Method and system for multi-object space debris removal

Abstract

To provide a method and system for multi-object space debris removal that enable safer and / or more economical operation.SOLUTION: There is provided a method for rendezvous with an orbiting object according to an embodiment of the present invention. This method includes: launching a tug and a servicer into a client orbit; separating the servicer from the tug; and docking the servicer with a client. There is provided a system for rendezvous with an orbiting object according to another embodiment of the present invention. This system comprises a first spacecraft including a tug capable of towing a second spacecraft. The second spacecraft is a servicer configured to dock with a tumbling client orbiting object.SELECTED DRAWING: Figure 1

Description

[Technical field]

[0001]

[0001] This application claims priority to U.S. Provisional Patent Application No. 63 / 137,905, filed January 15, 2021, which is incorporated herein by reference in its entirety.

[0002] FIELD OF THE DISCLOSURE

[0002] Embodiments of the present invention relate to methods and systems for multi-object space debris removal. [Background technology]

[0003]

[0003] There are many aging satellites and space debris in low Earth orbit, and new satellites launched periodically increase the risk of collisions.

[0004]

[0004] Various methods, devices and systems have been proposed for deorbiting aging satellites and removing other space debris. For example, the following patent document 1 proposes a device for stabilizing and deorbiting a satellite, including a pair of coplanar masts. The masts each support at least one membrane forming an aerodynamic braking sail, and are fixed to the satellite along non-parallel axes and comprise a mass for generating a gravity gradient at an end opposite to the end fixed to the satellite. The masts form a fixed angle with a bisector between the masts, adapted to align the bisector with the satellite velocity vector at any altitude. [Prior art documents] [Patent documents]

[0005] [Patent Document 1] U.S. Patent No. 10,723,490B2 Summary of the Invention

[0006]

[0007] However, prior art methods and systems suffer from significant limitations, including the inability to safely and economically de-orbit satellites and debris under some conditions.

[0007]

[0008] In view of the above, aspects of the present invention provide methods and systems for multi-object space debris removal that enable safer and / or more economical operations.

[0008]

[0009] According to one aspect of the invention, a method for rendezvous with an orbital object is provided, the method including launching a tag and a servicing vehicle into a client orbit, separating the servicing vehicle from the tag, and docking the servicing vehicle with the client.

[0009]

[0010] According to another aspect of the invention, a system for rendezvous with an orbital object is provided, the system comprising a first spacecraft including a tug capable of towing a second spacecraft, the second spacecraft being a servicing vehicle configured to dock with a tumbling orbital client object. [Brief description of the drawings]

[0010] [Figure 1]

[0011] This indicates that the tag will be launched into orbit along with the service vehicle on a dedicated launch vehicle. [Diagram 2]

[0011] Shown is the separation of the service aircraft from the tag and contact of the service aircraft with a client that is desired to be removed from orbit. [Diagram 3]

[0011] A tag is shown docking with a service aircraft while the service aircraft maintains hold of a client. [Figure 4]

[0011] A service vehicle attached to a client is shown being brought into a lower orbit by a tug. [Diagram 5]

[0011] The launch of Reentry Shepherd is shown. [Figure 6]

[0011] Shown is the docking of the re-entry Shepard with the client, held in a stable state by a service vehicle attached to the tug. [Figure 7]

[0011] A separation of the service vehicle attached to the tug from the client attached to the re-entry shepard is shown. [Figure 8] 1 is a flowchart showing the steps in a method for rendezvous, capture, and disposal of an orbital object that is the subject of an embodiment. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0011]

[0012] DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS The present invention will now be described with reference to the accompanying drawings, in which like reference numerals denote like elements and have the same or similar configurations.

[0012]

[0013] As used herein, delta-V is a measure of the impulse per unit of spacecraft mass that is required to perform a maneuver, such as an in-space orbit maneuver. It is a scalar with units of velocity.

