Method and System for Removing Multi-Object Space Debris

Abstract

According to one aspect of the invention, a method is provided for rendezvous with an orbital object. The method includes launching a tug and a servicing vehicle into a client orbit, separating the servicing vehicle from the tug, and docking the servicing vehicle with the client. According to another aspect of the invention, a system is provided for rendezvous with an orbital object. The system includes a first spacecraft including a tug capable of towing a second spacecraft. The second spacecraft is a servicing vehicle configured to dock with a tumbling orbital client object.

Description

Technical Field

[0001]

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

[0002]

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

Background Art

[0003]

[0003] Aging satellites and space debris are numerous in low Earth orbit, and newly launched satellites increase the risk of collision.

[0004]

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

Prior Art Documents

Patent Documents

[0005]

Patent Document 1

Summary of the Invention

[0006]

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

[0007]

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

[0008]

[0009] According to one aspect of the present invention, a method for rendezvousing with an object in orbit is provided. The method includes launching a tag and a service machine into a client orbit, separating the service machine from the tag, and docking the service machine with the client.

[0009]

[0010] According to another aspect of the present invention, a system for rendezvousing with an object in orbit is provided. The system includes a first spacecraft including a tag capable of towing a second spacecraft. The second spacecraft is a service machine configured to dock with a tumbling in-orbit client object.

Brief Description of the Drawings

[0010]

Figure 1

[0011] Shows launching a tag together with a service machine into orbit on a dedicated launcher.

Figure 2

[0011] Shows separation of the service machine from the tag and contact of the service machine with a client for which removal from orbit is desired.

Figure 3

[0011] Shows the tag docking with the service machine while the service machine maintains hold of the client.

Figure 4

[0011] Shows taking the service machine attached to the client by the tag to a lower orbit.

Figure 5

[0011] Shows the launch of a reentry shepherd.

Figure 6

[0011] Illustrates the docking of a reentry shepherd with a client maintained in a stable state by a service machine attached to a tag.

Figure 7

[0011] Illustrates the separation of a service machine attached to a tag from a client attached to a reentry shepherd.

Figure 8

[0011] A flowchart showing each step in a method for rendezvous, capture, and disposal of an object in orbit, which is the subject of an embodiment.

DETAILED DESCRIPTION OF THE INVENTION

[0011]

[0012] Embodiments of the present invention will be described with reference to the accompanying drawings. It should be noted that in each figure, components denoted by the same reference numerals have the same or similar configurations.

[0012]

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

[0013]

[0014] In a multi-object space debris removal mission, a significant amount of delta V is required, as well as the ability to rendezvous with and capture a large tumbling object, and to execute an accurately targeted atmospheric reentry trajectory in order to minimize the risk of harmful collisions with humans, animals, vulnerable habitats, and property located on the Earth's surface.

[0014]

[0015] In solutions that utilize a single monolithic spacecraft for rendezvous, capture, and disposal, these requirements are conflicting, as described below.

[0015]

[0016] To achieve a high delta-V, a large amount of fuel storage and / or a very efficient propulsion system such as solar electric propulsion (high specific impulse. Specific impulse is a measure of how efficiently fuel is utilized) may be required. In the case of solar electric propulsion, a large deployment area of solar panels using gimbal control is necessary to track the solar vector for sufficient power generation.

[0016]

[0017] For rendezvous with and capture of a large tumbling object, a low-mass and compact spacecraft with minimized deployed components and / or protruding components may be required to reduce the inertia of rotation and the risk of collision. This enables a sensitive spacecraft that can quickly cancel the speed and rotation with respect to a rapidly tumbling debris object while reducing the fuel used. Also, a smaller number of protruding components reduces the combined rigid body collision radius of the spacecraft and allows for a wider safety approach angle required when capturing a debris object.

