Patent ID: 12234043

It should be understood that the proportions and dimensions (either relative or absolute) of the various features and elements (and collections and groupings thereof) and the boundaries, separations, and positional relationships presented there between, are provided in the accompanying figures merely to facilitate an understanding of the various embodiments described herein and, accordingly, may not necessarily be presented or illustrated to scale, and are not intended to indicate any preference or requirement for an illustrated embodiment to the exclusion of embodiments described with reference thereto.

DETAILED DESCRIPTION

Reference will now be made in detail to representative embodiments. The following descriptions are not intended to limit the embodiments to one preferred embodiment. To the contrary, it is intended to cover alternatives, modifications, and equivalents as can be included within the spirit and scope of the described embodiments as defined, for example, by the appended claims.

The disclosed devices, systems, and methods of use will be described with reference toFIGS.1-9. Generally, a multi-vehicle system and associated method for removal of multi-object space debris is disclosed.FIGS.1-7provide representations of portions of a particular configuration of one embodiment of a multi-vehicle system for removal of multi-object space debris, each portion generally corresponding to a step described in the method of use of an embodiment of a multi-vehicle system for removal of multi-object space debris is described inFIG.8.

Although the disclosed devices, systems, and methods of use will be described relative to the removal of multi-object space debris, the devices, systems, and methods of use have other applications. For example, the method and/or devices may be used for the capture, servicing, and/or release of space objects. Other applications or uses are possible.

With attention toFIGS.1-7, a system for removal of multi-object space debris (also referred to as “multi-object space debris removal system,” or simply as “system”) comprises three space vehicles: a space tug vehicle1also referred to as “tug,”, a servicer vehicle3also referred to as “servicer,” and a reentry shepherd vehicle7also referred to as “reentry shepherd” or “shepherd.” The three vehicles operate in concert to capture and remove a targeted spacecraft, such as a targeted space debris object, referred to as a client spacecraft5or simply as “client.” The system and method distributes the aforementioned required capabilities across multiple, independent spacecraft which work together to achieve the mission goals. The method of use of the multi-vehicle system for removal of multi-object space debris may be referred to as “space debris removal method” or simply “method.”

The operation and features of the multi-vehicle system for removal of multi-object space debris will be described with reference to particular configurations of the system depicted inFIGS.1-7.

With attention toFIG.1, tug1is launched together with servicer3into orbit on a (wholly or partially) dedicated launch vehicle.

In some embodiments, tug1provides substantial amounts of delta V required for large orbit changes, rendezvousing with multiple objects, and re-orbiting massive debris objects. Tug1may be configured from several hundred kilograms up to a couple of tons in mass, devoting substantial mass fraction to fuel for electric propulsion (xenon or krypton). Tug1may also have multiple thrusters for increased thrust and redundancy, and a large tracking solar array with a wingspan of several meters to maximize solar power generation, which also increases thrust. Tug1may or may not include chemical propulsion which may be needed for docking operations only. Typical values for the specific impulse of tug1is 1200 s-2000 s, in a preferred configuration in the upper end of that range. Thrust generated by tug1may be between 60-240 mN depending on solar power generation and number of thrusters. Tug1may also have a very capable attitude/momentum control system that allows it to control a joint stack of multiple bodies that together comprise much more mass than the tug alone.

In some embodiments, servicer3is small, light, agile, and designed for docking with both tumbling debris objects and stabilized operational spacecraft to either provide services (such as refueling or repair), detumble, or provide other forms of movement or attitude control to the client.

Servicer3may be between 200-400 kg, with mostly body mounted solar panels, and a diverse array of thrusters for maximum agility, typically around a few Newtons in thrust. The bus size may be around a meter cubed, but servicer3may also have a device for capture, be it a robotic arm, magnetic capture device, or some other type of interface. The propulsion of servicer3may be primarily chemical, but some embodiments of servicer3could also have electric propulsion.