[0013]

[0014] Multi-object space debris removal missions require significant amounts of delta-V, as well as the ability to rendezvous with and capture large tumbling objects, and execute precisely targeted atmospheric re-entry trajectories to minimize the risk of harmful collisions with humans, animals, sensitive habitats and property located on the Earth's surface.

[0014]

[0015] In a solution utilizing a single monolithic spacecraft for rendezvous, capture, and disposal, these requirements are in conflict, as described below.

[0015]

[0016] Achieving high delta-V may require large fuel stores and / or highly efficient propulsion systems (high specific impulse, which is a measure of how efficiently fuel is utilized) such as solar electric propulsion. Solar propulsion requires a large deployment area of ​​solar cells with gimbal control to track the sun vector in order to generate sufficient power.

[0016]

[0017] Rendezvous with and capture of large tumbling objects may require a low mass, compact spacecraft with minimal deployed and / or protruding components to reduce rotational inertia and collision risk. This allows for an agile spacecraft that can counteract the velocity and rotation of a fast tumbling debris object while using less fuel. A smaller number of protruding components also reduces the combined rigid-body collision radius of the spacecraft, allowing for a wider safe approach angle to be achieved when capturing a debris object.

[0017]

[0018] To minimize the risk of a catastrophic accident on the ground, a precisely targeted re-entry trajectory may require that the debris object's remains only impact uninhabited areas on the Earth's surface during re-entry. Since the debris object has no ability to control its trajectory, this guidance must be performed by another spacecraft, which may consequently burn up together with the debris object in the atmosphere. In such a situation, the guidance spacecraft may not be reusable.

[0018]

[0019] In spacecraft design, achieving high delta-V necessarily entails large mass and size, while the ability to rendezvous with and capture large tumbling objects requires small mass and size. Both of these requirements increase cost, but a spacecraft executing a precisely targeted atmospheric re-entry trajectory may burn up with debris objects in the atmosphere, making the cost per use prohibitively high.

[0019]

[0020] The problem to be solved is that it is extremely difficult to design and build a single monolithic spacecraft that has high delta-V, is agile, and is cheap enough to be considered expendable. Meeting all three requirements is a unique challenge that has not been satisfactorily addressed thus far.

[0020]

[0021] Accordingly, one embodiment of the present invention provides a method for distributing the required functionality described above across multiple independent spacecraft that work together to accomplish a mission objective. Specifically, the embodiment utilizes a distributed architecture that includes two or more spacecraft.

[0021]

[0022] In FIG. 1 a tug 1 is launched (wholly or partly) into orbit together with a service vehicle 3 on a dedicated launch vehicle.

[0022]

[0023] In an embodiment, the tug 1 provides a significant amount of delta-V required for large orbital changes, rendezvous with multiple objects, and reorbiting large debris objects. The tug 1 can have a mass of several hundred kilograms up to a few tons, with a significant portion of the mass being spent on fuel (xenon or krypton) for electric propulsion. The tug 1 can also have multiple thrusters for thrust augmentation and redundancy, and a large tracking solar array with a wingspan of several meters to maximize solar power (which also augments thrust). The tug 1 may or may not include chemical propulsion, which may only be required for docking operations. Typical values ​​for the specific impulse of the tug 1 are 1200s-2000s, ideally at the upper end of this range. The thrust generated by the tug 1 can be 60-240mN depending on the solar power and number of thrusters. The tug 1 can also have a very capable attitude / momentum control system capable of controlling a multi-part combination that constitutes a much larger mass than the tag alone.

[0023]

[0024] In an embodiment, the Service Vehicle 3 is small, light, agile and designed to dock with both tumbling debris objects and stabilized operational spacecraft to provide servicing to clients (refueling or repair), detumbling, or other forms of movement or attitude control.

[0024]

[0025] The Service Vehicle 3 can be 200-400 kg, with mostly body mounted solar panels, plus various thruster arrays for maximum agility, typically with thrust on the order of a few Newtons. The bus size should be around one cubic meter, but the Service Vehicle 3 may also have a capture device, such as a robotic arm, a magnetic capture device, or some other type of interface. The propulsion of the Service Vehicle 3 is primarily chemical, although some embodiments of the Service Vehicle 3 may also have electric propulsion.