[0017]

[0018] To minimize the risk of a major accident on the ground, it may be necessary for the remnants of the debris object to collide only with unpopulated areas on the Earth's surface during atmospheric reentry in order to execute an accurately targeted atmospheric reentry trajectory. Since the debris object has no ability to control its trajectory, this guidance must be performed by another spacecraft, which may consequently burn up with the debris object in the atmosphere. In such a situation, the reusable nature of this guiding spacecraft may not be available.

[0018]

[0019] In the design of a spacecraft, achieving a high delta-V necessarily involves a large mass and size, while the function of rendezvousing with and capturing a large tumbling object requires a small mass and size. Both of these requirements increase costs, but a spacecraft that executes an accurately targeted atmospheric reentry trajectory may burn up with the debris object in the atmosphere, and the cost per use may be 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 a high delta V, is sensitive, and is inexpensive enough to be considered disposable. Meeting all three requirements is a unique problem that has not been satisfactorily addressed to date.

[0020]

[0021] Accordingly, one embodiment of the present invention provides a method of distributing the required functions described above among a plurality of independent spacecraft. These spacecraft cooperate to achieve the mission objectives. Specifically, the embodiment utilizes a distributed architecture that includes two or more vehicles.

[0021]

[0022] In FIG. 1, tag 1 is launched into orbit (either wholly or in part) with service vehicle 3 by a dedicated launch vehicle.

[0022]

[0023] In an embodiment, tag 1 provides a significant amount of delta V required for large orbital changes, rendezvous with multiple objects, and reorbiting of a large number of debris objects. Tag 1 has a mass ranging from several hundred kilograms up to a maximum of 2 - 3 tons, and a significant portion of its mass can be devoted to fuel (xenon or krypton) for electric propulsion. Also, tag 1 can have multiple thrusters for thrust augmentation and redundancy, and a large tracking solar array with a wingspan of several meters to maximize solar power generation (which also increases thrust). Tag 1 may or may not include chemical propulsion that may be required only for docking operations. A typical value for the specific impulse of tag 1 is 1200s - 2000s, ideally at the upper limit of this range. The thrust generated by tag 1 can be 60 - 240 mN depending on solar power generation and the number of thrusters. Also, tag 1 can have a very capable attitude / momentum control system that can control a combination of multiple parts that make up a mass much larger than the tag itself.

[0023]

[0024] In an embodiment, the service vehicle 3 is small, lightweight, and agile, and is designed to dock with both a tumbling debris object and a stabilized and operating spacecraft to provide services (fuel replenishment or repair) to a client, perform debris removal, or perform other forms of movement or attitude control.

[0024]

[0025] The service vehicle 3 can be 200 - 400 kg, mostly body-mounted solar panels, and further includes various thruster arrays for maximum agility typically with a thrust of approximately several newtons. The bus size must be about 1 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 mainly chemical, but some embodiments of the service vehicle 3 may also have electric propulsion.

[0025]

[0026] When approaching the tumbling client 5, the service vehicle 3 may be operable in all attitude profiles. This makes it difficult to use a directional antenna for ground communication, so an omnidirectional antenna can be used to establish a communication link with the ground in any direction. The disadvantage of an omnidirectional antenna is that the achievable data rate is low because the gain is low. However, since tag 1 does not need to perform a tumbling capture, it may have a directional antenna that enables higher bandwidth communication. Since tag 1 can be relatively close to the service vehicle 3 in space, it can act as an intermediate communication relay, enabling the service vehicle 3 to establish a high-bandwidth data link to tag 1 using an omnidirectional antenna, and then enabling a high-bandwidth link to the ground via the high-gain antenna on tag 1. In this way, the communication between the service vehicle and the ground can achieve a higher throughput than is possible without tag 1.

[0026]

[0027] Many different disposable launchers and reusable launchers that can be used for the launch of Tag 1 and Service Machine 3 are being used. Currently, government and private stakeholders in the United States and other countries are developing a variety of launchers, and each launcher is optimized for a specific mission.

[0027]

[0028] For example, Atlas V and Delta IV rockets can be used to launch some embodiments. Vulcan Centaur is a developing successor to Atlas V that is expected to include some Delta IV technology and have a similar capacity. Additionally, several small launchers have been developed to launch lightweight spacecraft at a lower overall cost, but the broad commercial market for using them has not yet been developed.