Servicer3may be capable of operating in all attitude profiles as it approaches tumbling client5. This makes a directional antenna for ground communication difficult to use, and as such, an omni-directional antenna may be used to close the communications link with the ground in any orientation. The drawback of an omni-directional antenna is that the gain is low, so achievable data rates are also low. Tug1, however, does not need to do a tumbling capture, so it may have a directional antenna which allows for higher bandwidth communications. Since tug1may be in space in relatively close proximity to servicer3, tug1may act as an intermediary comms relay, which allows servicer3to close a high-bandwidth data link to tug1using an omni-directional antenna, and then the high-bandwidth link to the ground may be enabled via high gain antenna on tug1. That way, servicer-to-ground communications may be higher throughput than it could be without tug1.

There are many different expendable and recoverable launch vehicles in use that may be used to launch tug1and servicer3. Currently, government and private parties in the United States and in other countries have developed a variety of launch vehicles, with each of the vehicles optimized for particular missions.

For example, the Atlas V and Delta IV rockets may be used to launch some embodiments. Vulcan Centaur is a successor to the Atlas V under development that includes some Delta IV technology and is expected to have similar capacity. In addition, certain smaller launch vehicles have been developed to launch lighter spacecraft at a lower overall cost, though they have not found a wide commercial market for their use.

Other launch vehicles potentially useful in embodiments include the Falcon family, which consists of three launch vehicles-Falcon 1, Falcon 9, and Falcon Heavy-built by the U.S. corporation SpaceX. Another privately developed launch vehicle which might be suitable in some embodiments is Virgin Orbit's LauncherOne, designed to launch a 300-kg payload to a 500-km altitude low-Earth orbit.

With attention toFIG.2, the separation of servicer3from tug1and contact by servicer3with client5desired to be deorbited is illustrated.

In some embodiments, client5is equipped an optical marker before being launched into space. The optical marker reflects light of a predetermined wavelength band radiated from a lighting device of servicer3; an image of the optical marker which has reflected the light of the predetermined wavelength band is acquired by an image acquisition device of servicer3; and the image is processed by an image processing device. As a result, servicer3can estimate the attitude of the target object.

In some embodiments, a capturing plate can be attached to client5. The capturing plate can be attached to a part of client5where a capturing device bonds with a bonding component of servicer3, with the bonding component including adhesive. The capturing plate is attached to client5, whereby the bonding component of servicer3can be bonded to the capturing plate of client5easily and reliably regardless of the structure or the material of the outer surface of client5. The capturing plate may include a guide structure to define the bonding position of client5with servicer3. This can improve the accuracy in bonding position of client5with servicer3.

The above-mentioned optical marker and/or capturing plate may utilize aspects and/or features as described in one or more of the following patent matters, all of Astroscale Japan Inc or associated companies, and each incorporated by reference in entirety for all purposes: U.S. Pat. Appl. Publ. No. 2018/0229865 filed 7 Apr. 2017 and entitled “Capturing System, Space Vehicle and Plate;” U.S. Pat. No. 10,882,643 issued 5 Jan. 2021 and entitled “Capturing Plate, On-Orbit Device and Method for Capturing;” U.S. Pat. Publ. Appl. No. 2019/0367192 filed 1 Feb. 2018 and entitled “Capturing System, Aerospace Vehicle, and Plate-Like Body;” and U.S. Pat. No. 10,309,798 issued 4 Jun. 2019 and entitled “Navigation System, Aerospace Vehicle and Optical Marker.”

In some embodiments, a set of capture arms extend from servicer3and operate to engage client5, including: at least one servicer umbilical with a servicer umbilical first end attached to the servicer3and a servicer umbilical second end fitted with a servicer umbilical end connector, the servicer umbilical end connector configured to form a connection with a client umbilical connector of client5; a manipulator arm with a manipulator arm first end coupled to servicer3and a manipulator arm second end configured to attach to and maneuver the servicer umbilical second end; and a processor operating to control the manipulator arm; wherein: the manipulator arm maneuvers the servicer umbilical second end to form a connection between the servicer umbilical end connector and the client umbilical connector. The engagement or use of an umbilical may, in some embodiments, be as described in U.S. Pat. Appl. Publ. No. 2021/0086923 of Astroscale Israel Ltd. filed 21 Sep. 2020 and entitled “In-Orbit Spacecraft Servicing Through Umbilical Connectors,” incorporated by reference in entirety for all purposes.