[0025]

[0026] The service aircraft 3 may be capable of operating in all attitude profiles when approaching the tumbling client 5. This makes it difficult to use directional antennas for ground communications, so an omnidirectional antenna may be used to close the communication link with the ground in any orientation. The drawback of an omnidirectional antenna is that the gain is low and therefore the achievable data rate is low. However, since the tag 1 does not need to perform a tumbling capture, it may have a directional antenna that allows for higher bandwidth communications. The tag 1 may be relatively close to the service aircraft 3 in space, acting as an intermediate communication relay, allowing the service aircraft 3 to close a high bandwidth data link to the tag 1 using an omnidirectional antenna, and then a high bandwidth link to the ground via a high gain antenna on the tag 1. In this way, the communication between the service aircraft and the ground can achieve a higher throughput than would be possible without the tag 1.

[0026]

[0027] There are many different expendable and recoverable launch vehicles in use that can be used to launch the tug 1 and service vehicle 3. Currently, government and private actors in the United States and other countries are developing a variety of launch vehicles, with each launch vehicle optimized for a particular mission.

[0027]

[0028] For example, the Atlas V and Delta IV rockets can be used to launch some embodiments. The Vulcan Centaur is an Atlas V successor in development that is expected to include some Delta IV technology and have similar capabilities. Additionally, several smaller launch vehicles are being developed to lower the overall cost of launching lighter spacecraft, but a broad commercial market for their use has yet to be developed.

[0028]

[0029] Other launch vehicles that may be useful in embodiments include the Falcon family, which consists of three launch vehicles, the Falcon 1, Falcon 9, and Falcon Heavy, developed by the American company SpaceX. Another privately developed launch vehicle that may be suitable for some embodiments is Virgin Orbit's LauncherOne, which is designed to launch a 300 kg payload into low Earth orbit at an altitude of 500 km.

[0029]

[0030] FIG. 2 illustrates the separation of the servicing vehicle 3 from the tug 1 and contact of the servicing vehicle 3 with a client 5 that is desired to be removed from orbit.

[0030]

[0031] In some embodiments, the client 5 is equipped with an optical marker before being launched into space. The optical marker reflects light of a predetermined wavelength band emitted from an illumination device of the service aircraft 3. An image of the optical marker reflecting the light of the predetermined wavelength band is captured by an image capture device of the service aircraft 3. The image is processed by an image processing device. As a result, the service aircraft 3 can estimate the attitude of the target object.

[0031]

[0032] In some embodiments, a capture plate may be attached to the client 5. The capture plate may be attached to a portion of the client 5 where the capture device interfaces with the interface component of the service machine 3. The interface component includes an adhesive. By attaching the capture plate to the client 5, the interface component of the service machine 3 may be easily and reliably interfaced to the capture plate of the client 5, regardless of the structure or material of the outer surface of the client 5. The capture plate may include a guide structure for defining the interface position between the client 5 and the service machine 3. This may improve the accuracy of the interface position between the client 5 and the service machine 3.

[0032]

[0033] In some embodiments, a set of capture arms extends from the service aircraft 3 and is operative to engage the client 5. The set of capture arms includes at least one umbilical, a first end of the umbilical attached to the service aircraft 3 and a second end of the umbilical attached with a service aircraft umbilical end connector, the service aircraft umbilical end connector configured to form a connection with the client umbilical connector of the client 5. The set of capture arms also includes a manipulator arm, a first end of the manipulator arm coupled to the service aircraft 3 and a second end of the manipulator arm configured to connect to and manipulate the second end of the umbilical. The set of capture arms also includes a processor operative to control the manipulator arm. The manipulator arm manipulates the second end of the umbilical to form a connection between the service aircraft umbilical end connector and the client umbilical connector.

