, to the control interface located at a station on Earth, or to another space vehicle, using the telecommunications means with which the device is equipped. Alternatively, the device according to the present invention can send this data to the satellite 20 ', 20 "with which it is associated, exploiting said means of communication of the satellite. This communication can also be bidirectional, so that the satellite 20 ', 20 "can send control signals to the device.
[0138] [00138] With reference to the modalities in Figures 13 and 15, the Figures show two containment means 492 ', 492 ", 592, in the form of parallelepiped containers coupled with the head part of the body to the device 40, 50. These containment means preferably contain means of supplying electrical power (such as non-rechargeable batteries), on-board control means and telecommunications means as previously described. These containment means 492 ', 492 ", 592 may also contain the means of monitoring status and the means to actively control the propulsion, as well as any additional means for detecting and controlling behavior, and the means to mitigate the misalignment of the impulse vector. These containment means 492 ', 492 ", 592 can be positioned either at the end of the device body or at the rear, close to the nozzle.
[0139] [00139] In an additional modality (not shown), the device may include a system to position the satellite with which it is coupled (or emergency tipping system - EDS) to ensure that the satellite's behavior is appropriate and stable before the device is turned on. This system may prove necessary if the satellite's behavior is out of control or if its behavior is such that the direction of propulsion generated by the device is not aligned with the orbital speed and the satellite makes any additional maneuvers impossible. The system is powered by one or more gas generator cartridges loaded with solid propellant, ignited by means of an electrical explosive device (with the addition of a suitable pyrotechnic mixture, if necessary) and is operationally connected with the control means on board. that govern the ignition of the device.
[0140] [00140] In its various possible modalities, the device according to the present invention is used to prevent the further accumulation of debris in orbit and the risk of damage to people or property caused by the controlled non-re-entry into Earth by satellites, space vehicles or parts of them.
[0141] [00141] In particular, device 10 is capable of changing the trajectory of satellite 20 ', 20 "to which it is associated when the latter reaches the end of its useful life or develops a defect. As illustrated in Figure 1, this The change in trajectory allows the satellite to be taken out of orbit directly towards Earth 1, where it disintegrates when it moves through the atmosphere or can land within an area of arbitrarily defined dimensions or previously established. may involve redirecting the satellite towards a safe area in space, defined as a graveyard or parking orbit. As shown in Figure 1, if the satellite's 20 'orbit is low 2 and the path change consists of removing the orbit even towards the celestial body at center 1, or towards a lower orbit, then device 10 generates a propulsion partially or completely turned in the opposite direction to that of the vector of orbital speed (in the direction of flight) of satellite 20 '. If, on the other hand, the orbit of satellite 20 "is high 3, then the change in trajectory consists of relocating satellite 20" to a parking orbit 4 or graveyard 4 further away from planet Earth 1. In this case, the direction of propulsion generated by device 10 will be in the same direction as the orbital velocity vector, for both types of orbit withdrawal procedure, satellite 20 ', 20 "will follow a specific calculated and established trajectory before any steps are taken to modify its orbit to avoid putting a risk to other satellites or space vehicles, people or properties, whether they are the same in space or on Earth 1.
[0142] [00142] Device 10 is installed on satellite 20 ', 20 "before it is launched. This assembly is carried out before launch using means 310 to mechanically couple said device 10 to satellite 20', 20" to be removed from orbit . In particular, the device 10 can be coupled in front of the satellite 20 'if it has to be moved from a low orbit 2 towards Earth 1; or device 10 can be coupled behind satellite 20 "if it is necessary to move the satellite from a high orbit 3 towards a parking orbit 4.
[0143] [00143] The mechanical coupling of the device to the 20 ', 20 "satellite does not bind any dependency. In fact, the device according to the present invention can operate independently or semi-independently, or it can be controlled remotely.
[0144] [00144] The solid propulsion means of propulsion provide the necessary impulse to move the satellite 20 ', 20 "from its mission orbit, while a fixed or adjustable mechanical coupling platform (it is adjusted before launch, and subsequently passive or active due to the fact that it is under control of information feedback during the operation of the propeller) allows the direction of the propulsion generated by the propulsion means to be controlled, if necessary.
[0145] [00145] The device can operate in three different operating modes: remote controlled mode, independent mode and semi-independent mode. All three modes perform orbit withdrawal / orbit change operations in just a few steps:
[0146] [00146] - calculate, by means of remote control or by said device, a new desired orbit in space or a desired trajectory that leads to an impact area on Earth, respectively;