The docking of the servicer3with the client5via a set of capture arms may be facilitated by geometries or configurations of the capture arms, such as the ends of the capture arms, as described in U.S. Pat. No. 10,611,504 issued 7 Apr. 2020 and entitled “Docking System and Method for Satellites;” U.S. Pat. No. 10,625,882 issued 21 Apr. 2020 and entitled “Service Satellite for Providing In-Orbit Services Using Variable Thruster Control;” and U.S. Pat. No. 11,117,683 issued 14 Sep. 2021 and entitled “Service Satellite for Providing In-Orbit Services Using Variable Thruster Control;” each of Astroscale Israel Ltd. and incorporated by reference in entirety for all purposes.

In some embodiments, servicer3may include a variety of structures for providing an attachment to client5. A rigid system of attachment hardware could include brackets, clamps, bolts, and screws. A non-rigid system could include a combination of tethers or elastic elements. “Lock and key” style elements may be attached to client5to facilitate the attachment process. For example, various forms of connectors could be attached to client5and then connected to the complementary part of tug1.

Client5may vary across a wide range of sizes and shapes. Client5may be a small piece of debris less than 10 cm in diameter or client5can be a large satellite weighing many tons. There are millions of pieces of debris in Low Earth Orbit (between 800-2,000 km) where most satellites operate that are too small to be tracked. Around 35,000 pieces of space debris over 10 cm in diameter have been individually tracked. Around 9,000 satellites have been launched into Low Earth Orbit of which only about 2,000 are operational, with the majority of the remainder still orbiting. Around 60% of satellites in orbit are non-functional and undesirable debris.

About 24% of the individually tracked debris objects are obsolescent satellites, and about 18% are spent upper stages and mission-related objects such as launch adapters and lens covers. Even bits of space debris only 1 cm in diameter can harm a satellite because collision would occur at a high velocity. It is believed that a fragment of space debris larger than 1 centimeter is capable of penetrating the outer walls of existing satellites and spacecraft, which can cause catastrophic failure if the walls are not reinforced with additional layers.

A typical impact with space debris occurs at a closing velocity of 10 km/s, equal to 36,000 km/hr. An object needs to be accelerated to a velocity of around 7 km/s to stay in low Earth orbit. Velocities of objects in space are determined by the laws of physics and the gravitational field of the body around which the objects orbit. Objects can be found in many different orbits around the Earth, some travelling in the opposite direction to others.

With attention toFIG.3, a docking by tug1with servicer3while servicer3maintains hold of client5is illustrated.

Docking of spacecraft is the joining of two space vehicles. This connection can be temporary, or partially permanent such as for space station modules. Docking specifically refers to joining of two separate free-flying space vehicles. For orbital rendezvous to occur, both spacecraft must be in the same orbit, and the position and the rotation of the spacecraft in the orbit must be matched. Prior to docking tug1and servicer3, the translational and angular speed of tug1and servicer3generally must be made equal.

To achieve the object, in one embodiment, a rotation suppressing device may be supplied for suppressing rotation of client5, and can include: a body; a shaft extending outward from the body and configured to rotate about a first rotation axis; a rotation part attached to an end of the shaft opposite to the body and configured to rotate about a second rotation axis together with the shaft; a capture part fixed to the rotation part and configured to capture the target; a braking part provided in the body and configured to suppress rotation of the shaft; and a body rotation suppressing part configured to suppress rotation of the body occurring when the braking part operates. The body rotation suppressing part may be a reaction wheel provided inside the body, for example. With this configuration, the capture part can capture client5, and the rotation part fixed to the capture part and the shaft attached to the rotation part can rotate together with client5. Then, the braking part gradually suppresses rotation of the shaft, and the body rotation suppressing part suppresses rotation of the body occurring when the braking part operates. That is, the capture part, for example, rotates integrally with client5, and this rotation is suppressed by the braking part, thereby suppressing rotation of the body occurring when the braking part operates (i.e., an angular momentum of client5is moved to the body rotation suppressing part and absorbed therein with the position of the body maintained). As a result, rotational motion of client5that is relatively large (has a large angular momentum) can be effectively suppressed.