[0033]

[0034] In some embodiments, the service machine 3 may include various structures for attachment to the client 5. Rigid systems of attachment hardware may include brackets, clamps, bolts, and screws. Non-rigid systems may include a combination of tethers or elastic elements. A "lock and key" type element may be attached to the client 5 to facilitate the attachment process. For example, various forms of connectors may be attached to the client 5 and then connected to a complementary portion of the tag 1.

[0034]

[0035] Clients 5 can vary widely in size and shape. They can be small debris pieces less than 10 cm in diameter or large satellites weighing several tons. In low Earth orbit (800-2,000 km), where most satellites operate, there are millions of debris pieces that are too small to track. About 35,000 space debris pieces larger than 10 cm in diameter have been tracked individually. About 9,000 satellites have been launched into low Earth orbit to date, of which only about 2,000 are operational, while the majority of the rest are still in orbit. About 60% of the satellites in orbit are non-functional and are unwanted debris.

[0035]

[0036] About 24% of the debris objects tracked individually are aging satellites, and about 18% are used upper stages and mission-related objects, such as launch adapters and lens covers. Because impacts occur at high speeds, even space debris pieces with a diameter of only 1 cm can harm satellites. Space debris pieces larger than 1 cm are believed to be capable of penetrating the exterior walls of existing satellites and spacecraft, causing catastrophic failure if the exterior walls are not reinforced with additional layers.

[0036]

[0037] A typical collision with space debris occurs with a closure velocity of 10 km / s, which is equivalent to 36,000 km / hour. The object needs to be accelerated to a speed of about 7 km / s to remain in low Earth orbit. The speed of an object in space is determined by the laws of physics and the gravitational field of the body around which it orbits. Objects can be found in many different orbits around the Earth, some moving in the opposite direction to others.

[0037]

[0038] FIG. 3 shows the tag 1 docking with the service vehicle 3 while the service vehicle 3 maintains possession of the client 5 .

[0038]

[0039] Spacecraft docking is the joining of two spacecraft. This connection can be temporary or partially permanent, such as for space station modules. Docking specifically refers to the joining of two separate free-flying spacecraft. For an orbital rendezvous to occur, both spacecraft must be in the same orbit, and the positions and rotations of the spacecraft within the orbit must match. Before the tug 1 and servicing vehicle 3 can dock, the translational and angular velocities of the tug 1 and servicing vehicle 3 must be approximately equal.

[0039]

[0040] To achieve this objective, a rotation suppression device for suppressing the rotation of the client 5 can be provided. The rotation suppression device may include a main body, a shaft extending outward from the main body and configured to rotate about a first rotation axis, a rotating part attached to an end of the shaft opposite the main body and configured to rotate together with the shaft about a second rotation axis, a capture part fixed to the rotating part and configured to capture a target, a braking part provided on the main body and configured to suppress the rotation of the shaft, and a main body rotation suppression part configured to suppress the rotation of the main body that occurs when the braking part operates. The main body rotation suppression part may be, for example, a reaction wheel provided inside the main body. With this configuration, the capture part captures the client 5, and the rotating part fixed to the capture part and the shaft attached to the rotating part can rotate together with the client 5. Then, the braking part gradually suppresses the rotation of the shaft, and the main body rotation suppression part suppresses the rotation of the main body that occurs when the braking part operates. That is, the capture unit rotates integrally with the client 5, for example, and this rotation is suppressed by the braking unit, thereby suppressing the rotation of the main body that occurs when the braking unit operates (that is, while maintaining the position of the main body, the angular momentum of the client 5 is transferred to the main body rotation suppressing unit and absorbed there). As a result, it is possible to effectively suppress the relatively large rotational motion of the client 5 (having a large angular momentum).

[0040]

[0041] The docking and berthing systems on the tug 1, service vehicle 3, and client 5 can be asexual or sexual, depending on the design of the attachments. Early systems for mating spacecraft were all sexual docking system designs. A sexual design is a form of gender mating, where each component of the mating interface has a unique design (e.g., a "male" or "female" shape) and plays a specific role in the docking process. These roles are not interchangeable, and two spacecraft with identical connection attachments cannot be mated. In contrast, an asexual docking (and asexual berthing) scheme uses identical interfaces on all spacecraft so equipped, and these can be used to connect spacecraft.