[0147] [00147] - calculate, by means of a remote control or by said device, the time to activate the device to move / remove to reach the new orbit or the desired trajectory based on the current orbit of the satellite;
[0148] [00148] - send from said remote control means to said device to move / remove a control signal or start the orbit withdrawal / orbit change procedure, through said device to move the satellite to the new orbit or calculated trajectory;
[0149] [00149] - ignite the means of propulsion of the device to move / remove at the calculated time when it receives the control signal.
[0150] [00150] In the preferred mode of operation, the device according to the present invention is controlled remotely. As shown in the flowchart in Figure 17, the control procedure required for device 10 to effect orbit withdrawal / orbit change is guided by means of remote control, which in the simplest case consists of a ground control station.
[0151] [00151] During the lifetime of satellite 20 ', device 10 remains in a WAITING state, until it receives a signal from the station on Earth 100, as the means of remote control. When it receives the signal, the onboard control means checks the status of the SIU (if any). If its status is START, this means that the system is armed and the procedure for igniting device 10 starts at the time established by the received signal and the satellite is safely removed from orbit / orbit and in a controlled manner, that is, is placed in a specifically calculated reentry orbit. If this is not the case, due to an unintended signal, for example, or if device 10 is not yet armed, then the ignition signal is ignored and device 10 returns to its WAIT state. Sending a safety signal from the remote control means to the means for receiving control signals from the device to move / remove allows the device to switch to an operational state before sending the control signal to move the satellite to the new orbit or calculated trajectory. Said signals can be in an encrypted mode for transmission security and prevent unauthorized intrusion and ignition. With reference to Figure 16, when it has been decided to remove satellite 20 'from orbit, a sequence of signals is sent to device 10 from the station on Earth 100 and / or from another vehicle in space. These signals are received by the means for receiving control signals operationally connected to the control means on board the device 10. Upon receipt of this signal sequence, the control means on board enable the SIU, which in turn enables the igniter 216 , which consequently ignites the propellant. In particular, a high voltage igniter is preferably used. The SIU, enabled by the activation signal, charges a capacitor unit until it reaches a high voltage state characteristic of the armed status. When the capacitor is suddenly discharged as a result of a subsequent independent ignition signal, the metallic conductive layer on the membrane is vaporized by the high voltage current and the plastic membrane is thrown at high speed against the main explosive layer, causing it to ignite .
[0152] [00152] Alternatively, device 10 can be fitted with a low voltage igniter. In this case, the igniter's electrical wire is heated and dissolved by a current flow (typically a few amps at low voltage for a few milliseconds) that represents the ignition signal. This ignites the pyrotechnic mixture contained in the igniter. As a result, the capsule releases the combustion products, sending them towards the main charge of the igniter and as a result ignites the solid propellant 212 in the propellant. This provides sufficient propulsion to position the satellite 20 'in a desired reentry orbit so that it will land on Earth 1 or disintegrate in the atmosphere. Alternatively, the propulsion of the device will be sufficient to relocate the device-satellite system to a previously established safe orbital space, such as a graveyard or parking orbit 4. In this preferred mode of operation, the allowable orbit withdrawal window is calculated by the station in question. Earth 100. Alternatively, it can be calculated by another vehicle located in space. This permissible orbit withdrawal window takes into account the position of the device-satellite system when the decision is made to proceed with the orbit withdrawal. The calculation of the permissible orbit withdrawal window also takes into account the other objects in space, such as other satellites or spacecraft, or debris, to prevent the device-satellite system from occupying a trajectory that could cause it to collide with any of these objects Therefore, the sequence of signals sent by station 100 to device 10 comprises at least one activation signal and subsequently at least one ignition signal, both sent once the decision to withdraw orbit has been made and when the permissible window is available. The orbital trajectory is calculated taking into account the orbital position and orientation of the device at the moment when the orbit withdrawal should take place, possibly exploring the four sensors, if available. The flowchart shown in Figure 18 extends the operations already illustrated in Figure 17. After receiving the signal and verifying that the SIU status is STOPPED, the device sends an alarm signal to the station on Terra 100 to alert that an attempt was made to ignite device 10, but the conditions necessary for this to occur have not been met, in the flowchart shown in Figure 19 the number of operations performed or required during the period in which the device is operational is further extended. To reduce any risk of SIU failure, if the status of the latter has been identified as STOPPED after the command to start ignition procedures has been received from the station on Terra 100, then a request is sent for further confirmation of the signal by station on Terra 100. Finally, device 10 can include a system to monitor its status with the means and sensors previously described, and the collected data can be sent to the station on Terra 100, as shown schematically in Figure 20, or at intervals regular or on request.