Docking and berthing systems on tug1, servicer3, and client5may be either androgynous or non-androgynous, depending on the design of attachment pieces. Early systems for conjoining spacecraft were all non-androgynous docking system designs. Non-androgynous designs are a form of gender mating where each component of an interface to be joined has a unique design (e.g., “male” shape or “female” shape) and a specific role to play in the docking process. The roles cannot be reversed and two spacecraft with identical connecting attachments cannot be joined. An androgynous docking (and androgynous berthing) scheme by contrast uses an identical interface on all spacecraft so equipped, which can be used to connect the vehicles.

In one embodiment, the servicer3employs techniques, components, and/or elements described in U.S. Pat. Appl. No. 63/137,680 filed 14 Jan. 2021 and entitled “Method for Capture of Tumbling Space Debris” to Astroscale Holdings Inc., incorporated by reference in entirety for all purposes.

In one embodiment, the client5is a tumbling spacecraft and the positioning of the servicer3relative to the client5is as described in ELSA-D: An In-orbit End-of-Life Demonstration Mission, Blackerby et al, IAC-18, Sep. 14, 2018, incorporated by reference in entirety for all purposes.

In some embodiments, client5may lack any means of propulsion, such as if client5has no remaining fuel stores. Servicer3can guide the motion of client5to facilitate formation of a complex with tug1. Tug1and servicer3may be equipped with computerized control systems, which can use data gathered by scanning LIDAR ranging sensors and infrared and visible cameras in guiding motion of client5.

With attention toFIG.4, tug1is illustrated bringing servicer3attached to client5to a lower orbit.

If tug1generates thrust in the opposite direction to its current direction of motion, tug1, servicer3, and client5can drop into a lower-energy elliptical transfer orbit. Tug1can then generate thrust to insert tug1, servicer3, and client5into a corresponding lower-energy circular orbit. Conversely, if tug1generates thrust in the same direction as its current direction of motion, tug1, servicer3, and client5can rise into a higher-energy elliptical transfer orbit. Tug1can then generate thrust to insert tug1, servicer3, and client5into a corresponding higher-energy circular orbit.

In embodiments a Hohmann transfer orbit may provide one manner of moving tug1, servicer3, and client5into a higher or lower orbit. A Hohmann transfer orbit is tangent to both the current orbit of tug1, servicer3, and client5and a desired orbit. A change of orbit is initiated by propulsion of tug1, servicer3, and client5in the direction of motion to accelerate tug1, servicer3, and client5along the elliptical Hohmann transfer orbit if going to a higher orbit, or by propulsion against the direction of motion to decelerate tug1, servicer3, and client5if transition to a lower orbit is desired. When tug1, servicer3, and client5arrive at the point of tangency between the Hohmann transfer orbit and the desired orbit, tug1, servicer3, and client5can accelerate to change its motion so that tug1, servicer3, and client5travel in the desired orbit.

As used herein, lower orbit means an orbit with lower altitude when measured from the center of the Earth, and higher orbit means are orbit with higher altitude when 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 Earth-centered orbit with an altitude of 2,000 km (about 1,200 mi) or less. Higher orbit classes include medium Earth orbits, sometimes called intermediate circular orbits (ICO), and further above, geostationary orbits. An object in geostationary orbit moves at the same angular velocity as the rotation of the Earth, and so from the vantage point of the surface of Earth remains at a single point in the sky. A high Earth orbit is a geocentric orbit with an altitude entirely above that of a geosynchronous orbit (35,786 kilometers (22,236 mi)).