[0041]

[0042] In an embodiment, the client 5 may not include a propulsion means, for example if the client 5 has no remaining fuel stores. The service vehicle 3 may guide the movement of the client 5 to facilitate the formation of a complex with the tug 1. The tug 1 and the service vehicle 3 may be equipped with a computerized control system. This system may use data collected by the LIDAR ranging sensor and the infrared and visible light camera in guiding the movement of the client 5.

[0042]

[0043] FIG. 4 shows a service vehicle 3 attached to a client 5 being brought into a downward orbit by a tug 1.

[0043]

[0044] If the tag 1 generates thrust in the opposite direction to the current direction of motion, the tag 1, the service vehicle 3, and the client 5 can fall into a low-energy elliptical transfer orbit. The tag 1 can then generate thrust to insert the tag 1, the service vehicle 3, and the client 5 into a corresponding low-energy circular orbit. Conversely, if the tag 1 generates thrust in the same direction as the current direction of motion, the tag 1, the service vehicle 3, and the client 5 can rise into a high-energy elliptical transfer orbit. The tag 1 can then generate thrust to insert the tag 1, the service vehicle 3, and the client 5 into a corresponding high-energy circular orbit.

[0044]

[0045] In an embodiment, the Hohmann transfer orbit can provide one way to move the tug 1, service vehicle 3, and client 5 into a high or low orbit. The Hohmann transfer orbit is tangent to both the current and desired orbits of the tug 1, service vehicle 3, and client 5. To initiate the orbit change, the tug 1, service vehicle 3, and client 5 are propelled in the direction of motion to accelerate the tug 1, service vehicle 3, and client 5 along the elliptical Hohmann transfer orbit if moving to a high orbit, and are propelled in the opposite direction to decelerate the tug 1, service vehicle 3, and client 5 if a transfer to a low orbit is desired. Once the tug 1, service vehicle 3, and client 5 reach the tangent point between the Hohmann transfer orbit and the desired orbit, the tug 1, service vehicle 3, and client 5 can accelerate and change their motion so that they are moving within the desired orbit.

[0045]

[0046] As used herein, low orbit means an orbit having a low altitude as measured from the center of the Earth, and high orbit means an orbit having a high altitude as measured from the center of the Earth. Orbits can be described and classified according to a number of classification systems. A low-Earth orbit (LEO) is an orbit centered on the Earth with an altitude of 2,000 km (about 1,200 miles) or less. The high orbit class includes intermediate Earth orbits, sometimes called intermediate circular orbits (ICOs), and even higher geostationary orbits. Objects in geostationary orbits move with the same angular velocity as the Earth rotates, so that they remain at a single point in the sky as seen from the Earth's surface. A high-Earth orbit is a geocentric orbit with an altitude generally higher than a geosynchronous orbit (35,786 kilometers (22,236 miles)).

[0046]

[0047] Objects in low Earth orbit (less than about 500 km) are affected by atmospheric drag. As atmospheric drag reduces the kinetic energy of an object in orbit, the object will lose altitude accordingly until it re-enters the atmosphere. Thus, atmospheric drag moves objects out of orbit without human intervention. Objects in low Earth orbit are affected more by atmospheric drag and as a result decay faster. The decay lifetime of a space object depends on its altitude, solar activity level, and mass to cross-sectional area. Objects with a large mass to area ratio decay slower as they are less affected by drag. High solar activity increases atmospheric density and drag in low Earth orbit. For objects in orbit at relatively low altitudes, atmospheric drag is strong enough that it may cause a re-entry before the intended end of the mission if maneuvers to raise the orbit are not performed from time to time. On average, if client 5 is initially in a 300 km high orbit, client 5 has a decay lifetime of only a few months. If client 5 is in a 500 km orbit, its lifetime is about 10 years. If Client 5 is at an altitude of 1000 km, it can remain in orbit for thousands of years without the intervention of any external forces acting on it.