[0153] [00153] In a second mode of operation, the device is an autonomous device, that is, without any control from the ground station or from another vehicle in space, or from any means of remote control. Device 40 independently generates the ignition sequence by means of emitting control signals with its own control means on board after a predetermined time. This time interval usually corresponds to the useful life of the satellite 20 ', 20 "with which the device is associated and / or the duration of its mission. Alternatively, instead of being established in advance, said time may depend on limit values being exceeded and alarm signals are sent by the sensors, said limit values can be exceeded, for example, as a result of a severe defect, an impact (a mechanical shock identified by an accelerometer), or an imminent collision with another object in orbit (detected by a radar system at the ground station, or by proximity sensors on board the device, if any), in said second mode of operation, the control means on board will therefore independently generate at least the activation signal and at least the ignition signal. In this mode of operation, provision is also made for the ground station, or another space vehicle to be able to stop the ignition procedure by sending a PARA signal R to the device, which is received by the device's control signal receiving means. In addition, said ground station, or said other space vehicle, can reprogram the ignition sequence, sending the control signals as described for the preferred mode of operation described previously. For independently managed safe removal from orbit in accordance with this second mode of operation, device 40 must be able to establish its own position and orientation, and calculate a safe path for reentry into the atmosphere or towards an orbit that is, the device must be equipped with an orientation and positioning unit. In addition, the device must be able to access the permissible orbit withdrawal window in relation to other objects in space that can be found along its path, as previously explained.
[0154] [00154] A third mode of operation involves semi-independent operation of the device.
[0155] [00155] In this operating mode, some of the control signals are generated independently by the device with its own on-board control means. In particular, through the means for emitting control signals, the device generates signals useful for arming the device. Once armed, the ignition signal is sent from a station on Earth, or in space, or from other space vehicles. In particular, the ignition signals can also be sent by the satellite to be removed from orbit / changed orbit in the modalities with an operational connection for data exchange between the two.
[0156] [00156] According to a modality (not shown), before the device was turned on, if the satellite's behavior was out of control (for example, as a result of a defect in its behavior control system), the satellite positioning should reduce the angular tilt and rotation speeds to insignificant values and should align the direction of propulsion generated by the device with the orbital speed. If, even with active control of the satellite's behavior, the angles of inclination and rotation are sufficient to induce a misalignment between the propulsion generated by the device and the orbital speed, and the satellite is unable to modify its behavior to cancel the said misalignment, the satellite's positioning system can take action on the device itself to restore the satellite to the required behavior, enabling the device to generate propulsion in the required direction. The combustion of the cartridges installed in the system occurs in one or more cavities connected with two pairs of nozzles that allow rotation maneuvers around the tilt and rotation axes. Each cartridge cavity is separated from the nozzle pair by suitable valves. The valves can be of the solenoid type, open and closed by regulating the current in a coil, but any other type of valve could be used as long as it can be acted fast enough. When the single cartridge is lit, combustion takes place inside one of the cavities. With the valve closed, the gas pressure generated inside the cavity increases until the cartridge burns. When the valve opens, the gas under pressure tends to flow from the cavity into the nozzle. The diameters of the valve, the connection pipe leading to the nozzle and the neck on the nozzle are dimensioned to guarantee the expansion of the gas and its release through the opening in the nozzle at supersonic speed, in order to generate the necessary propulsion. The nozzles are located at an appropriate distance from the tilt and rotation axes so that the required torque is generated along each axis after the gas has been discharged through the single nozzle. The ignition of the single cartridge is governed by the control means on board. The opening and closing of the valves is managed by a control information feedback system that uses the values of behavior angles provided in real time by the behavior detection and / or control means. These means are operationally connected to the satellite's positioning system and work until the required behavior is achieved. Solid propellant cartridges are sized to generate gas at a sufficient pressure and for a sufficient time to allow the required behavior to be achieved, starting from any starting condition! in terms of angles and angular velocities. In addition, said satellite positioning system can be used to actively correct the misalignment of the impulse vector.