Objects in low altitude orbits (below about 500 km) are affected by atmospheric drag. Atmospheric drag reduces the kinetic energy of orbiting objects, so that the objects correspondingly decrease in altitude until re-entering the atmosphere. Atmospheric drag can thus remove objects from orbit without human intervention. Objects in lower orbits will be affected to a greater degree by atmospheric drag and will accordingly decay faster. The decay lifetime of a space object depends on its altitude, the level of solar activity, and the object's mass to cross-sectional area. Objects with a large mass to area ratio will decay more slowly as they are less affected by drag. High solar activity increases the density of the atmosphere and atmospheric drag in low Earth orbits. For objects orbiting at relatively low altitudes, atmospheric drag can be sufficiently strong to cause a re-entry before the intended end of mission if orbit raising maneuvers are not executed from time to time. On the average, if client5is in an initial 300 km high orbit, client5will have a decay lifetime of only a few months. If client5is in a 500 km orbit, its lifetime will be around 10 years, and if client5is at 1000 km altitude client5can stay in orbit for thousands of years without the intervention of outside forces acting upon client5.

In some embodiments, rather than moving client5to a lower orbit, the complex formed by tug1and servicer3move client5to a higher orbit known as a graveyard, junk, or disposal orbit. For satellites in geostationary orbit and geosynchronous orbits, the graveyard orbit is a few hundred kilometers above the operational orbit.

With attention toFIG.5, reentry shepherd7is launched.

In some embodiments, reentry shepherd7may be designed to provide guidance, control, and/or thrust to a debris object for the specific purpose of targeted atmospheric reentry. Reentry shepherd7may accordingly be equipped with chemical or electric propulsion systems. In some embodiments, reentry shepherd7may be outfitted with attachments which can be applied to a client in order to increase atmospheric drag. Drag devices can increase the cross-sectional area of a satellite to cause the atmosphere to slow the satellite and lower the satellite's altitude, but these devices may not assure a particular re-entry location such as the South Pacific Ocean Uninhabited Area.

Reentry shepherd7may be anywhere from a few hundred kg to a couple of tons in mass. Debris objects around 3 tons will require a smaller reentry shepherd than debris objects that are 8 tons, for example. The design of reentry shepherd7is assumed to be based on the kick stages of launch vehicles, such as the Photon by Rocket Lab, or Fregat by Soyuz. The last stage of typical launch vehicles has significant thrust (hundreds of Newtons) which helps with reentry burns because a large amount of impulse must be imparted quickly. The final reentry process cannot take more than one burn because the final maneuver usually must lower the perigee from ˜180 km to 50 km to be successful. Having a perigee between those altitudes will usually result in a loss of control of due to aerodynamic disturbances, but not an immediate atmospheric reentry and thus a mission failure. A large thruster helps make sure a single burn will impart enough impulse to sufficiently change the perigee.

Reentry shepherd7may typically have a mass in between the tug and servicer. Reentry shepherd7does not need the agility of servicer3and does not need the longevity and fuel efficiency of tug1. Reentry shepherd7just needs substantial impulse and thrust. Reentry shepherd7is, in one embodiment, an augmented kick stage of launch vehicles. It is possible that that all three vehicles—tug1, servicer3, and reentry shepherd7—could be launched together in some embodiments, while in others they can all be launched separately. The architecture is flexible. Reentry shepherd7may also be launched several months after the tug and servicer, waiting for an optimal time to bring client5down from a higher altitude to lower altitude. This allows some flexibility in development, deployment, and operation timelines, as deploying multiple spacecraft at the same time can be burdensome. Having the architecture be configurable in different ways also allows for flexibility in launch vehicle selection.

With attention toFIG.6, a reentry shepherd7docking with client5, which is held stable by servicer3attached to tug1, is illustrated.

Client5may be held stable by means of magnetic, mechanical, or other forces. Servicer3can for example include permanent magnets to facilitate holding after successful docking. Magnetic force can be generated in servicer3using superconducting wires cooled to cryogenic temperatures. Client5may have on-board magnets intended to adjust the orientation of client5using Earth's magnetic field which can be utilized by servicer3for the purpose of attracting or repelling client5or to shift the orbit of client5. Servicer3may also be equipped with robotic arms or other mechanical means for holding client5.