[0047]

[0048] In some embodiments, rather than moving the client 5 to a low Earth orbit, the complex formed by the tug 1 and the servicing vehicle 3 moves the client 5 to a high Earth orbit known as a graveyard orbit, junk orbit, or disposal orbit. For satellites in geosynchronous and geosynchronous orbits, the graveyard orbit is several hundred kilometers above the operational orbit.

[0048]

[0049] In Figure 5, reentry Shepard 7 is launched.

[0049]

[0050] In embodiments, the Re-entry Shepherd 7 can be designed to provide guidance, control, and / or thrust to a debris object for the specific purpose of targeted atmospheric re-entry. Thus, the Re-entry Shepherd 7 can be equipped with chemical or electric propulsion systems. In some embodiments, the Re-entry Shepherd 7 can be equipped with attachments that can be used on a client to increase atmospheric drag. Drag devices increase the cross-sectional area of ​​the satellite, allowing the atmosphere to slow the satellite down and lower its altitude, although these devices may not guarantee a specific re-entry location, such as an uninhabited area of ​​the South Pacific Ocean.

[0050]

[0051] The mass of the Reentry Shepherd 7 can range from a few hundred kilograms to 2-3 tons. For example, a debris object of about 3 tons would require a smaller Reentry Shepherd 7 than an 8 tons debris object. The Reentry Shepherd 7 design is expected to be based on the kick stage of a launch vehicle, such as Rocket Lab's Photon or Soyuz's Fregat. A typical launch vehicle's final stage has a large thrust (hundreds of Newtons) that is useful for the reentry burn, since a large amount of impulse must be applied quickly. More than one burn is not possible during the final reentry process, since the final maneuver must usually lower the perigee from about 180 km to 50 km to be successful. Normally, with a perigee at these altitudes, control would be lost due to aerodynamic disturbances, but the mission would fail since no immediate reentry would occur. The large thrust helps ensure that a single burn can apply enough impulse to change the perigee sufficiently.

[0051]

[0052] The Re-entry Shepherd 7 typically has a mass between the tug and the servicing vehicle. The Re-entry Shepherd 7 does not require the agility of the servicing vehicle 3, nor the long life and fuel efficiency of the tug 1. The Re-entry Shepherd 7 only requires a large impulse and thrust. The Re-entry Shepherd 7 can be thought of as an extended kick stage of the launch vehicle. In some embodiments, all three spacecraft, i.e., the tug 1, the servicing vehicle 3, and the Re-entry Shepherd 7, can be launched together, while in other embodiments, they are all launched separately. The architecture is flexible. The Re-entry Shepherd 7 may also be launched several months after the tug and servicing vehicle, waiting for the optimal time to lower the client 5 from a high altitude to a low altitude. This allows some flexibility in the development, deployment, and operation timelines, since deploying multiple spacecraft simultaneously is laborious. The configurable architecture also allows flexibility in the choice of launch vehicle.

[0052]

[0053] FIG. 6 shows the docking of a re-entry Shepard 7 with a Client 5 held steady by a Service Vehicle 3 attached to a Tug 1.

[0053]

[0054] The client 5 may be held stable by magnetic, mechanical, or other forces. The service vehicle 3 may, for example, include permanent magnets to facilitate holding after successful docking. Cryogenically cooled superconducting wires may be used to generate magnetic forces in the service vehicle 3. The client 5 may have an internal magnet intended to orient the client 5 with the Earth's magnetic field. This internal magnet may be utilized by the service vehicle 3 to attract or repel the client 5 or to shift the orbit of the client 5. The service vehicle 3 may also be equipped with a robotic arm or other mechanical means to hold the client 5.