[0157] [00157] The device according to the present invention operates without any support from the satellite to be removed from the orbit with which it is associated. The device can communicate directly with another space vehicle or orbiting station, or control station on Earth. The station on Earth can send commands to arm and ignite the device, or request device-satellite system status data and (where available) satellite status data only. An advantage of the device according to the present invention lies in its ability to generate orbit withdrawal maneuvers in a few simple steps.
[0158] [00158] Therefore the device according to the present invention allows an appropriate disposal of the satellite with which it is coupled at the end of the latter's mission. In particular, said disposal is achieved by removing the satellite from its working orbit or removing the satellite from orbit towards Earth.
[0159] [00159] An advantage of the device according to the present invention relates to its ability to release the orbit previously occupied by a satellite, avoiding any interference with other satellites or space vehicles in the vicinity, which is an important advantage particularly for geostationary orbits and geosynchronous.
[0160] [00160] Another advantage of the device according to the present invention is that it produces a safe and controlled reentry of the satellite to be removed from orbit, avoiding collision damage and preventing impact with people or property on Earth, or with other objects orbiting in the neighborhood.
[0161] [00161] An orbit removal / orbit change device according to the present invention has the advantage of eliminating malfunctioning satellites with which it is associated, which would have no other way to re-enter the atmosphere, in this case of a satellite that is severely damaged during its lifetime, the device according to the present invention allows the damaged satellite to be moved into a parking orbit or to be brought back into the Earth's atmosphere, reducing the risk of it colliding with other satellites in operation.
[0162] [00162] The device according to the present invention advantageously eliminates the risk of satellites accidentally colliding with other known objects during their orbiting life cycles. If it is associated with a proximity sensor (for example, radar), the device allows collisions with any objects, even previously unknown, to be avoided in case the latter enters a certain safety distance from the satellite in question.
[0163] [00163] The device according to the present invention can be used advantageously to prevent intentional (or unintended) destructive activities or other dangerous operations, including the intentional destruction of a satellite, space vehicle or orbital by means of a deliberate collision, for example, or other activities that could increase the risk of collision with other objects in space.
[0164] [00164] The device can be used at any time to change the orbit of the satellite with which it is associated, or as a support device in the event of a failure of the satellite's propulsion means when it is launched. In case the engine of the final stage is defective, the device can be used as a backup propulsion system, allowing the satellite to reach its planned orbit or a reserve orbit to complete all or part of the planned mission.
[0165] [00165] Therefore the device according to the present invention allows the removal of unassisted orbit of a space vehicle simply by pre-programming the device itself, or by receiving at least one remote control directly from a station on Earth and without any limitation. The same control can also be sent by a vehicle in space, or even by the satellite to be removed from the orbit with which the device is mechanically coupled.
[0166] [00166] The device according to the present invention advantageously allows the satellite to be taken out of orbit to return to a predefined safe location on Earth, quite far from areas populated by humans or densely occupied by buildings.
[0167] [00167] Its construction is designed to ensure that it lasts longer than the mission of the satellite with which it is associated and, in extreme situations, it can be used as a backup system to increase the mission's useful time satellite by 20% to 80% in the event of a failure of the satellite's propulsion systems at the time of its launch.
[0168] [00168] Finally, the modular design of the device according to the present invention allows it to be adapted to the needs of orbiting the satellite with which it is associated.
[0169] [00169] In addition, the method according to the invention allows to implement the action of withdrawing orbit / orbiting a satellite by an associated autonomous device.
[0170] [00170] The method allows a simple and controlled withdrawal of orbit / orbit change from the satellite, safely calculating "a priori" an impact area on Earth or a new orbit in space.