Docking servicer3with client5may be difficult, but after servicer3has control, subsequent docking operations will be simplified. Servicer3may grasp a small section of client5, and leave significant area for another spacecraft, such as reentry shepherd7, to approach and grasp that same interface. This may be achieved by extending the robotic arm of servicer3which allows servicer3to get out of the way of another approaching spacecraft, as depicted inFIG.6. In this operation, servicer3may even provide optical guidance to reentry shepherd7via retroreflectors or LEDs positioned on a robotic arm or end-effector of servicer3. As client5may not have optical fiducials which aid docking, the ability of servicer3to provide this guidance to reentry shepherd7is a significant advantage.

With attention toFIG.7, separation of servicer3attached to tug1from client5attached to reentry shepherd7is illustrated. Tug1and servicer3accelerate in the direction of a next client object. Reentry shepherd7uses drag to lower the orbit of client5until the orbit is low enough for direct reentry burn.

In an ideal case, client5is vaporized entirely during transit through the atmosphere. Heat from the friction of the gasses in Earth's atmosphere burns up client5as client's altitude decreases. An approximate rule-of-thumb is that the air temperature in Kelvin around client5is equal to the entry speed in meters per second. Thus, at an orbital reentry velocity of 7800 m/s, the temperature may be as high as 7800 K. However, in embodiments, some pieces of client5may withstand even this high temperature to reach the surface of Earth and pose a risk of damage to persons and property on the ground.

Some components of client5(especially parts which are made of heat resistant materials like titanium) may survive atmospheric re-entry and fall down to Earth. In that case, it may be desirable for reentry shepherd7to escort client5to a position which will result in reentry over uninhabited areas of the Earth. In embodiments, reentry shepherd7may target a particular atmospheric entry point, in order to reduce the likelihood of casualties to persons or property on the surface of the earth. If client5has maneuvering capability and still has remaining fuel at the end of its life, client5may be positioned so that client5reenters over a large area of ocean. If client5does not have maneuvering capability and remaining fuel, a new reentry shepherd must be used for each client5requiring direct atmospheric reentry, to provide the guidance and control of client5required during reentry.

Reentry shepherd7may be reusable in the case where client5does not need guidance and control once being placed on the desired reentry trajectory intersecting earth. In some embodiments, reentry shepherd7may provide a reentry burn at apogee, then undock from client5, and quickly raise its altitude before it reaches perigee. This maneuver process may have to take a fraction of an orbit period and may have to be automated. Reentry shepherd7may use chemical propulsion in this maneuver (in one preferred embodiment, biprop) with a thrust of at least 400N. If reentry shepherd7undocks from client5, significant precision would be required on the reentry burn to ensure the target corridor in the Pacific Ocean is hit. After recovery, reentry shepherd7would likely remain at a low orbit, waiting for the next client. Reentry shepherd7may potentially be carried by tug1to a different altitude.

Tug1and servicer3are designed to be reused multiple times for multiple clients and may be used with or without the reentry shepherd7, as some clients may not require direct atmospheric reentry. This reusability of both of these highly capable platforms helps to reduce the cost of disposal per object.

FIG.8provides a flow diagram of one method of use800of embodiment of a multi-vehicle system for removal of multi-object space debris described above with respect toFIGS.1-7.

Generally, the method800starts at step804and ends at step848. Any of the steps, functions, and operations discussed herein can be performed continuously and automatically. In some embodiments, one or more of the steps of the method of use800, to include steps of the method800, may comprise computer control, use of computer processors, and/or some level of automation. However, in some embodiments some of the steps or parts of some of the steps are performed in concert with or exclusively by human intervention. For example, a final approval (by a human) may be required to initiate a docking, a separation, or transfer of custody from one space vehicle to another.

The steps are notionally followed in increasing numerical sequence, although, in some embodiments, some steps may be omitted, some steps added, and the steps may follow other than increasing numerical order. A user may interact or perform one or more of the described steps be using a display/GUI. The phrase “user interface” or “UI”, and the phrase “graphical user interface” or “GUI”, means a computer-based display that allows interaction with a user with aid of images or graphics.

After starting at step804, the method800proceeds to step808. At step808, a multi-vehicle system for removal of multi-object space debris is provided, such as described above with respect toFIGS.1-7. After completing step808, the method800proceeds to step812.