[0054]

[0055] Docking between the servicing vehicle 3 and the client 5 can be difficult, but after the servicing vehicle 3 takes control, subsequent docking operations are simplified. The servicing vehicle 3 can grasp a small portion of the client 5, leaving a large area for another spacecraft, such as the reentry shepherd 7, to approach and grasp the same interface. This can be accomplished by extending the robotic arm of the servicing vehicle 3 to avoid another approaching spacecraft, as shown in FIG. 6. In this operation, the servicing vehicle 3 can also provide optical guidance to the reentry shepherd 7 by means of a retroreflector or LED positioned on the robotic arm or end effector of the servicing vehicle 3. The ability of the servicing vehicle 3 to provide this guidance to the reentry shepherd 7 is a significant advantage, since the client 5 may not have optical fiducials to assist in docking.

[0055]

[0056] Figure 7 shows the separation of a service vehicle 3 attached to a tug 1 from a client 5 attached to a reentry shepherd 7. The tug 1 and service vehicle 3 accelerate in the direction of the next client object. The reentry shepherd 7 uses drag to reduce the orbit of the client 5 until it is low enough for a direct reentry burn.

[0056]

[0057] In an ideal case, the client 5 would evaporate completely during its passage through the atmosphere. As the client 5 loses altitude, heat from friction with the gases in the Earth's atmosphere burns up the client 5. A rough rule of thumb is that the temperature in Kelvin around the client 5 is equal to the entry velocity in meters per second. Thus, at an orbital re-entry velocity of 7800 m / s, temperatures can be as high as 7800 K. However, in an embodiment, some parts of the client 5 may survive this high temperature and reach the Earth's surface, risking damage to people and property on the ground.

[0057]

[0058] Some components of client 5 (especially those made of heat-resistant materials such as titanium) may survive atmospheric re-entry and fall to the ground. In that case, it may be desirable for a re-entry shepherd 7 to take client 5 to a position that will cause re-entry over an uninhabited area of ​​the Earth. In an embodiment, re-entry shepherd 7 may target a specific re-entry point to reduce the chance of a catastrophic incident to people or property on the Earth's surface. If client 5 has steering capabilities and has fuel remaining at the end of its life, client 5 may be positioned for re-entry over a large area of ​​ocean. If client 5 does not have steering capabilities and has no fuel remaining, a new re-entry shepherd must be used for each client 5 requiring direct atmospheric re-entry to provide the necessary guidance and control of client 5 during re-entry.

[0058]

[0059] If the client 5 does not require guidance and control after being placed on the desired re-entry trajectory across the Earth, the re-entry shepherd 7 may be reusable. In some embodiments, the re-entry shepherd 7 may provide a re-entry burn at apogee and then detach from the client 5 to quickly gain altitude before reaching perigee. This maneuvering process may require only a small portion of the orbital period and may need to be automated. The re-entry shepherd 7 may use chemical propulsion for this maneuver, most likely biprop with a thrust of at least 400N. If the re-entry shepherd 7 detaches from the client 5, the re-entry burn will require high precision to ensure it hits the target re-entry trajectory in the Pacific Ocean. After recovery, the re-entry shepherd 7 will likely remain in low orbit to await the next client. The re-entry shepherd 7 may be carried to different altitudes by the tug 1.

[0059]

[0060] The Tug 1 and Service Vehicle 3 are designed to be reused multiple times for multiple clients, and may be used with or without the Re-entry Shepherd 7, as some clients may not require direct atmospheric re-entry. This reusability of both of these highly capable platforms helps reduce disposal costs per object.

[0060]

[0061] 8 is a flow chart illustrating steps in a computer-implemented method for designing a spacecraft useful for orbital object rendezvous, capture, and disposal that is the subject of one embodiment. In embodiments, many variations of the method are possible, and in fact the order of steps may vary from embodiment to embodiment.

[0061]

[0062] One embodiment is a computer-implemented method for calculating spacecraft parameters, the method including inputting one or more target client attributes via one or more devices (101) and calculating, with one or more processors, desired properties of one or more tags and / or one or more service vehicles according to a previously defined ranking (103). The devices can be various computer hardware such as a mouse, a keyboard, or various memory types.