权利要求:
Claims (13)
[0001]
Device (10) to be coupled to a space satellite (20 ', 20 ") to be orbited around a celestial body in a mission orbit before the launch of said space satellite to remove said space satellite from said mission orbit towards said celestial body or moving said space satellite from said mission orbit to a different space orbit when said space satellite reaches an end of life or when said space satellite becomes defective, said device operating in a mode independent or in a remote controlled mode applying remote control means, the mission orbit being an orbit intended for said space satellite for operations required during its stay in space, characterized by the fact that said device is independent and autonomous with respect to said space and independent satellite in relation to said means of remote control, in which said device comprises: control means on board said device (10), independent of the control means of said space satellite, comprising an electronic board including microcontrollers and at least one of an electronic interface and an electrical interface; means for receiving control signals from said remote control means or means for emitting control signals, operationally connected to said control means on board via at least one of said electronic interface and said electrical interface; wherein said control signals activate a movement / removal sequence; means of propulsion, independent of the propulsion of said space satellite, operationally connected to said means of control on board through at least one of said electronic interface and said electrical interface, in which said means of propulsion are activated by said means of control a board upon receipt of said control signals to remove said space satellite (20 ', 20 ") from said mission orbit to a predefined area of Earth or to a different predefined space orbit; means of supplying electric power to said device (10), independent of the means of supplying electric power to said space satellite, operationally connected to said control means on board by at least one of said electronic interface and said electrical interface; and pre-launch mechanical coupling means (310) coupling said device to said space satellite, providing an adjustable connection between said device and said space satellite; wherein said pre-launch mechanical coupling means is operationally connected to said control means on board said device via at least one of said electronic interface and said electrical interface; wherein said adjustable connection is configured to be adjusted to mitigate the misalignment of a drive vector of said means of propelling said device in relation to a center of gravity of a combination of said space satellite and said device, without employing a component of said space satellite, when said means of propulsion is activated; where said space satellite is a single satellite, and wherein the device is permanently coupled to said single satellite to be located external to said single satellite and protruding from it while remaining connected to said single satellite during operation in said drifting orbit and while removing said single satellite from said mission orbit towards said predefined area of Earth or said different predefined space orbit.
[0002]
Device according to claim 1, characterized by the fact that said device is coupled to said space satellite before launch, and said means of propulsion is activated by means of control on board said device upon receipt of said signals control to remove the space satellite from space to said predefined area of Earth.
[0003]
Device according to claim 1 or 2, characterized by the fact that it additionally comprises: means for detecting and / or controlling the behavior of said space satellite operationally connected to said control means on board and configured to carry out directional control of said device and said space satellite (20 ', 20 ") when said means of propulsion are activated.
[0004]
Device according to any one of claims 1 to 3, characterized in that said control means on board said device comprise at least one among means for emitting pre-programmed control signals and means for calculating said control signals. control to be sent to said means to receive control signals.
[0005]
Device according to any one of claims 1 to 4, characterized in that it additionally comprises: sensor means, operationally connected to said control means on board said device, for the detection of other objects within a defined security zone around at least one of said device and said space satellite (20 ', 20 ").
[0006]
Device according to any one of claims 1 to 5, characterized in that it additionally comprises: sensing means, operationally connected to said control means on board said device, for detecting impact damage in at least one of said device and / or on said space satellite (20 ', 20 ").
[0007]
Device according to any one of claims 1 to 6, characterized by the fact that it additionally comprises: sensing means, operationally connected to said control means on board the device, for detecting failures of said space satellite (20 ', 20 ").
[0008]
Device according to any one of claims 1 to 7, characterized by the fact that said pre-launch mechanical coupling means is configured to effect an adjustable connection between said space satellite and said device, wherein said means of pre-launch mechanical coupling is an adjustable platform located between the device and the space satellite.
[0009]
Device, according to claim 8, characterized by the fact that said means of mechanical pre-launch coupling comprise at least one of the active and passive control of said impulse vector that makes the adjustable connection between said space satellite (20 ', 20 ") and said device for effecting an alignment of said drive vector.
[0010]
Device according to any of claims 1 to 9, characterized in that said means of propulsion comprise at least one solid propellant engine and at least one igniter for said solid propellant engine, wherein said ignition device is operationally connected to said onboard control means to receive an ignition control signal.
[0011]
Device according to claim 10, characterized in that said means of propulsion comprise one or more independent cartridges or charges provided with solid propellant.
[0012]
Device according to any one of claims 1 to 11, characterized in that said means of propulsion comprise hybrid means of propulsion.
[0013]
Device according to any one of claims 1 to 11, characterized in that said means of propulsion comprise means of propelling gel propellant.