At step812, a combined tug1and servicer3vehicle is formed or is presented. In one embodiment, as described above with respect toFIG.1, the combined tug-servicer vehicle is launched as a common or combined or attached or coupled unit. After completing step812, the method800proceeds to step816.

At step816, as described above with respect toFIG.2, the servicer3separates from the combined tug-servicer vehicle. After completing step816, the method800proceeds to step820.

At step820, the servicer3attaches or docks with the client5, the client5operating in a first orbit, as described above with respect toFIG.2. In one embodiment, the client5is a tumbling client. In one embodiment, the servicer3docks with or attaches to the client5while the client5is tumbling. In one embodiment, the servicer3detumbles to tumbling client5. After completing step820, the method800proceeds to step824.

At step824, the tug1attaches to the servicer3, the servicer3in turn still attached or docked to the client5, as described inFIG.3above. After completing step824, the method800proceeds to step828.

At step828, the tug1brings the combined tug-servicer-client vehicle to a second orbit, as described above with respect toFIG.4. After completing step828, the method800proceeds to step832.

At step832, the reentry shepherd is attached to the combined tug-servicer-client, as described above with respect toFIG.6. After completing step832, the method800proceeds to step836.

At step836, the tug vehicle and servicer vehicle are separated from the combined tug-servicer-client-shepherd vehicle to form each of a tug-servicer vehicle and a client-shepherd vehicle, as described above with respect toFIG.7. After completing step836, the method800proceeds to step840.

At step840, the client-shepherd vehicle executes atmospheric reentry, as described above with respect toFIG.7. After completing step840, the method800proceeds to step844.

At step844, the tug-servicer vehicle maneuvers to rendezvous with a second client, as described above with respect toFIG.7. After completing step844, the method800proceeds to step848and ends.

With attention toFIG.9, a flow diagram illustrating steps in a computer-implemented method of designing spacecraft useful for rendezvous, capture, and disposal of an orbiting object that is the subject of one embodiment. In embodiments, many variations on this method are possible, and in fact it is likely that the order of steps may be changed from one embodiment to another. Any of the steps, functions, and operations discussed herein can be performed continuously and automatically. In some embodiments, one or more of the steps of the method of use900, to include steps of the method900, may comprise computer control, use of computer processors, and/or some level of automation.

One embodiment is computer-implemented method of calculating parameters of a spacecraft comprising: inputting, via one or more devices101, attributes of one or more target clients; and calculating, at one or more processors103, desired qualities of one or more tugs and/or one or more servicers according to a ranking previously defined. Devices can be various computer hardware such as a mouse, a keyboard, or various memory types.

Desired qualities can be various design and aeronautical parameters associated with the projected performance of a tug and/or a servicer. Thus, output of desired qualities can provide useful information about manufacturing processes for tugs and servicers, including materials used, sizes of components and assemblies, and shapes of components and assemblies. As well, reentry shepherds can be constructed according to desired qualities output from embodiments.

Desired qualities can be output in various data formats, including various spreadsheet and various word processing programs, as well as various file formats which can be used in Computer-Aided Design.

Previously defined rankings can weight, in response to data received about a client, different design and aeronautical parameters in different amounts according to mission objectives associated with rendezvous and disposal with that client. Previously defined rankings can be input via an I/O device such as a keyboard, mouse, or memory.

Other embodiments can comprise calculating105target weights of a fuel for one or more of the tug, servicer, and reentry shepherd. Still other embodiments can comprise calculating107a target time of one or more of launch, rendezvous and disposal of among one or more of tugs, servicers, reentry shepherds and clients.

Computer hardware such as a processor useful for implementing an embodiment may also be equipped for calculating109desired qualities of a reentry shepherd according to a ranking previously defined. In embodiments desired qualities of one or more reentry shepherds can but need not be calculated at the same time. In other words, step109can occur simultaneously with steps101and103, or step109can occur later, and then can be followed by step111and step113.

The embodiments described above are given for the purpose of facilitating the understanding of the present invention and does not intend to limit the interpretation of the present invention. The respective elements and their arrangements, materials, conditions, shapes, sizes, or the like of the embodiment are not limited to the illustrated examples but may be appropriately changed. Further, the constituents described in the embodiment may be partially replaced or combined together.