[0062]

[0063] The desired properties may be various design and aviation parameters related to the predicted performance of the tug and / or servicing vehicle. Thus, the desired properties output may provide useful information for the manufacturing process of the tug and servicing vehicle, including materials used, component and assembly sizes, and component and assembly shapes. Additionally, re-entry shepherds may be constructed according to the desired properties output from the embodiments.

[0063]

[0064] The desired properties can be output in a variety of data formats, including a variety of spreadsheets and word processing programs, as well as a variety of file formats that can be used in computer-aided design.

[0064]

[0065] The previously defined rankings may be responsive to received data regarding a client, and may assign different amounts of weight to different design and aviation parameters depending on the mission objectives associated with rendezvous with and abandonment of that client. The previously defined rankings may be input via an I / O device such as a keyboard, mouse, or memory.

[0065]

[0066] Other embodiments may include calculating a target weight of fuel for one or more of the tug, servicing vehicle, and re-entry shepherd (105). Other embodiments may include calculating a target time for one or more of launch, rendezvous, and disposal for one or more of the tug, servicing vehicle, re-entry shepherd, and client (107).

[0066]

[0067] Also, computer hardware, such as a processor, useful for implementing the embodiments can be included to calculate 109 the desired properties of the re-entry shepherds according to the rankings previously defined. In embodiments, it may not be necessary to calculate the desired properties of one or more re-entry shepherds simultaneously. In other words, step 109 may be performed simultaneously with steps 101 and 103, or step 109 may be performed at a later time, followed by steps 111 and 113.

[0067]

[0068] The above-described embodiments are provided for the purpose of facilitating understanding of the present invention, and are not intended to limit the interpretation of the present invention. The elements of the embodiments and their arrangements, materials, conditions, shapes, sizes, etc. are not limited to the examples described, and may be changed as appropriate. Furthermore, the components described in the embodiments may be partially replaced or combined.

Claims

1. 1. A method for rendezvous with an orbital object, comprising: launching the tug and the service vehicle into a client orbit; Separating the service machine from the tag; docking the service aircraft with a client; A method comprising:

2. The method of claim 1 , further comprising detumbling the client.

3. The method of claim 1 or 2, further comprising docking the tag with the client.

4. The method of claim 1 , wherein the tag docks with the client while the service aircraft maintains hold.

5. 5. The method of claim 1, wherein the tag docks with the service machine while the service machine maintains possession of the client.

6. The method of claim 1 , further comprising moving the service vehicle and client to a desired trajectory by the tag.

7. The method of claim 1 , further comprising docking a re-entry shepherd with the client.

8. 8. The method of claim 1, further comprising positioning the client on a desired re-entry trajectory across the Earth by the re-entry shepherd.

9. The method of claim 7 or 8, further comprising guiding and controlling the client during atmospheric re-entry by the re-entry shepherd.

10. The method of claim 7 , further comprising: detaching the re-entry shepherd from the client.

11. The method of claim 1 , further comprising refueling the client.

12. The method of claim 1 , further comprising repairing the client.

13. The method of claim 1 , further comprising modifying the trajectory of the client.

14. The method of claim 1 , further comprising coupling one or more components to the client.

15. said service aircraft closing a high bandwidth data link to said tag using an omnidirectional antenna; the tag acts as an intermediate communications relay, enabling a high bandwidth link to ground via a directional antenna on the tag; 15. The method of claim 1 , further comprising:

16. 1. A system for rendezvous with an orbital object, comprising: a first spacecraft including a tug capable of towing a second spacecraft; The system, wherein the second spacecraft is a servicing vehicle configured to dock with a tumbling on-orbit client object.

17. The service aircraft is equipped with an omnidirectional antenna, The system of claim 16 , wherein the tag is equipped with a directional antenna.

18. 18. The system of claim 16 or 17, further comprising a re-entry shepherd configured to guide the on-orbit client object on a desired trajectory.

19. 19. The system of claim 16, 17 or 18, wherein the re-entry shepherd is reusable.

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