类似技术:

---

公开号 | 公开日 | 专利标题

---

BR112014001459B1|2021-01-12|device to move or remove artificial satellites

---

US8242422B2|2012-08-14|Modular divert and attitude control system

---

CN108367815B|2021-11-26|Propulsion system for small artificial satellite

---

EP2574557A2|2013-04-03|System and method for creating an artificial atmosphere for the removal of space debris

---

JP5505829B2|2014-05-28|Space debris reduction device

---

JP2022027800A|2022-02-14|Service satellites for providing orbital services with variable thruster control

---

JP2018527246A|2018-09-20|Operating system for Earth-orbiting satellite with electric thruster

---

Lin2003|Development of US air force intercontinental ballistic missile weapon systems

---

Schonenborg et al.2009|Solid propulsion de-orbiting and re-orbiting

---

JP3828072B2|2006-09-27|Thrust generating method and apparatus

---

Foelsche et al.2006|Flight results from a program to develop a freeflight atmospheric scramjet test technique

---

Toki et al.1997|Electric propulsion experiment | of a repetitively pulsed MPD thruster system onboard Space Flyer Unit |

---

WO2012125597A2|2012-09-20|Tether for spacecraft reaction control system

---

Thurman1995|Return to the red planet-An overview of the Mars Pathfinder mission

---

Strong et al.2010|Transport and use of a Centaur second stage in space

---

Schonenborg2004|Solid propellant de-orbiting for constellation satellites

---

JP2021511255A|2021-05-06|How to eject a payload from a spacecraft driven by continuous propulsion

---

PORT0|e~ c, i,-CCCI 22c

---

Lewis1971|NASA TM X-2188

---

Orloff et al.2006|Apollo 13: The seventh manned mission: in-flight abort 11–17 April 1970

---

PEKFOKMAN-CE0|FOE THE NIMBUS IT1 MISSION

---

Pelt2007|Launch

---

Janovsky2004|Mod-Protection and Mitigation Measures in Spacecraft Design

---

Orloff et al.2006|Apollo 6: The second test of the Saturn V 4 April 1968

---

STAGE0|Systems and scenarios

同族专利:

---

公开号 | 公开日

---

RU2014105836A|2015-08-27|

---

ES2728276T3|2019-10-23|

---

JP6349254B2|2018-06-27|

---

JP2014520724A|2014-08-25|

---

US20140158830A1|2014-06-12|

---

CN103732496B|2015-10-14|

---

AU2012285783B2|2017-02-02|

---

ITMI20111332A1|2013-01-19|

---

WO2013011073A1|2013-01-24|

---

CA2841893A1|2013-01-24|

---

BR112014001459A2|2017-02-21|

---

PT2734448T|2019-06-05|

---

CA2841893C|2017-11-21|

---

US9809327B2|2017-11-07|

---

CN103732496A|2014-04-16|

---

EP2734448B1|2019-02-06|

---

AU2012285783A1|2014-01-30|

---

KR20140066707A|2014-06-02|

---

UA116763C2|2018-05-10|

---

RU2600971C2|2016-10-27|

---

KR101920525B1|2018-11-20|

---

EP2734448A1|2014-05-28|

---

US20170081047A1|2017-03-23|

---

US9751643B2|2017-09-05|

引用文献:

---

公开号 | 申请日 | 公开日 | 申请人 | 专利标题

---

---

US3200587A|1961-09-15|1965-08-17|Gen Motors Corp|Rocket vehicle attitude control|

---

US3534686A|1968-10-04|1970-10-20|Nasa|Payload/burned-out motor case separation system|

---

US3589644A|1970-01-06|1971-06-29|Robert J Salkeld|Self-referencing system|

---

US3652042A|1970-12-28|1972-03-28|Trw Inc|Spacecraft for deploying multiple objects|

---

US3907225A|1973-12-17|1975-09-23|Tru Inc|Spacecraft for deploying objects into selected flight paths|