The exemplary systems and methods of this disclosure have been described in relation to systems and methods of use of providing multi-object space debris removal, such as removal of one or more client satellites. Other uses or applications to the disclosed systems and methods are possible, such as servicing of satellites. Also, to avoid unnecessarily obscuring the present disclosure, the preceding description omits a number of known structures and devices, and other application and embodiments. This omission is not to be construed as a limitation of the scopes of the claims. Specific details are set forth to provide an understanding of the present disclosure. It should however be appreciated that the present disclosure may be practiced in a variety of ways beyond the specific detail set forth herein.

Furthermore, it should be appreciated that the various links connecting the elements can be wired or wireless links, or any combination thereof, or any other known or later developed element(s) that is capable of supplying and/or communicating data to and from the connected elements. These wired or wireless links can also be secure links and may be capable of communicating encrypted information. Transmission media used as links, for example, can be any suitable carrier for electrical signals, including coaxial cables, copper wire and fiber optics, and may take the form of acoustic or light waves, such as those generated during radio-wave and infra-red data communications.

Also, while the methods have been discussed and illustrated in relation to a particular sequence of events, it should be appreciated that changes, additions, and omissions to this sequence can occur without materially affecting the operation of the disclosed embodiments, configuration, and aspects.

A number of variations and modifications of the disclosure can be used. It would be possible to provide for some features of the disclosure without providing others.

Although the present disclosure describes components and functions implemented in the aspects, embodiments, and/or configurations with reference to particular standards and protocols, the aspects, embodiments, and/or configurations are not limited to such standards and protocols. Other similar standards and protocols not mentioned herein are in existence and are considered to be included in the present disclosure. Moreover, the standards and protocols mentioned herein, and other similar standards and protocols not mentioned herein are periodically superseded by faster or more effective equivalents having essentially the same functions. Such replacement standards and protocols having the same functions are considered equivalents included in the present disclosure.

The present disclosure, in various aspects, embodiments, and/or configurations, includes components, methods, processes, systems and/or apparatus substantially as depicted and described herein, including various aspects, embodiments, configurations embodiments, sub-combinations, and/or subsets thereof. Those of skill in the art will understand how to make and use the disclosed aspects, embodiments, and/or configurations after understanding the present disclosure. The present disclosure, in various aspects, embodiments, and/or configurations, includes providing devices and processes in the absence of items not depicted and/or described herein or in various aspects, embodiments, and/or configurations hereof, including in the absence of such items as may have been used in previous devices or processes, e.g., for improving performance, achieving ease and\or reducing cost of implementation.

The foregoing discussion has been presented for purposes of illustration and description. The foregoing is not intended to limit the disclosure to the form or forms disclosed herein. In the foregoing Detailed Description for example, various features of the disclosure are grouped together in one or more aspects, embodiments, and/or configurations for the purpose of streamlining the disclosure. The features of the aspects, embodiments, and/or configurations of the disclosure may be combined in alternate aspects, embodiments, and/or configurations other than those discussed above. This method of disclosure is not to be interpreted as reflecting an intention that the claims require more features than are expressly recited in each claim. Rather, as the following claims reflect, inventive aspects lie in less than all features of a single foregoing disclosed aspect, embodiment, and/or configuration. Thus, the following claims are hereby incorporated into this Detailed Description, with each claim standing on its own as a separate preferred embodiment of the disclosure.

Moreover, though the description has included description of one or more aspects, embodiments, and/or configurations and certain variations and modifications, other variations, combinations, and modifications are within the scope of the disclosure, e.g., as may be within the skill and knowledge of those in the art, after understanding the present disclosure. It is intended to obtain rights which include alternative aspects, embodiments, and/or configurations to the extent permitted, including alternate, interchangeable and/or equivalent structures, functions, ranges or steps to those claimed, whether or not such alternate, interchangeable and/or equivalent structures, functions, ranges or steps are disclosed herein, and without intending to publicly dedicate any patentable subject matter.