---

RU2026243C1|1987-06-10|1995-01-09|Шота Николаевич Хуцишвили|Method of delivery of celestial body to assignment planet and space transport facility for its realization|

---

JP2802130B2|1988-08-12|1998-09-24|日本電信電話株式会社|Orbit change method and orbit change device for artificial satellite|

---

US5120008A|1989-07-28|1992-06-09|Kumar Ramohalli|Orbital debris processor and method therefor|

---

US5056740A|1989-09-22|1991-10-15|The Johns Hopkins University|Over-the-horizon targeting system and method|

---

US4991799A|1990-02-16|1991-02-12|The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration|Orbital debris sweeper and method|

---

US5082211A|1990-10-31|1992-01-21|Teledyne Industries, Inc.|Method and apparatus for mitigating space debris|

---

JP3165455B2|1991-03-26|2001-05-14|三菱重工業株式会社|Method and apparatus for detecting collision of meteoroid or debris with spacecraft|

---

US5421540A|1992-08-26|1995-06-06|Ting; Paul C.|Method and apparatus for disposal/recovery of orbiting space debris|

---

JPH0789497A|1993-09-24|1995-04-04|Nec Corp|Automatic collision avoidance device and artificial satellite provided therewith|

---

US6102337A|1995-12-22|2000-08-15|Hughes Electronics Corporation|Spacecraft attitude control with gimbaled thrusters|

---

US5740985A|1996-09-16|1998-04-21|Scott; Harry|Low earth orbit payload launch system|

---

US6024328A|1996-12-18|2000-02-15|Hughes Electronics Corporation|Satellite control method using single species of bipropellant|

---

FR2814817B1|2000-10-04|2003-01-31|Sagem|DETECTION OF SPACE DEBRIS IN ORBIT FROM AN INSTRUMENT ON BOARD ON SATELLITE|

---

US7587325B1|2000-11-15|2009-09-08|Scott David R|Method of insuring against satellite launch failure|

---

US6629673B2|2001-11-28|2003-10-07|United Technologies Corporation|Adaptable solid-hybrid rocket for crew escape and orbital injection propulsion|

---

US6565043B1|2001-12-21|2003-05-20|The Boeing Company|Redundant system for satellite inclination control with electric thrusters|

---

DE10208724C1|2002-02-28|2003-09-25|Astrium Gmbh|Method for detecting damage to spacecraft caused by the impact of foreign objects|

---

US6655637B1|2002-06-24|2003-12-02|The Aerospace Corporation|Spacecraft for removal of space orbital debris|

---

US7093800B2|2002-09-26|2006-08-22|The Boeing Company|Method of operating a satellite for end-of-life maneuvers|

---

US7669802B2|2005-05-23|2010-03-02|Scapa Flow, Llc|Space based orbital kinetic energy weapon system|

---

US20070120020A1|2005-11-29|2007-05-31|Mei Technologies, Inc.|Small reusable payload delivery vehicle|

---

US7966805B2|2007-05-15|2011-06-28|Raytheon Company|Hydroxyl amine based staged combustion hybrid rocket motor|

---

JP2009204419A|2008-02-27|2009-09-10|Mitsubishi Electric Corp|Failure detector|

---

CN201165916Y|2008-03-24|2008-12-17|沈阳理工大学|Solid rocket engine|

---

JP2010069973A|2008-09-17|2010-04-02|Kyushu Institute Of Technology|Space debris removal method and its device|

---

US9233765B2|2011-06-16|2016-01-12|The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration|Multi-dimensional damage detection|

---

CN110010765A|2012-02-14|2019-07-12|内克斯特能源技术有限公司|Use the electronic device of organic small molecular semiconductor compound|US20140312160A1|2011-06-07|2014-10-23|Raytheon Company|Flight vehicles including scribed frangible seals and methods for the manufacture thereof|

---

US9187189B2|2012-10-12|2015-11-17|The Aerospace Corporation|System, apparatus, and method for active debris removal|

---

JP6472600B2|2014-03-18|2019-02-20|株式会社アストロスケール|Space device, debris removal system, and debris removal method|

---

JP6426364B2|2014-04-04|2018-11-21|株式会社アストロスケール|Debris removal device and debris removal system|

---

CN103963999B|2014-04-08|2016-02-10|北京控制工程研究所|A kind of thrust vectoring regulating mechanism locking and expanding unit|

---

RU2750349C2|2014-08-26|2021-06-28|Астроскейл Израэл Лтд.|Satellite docking system and method|

---

CN104309560A|2014-10-25|2015-01-28|贺州学院|Automotive solid propellant rocket protection system|

---

US9745083B2|2015-04-01|2017-08-29|Worldvu Satellites Limited|Method for thermal stabilization of a communications satellite|

---

US10254097B2|2015-04-15|2019-04-09|Raytheon Company|Shape memory alloy disc vent cover release|

---

KR101700539B1|2015-12-10|2017-01-26|한국항공우주연구원|Apparatus and method for satellite deorbiting|

---

CN105739511B|2016-01-31|2019-02-12|中国人民解放军国防科学技术大学|A kind of missing mark hovers asymptotic control method to the underactuated spacecraft of control|

---

EP3208199B1|2016-02-22|2020-01-29|Airbus Defence and Space GmbH|Device for controlling change of an orbit of a space object in space or removal of a space object from space and system and method for changing an orbit of a space object in space or for removing a space object from space|

---

KR101872612B1|2016-11-10|2018-06-28|한국항공우주연구원|Spacecraft for space debris removal|

---

US10625882B2|2017-03-06|2020-04-21|Effective Space Solutions Ltd.|Service satellite for providing in-orbit services using variable thruster control|

---

EP3376390B1|2017-03-17|2019-10-30|TTTech Auto AG|Fault tolerant method for controlling an autonomous controlled object|

---

EP3376330B1|2017-03-17|2020-11-04|TTTech Auto AG|Fault-tolerant method for detecting faults in an electronic system for controlling a controlled object|

---

JP6863797B2|2017-03-31|2021-04-21|三菱重工業株式会社|Artificial satellite and debris remover|

---

CA3069541A1|2017-07-21|2019-01-24|Northrop Grumman Innovation Systems, Inc.|Spacecraft servicing devices and related assemblies, systems, and methods|

---

CN108820256A|2018-06-23|2018-11-16|中国卫星海上测控部|A kind of spacecraft bias track design method|

---

EP3831724A4|2018-07-27|2021-07-28|IHI Aerospace Co., Ltd.|Emergency deorbit device and emergency deorbit method|

---

WO2020157807A1|2019-01-28|2020-08-06|三菱電機株式会社|Satellite constellation formation system, satellite constellation formation method, debris removal scheme, satellite constellation construction scheme, and ground equipment|

---

KR102049504B1|2019-09-02|2019-11-28|한국과학기술원|Method and apparatus for estimating orbital lifetime and survivability of orbiting objects|

---

RU2726340C1|2019-12-18|2020-07-13|Вадим Дмитриевич Зеленов|Device for docking of spacecrafts|

---

RU2726338C1|2019-12-20|2020-07-13|Вадим Дмитриевич Зеленов|Docking-mounting module|

---

WO2021193411A1|2020-03-26|2021-09-30|三菱電機株式会社|Satellite constellation forming system, mega constellation business device, satellite constellation, satellite, and ground facility|

法律状态:
2018-12-11| B06F| Objections, documents and/or translations needed after an examination request according [chapter 6.6 patent gazette]|

---

2020-01-28| B06U| Preliminary requirement: requests with searches performed by other patent offices: procedure suspended [chapter 6.21 patent gazette]|

---

2020-11-17| B09A| Decision: intention to grant [chapter 9.1 patent gazette]|

---

2021-01-12| B16A| Patent or certificate of addition of invention granted|Free format text: PRAZO DE VALIDADE: 20 (VINTE) ANOS CONTADOS A PARTIR DE 18/07/2012, OBSERVADAS AS CONDICOES LEGAIS. |

优先权:

---

申请号 | 申请日 | 专利标题

---

ITMI2011A001332|2011-07-18|

---

IT001332A|ITMI20111332A1|2011-07-18|2011-07-18|DEVICE FOR THE DEORBITATION OF ARTIFICIAL SATELLITES.|

---

PCT/EP2012/064123|WO2013011073A1|2011-07-18|2012-07-18|Device for moving or removing artificial satellites|

---