[0028]

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

[0029]

[0030] FIG. 2 shows the separation of Service Machine 3 from Tag 1 and the contact of Service Machine 3 with Client 5 that is desired to be removed from orbit.

[0030]

[0031] In some embodiments, Client 5 is equipped with an optical marker before being launched into space. The optical marker reflects light in a predetermined wavelength band emitted from the illumination device of Service Machine 3. An image of the optical marker that has reflected light in the predetermined wavelength band is acquired by the image acquisition device of Service Machine 3. This image is processed by the image processing device. As a result, Service Machine 3 can estimate the attitude of the target object.

[0031]

[0032] In some embodiments, a capture plate can be attached to the client 5. The capture plate can be attached to a portion of the client 5 where the capture device joins with the joining component of the service device 3. The joining component includes an adhesive. By attaching the capture plate to the client 5, the joining component of the service device 3 can be easily and reliably joined 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 guiding structure for defining the joining position between the client 5 and the service device 3. Thereby, the accuracy of the joining position between the client 5 and the service device 3 can be improved.

[0032]

[0033] In some embodiments, a set of capture arms extends from the service device 3 and operates to engage the client 5. The set of capture arms includes at least one service umbilical, the first end of the service umbilical is attached to the service device 3, the second end of the service umbilical is equipped with a service umbilical end connector, and the service umbilical end connector is configured to form a connection with the client umbilical connector of the client 5. Also, the set of capture arms includes a manipulator arm, the first end of the manipulator arm is coupled to the service device 3, and the second end of the manipulator arm is connected to the second end of the service umbilical and configured to operate it. Also, the set of capture arms includes a processor that operates to control the manipulator arm. The manipulator arm operates the second end of the service umbilical to form a connection between the service umbilical end connector and the client umbilical connector.

[0033]

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

[0034]

[0035] The client 5 can vary widely in size and shape. The client 5 can be a small debris piece with a diameter of less than 10 cm or a large satellite weighing several tons. In the low Earth orbit (800 - 2,000 km) where most satellites operate, there are millions of debris pieces that are too small to be tracked. Approximately 35,000 space debris pieces with a diameter exceeding 10 cm have been individually tracked. To date, approximately 9,000 satellites have been launched into low Earth orbit, of which only about 2,000 are in operation, while most of the rest are still orbiting. Approximately 60% of the satellites in orbit are non-functional and are undesirable debris.

[0035]

[0036] Approximately 24% of the individually tracked debris objects are aging satellites, and approximately 18% are used upper stages and mission-related objects, such as launch adapters and lens covers. Since collisions occur at high speeds, even a fragment of space debris with a diameter of only 1 cm can potentially damage a satellite. Fragments of space debris larger than 1 cm can penetrate the outer walls of existing satellites and spacecraft, and if the outer walls are not reinforced with additional layers, it is considered likely to cause a catastrophic failure.

[0036]

[0037] Typical collisions with space debris occur at a closing speed of 10 km / s, which is equivalent to 36,000 km / h. An 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 matter (body) around which the object orbits. Objects can be found in many different orbits around the Earth, some of which are traveling in the opposite direction to others.

[0037]

[0038] Figure 3 shows the tag 1 docking with the service machine 3 while the service machine 3 maintains the holding of the client 5.

[0038]

[0039] The docking of spacecraft is the connection of two spacecraft. This connection can be temporary or can be partially permanent, such as for a space station module. Docking specifically refers to the connection of two separate freely flying spacecraft. For an orbital rendezvous to occur, both spacecraft need to be in the same orbit and the positions and rotations of the spacecraft within the orbit need to be matched. Before docking the tag 1 and the service machine 3, the translational speeds and angular velocities of the tag 1 and the service machine 3 need to be made approximately equal.

[0039]

[0040] To achieve this object, a rotation suppression device for suppressing the rotation of the client 5 can be provided. The rotation suppression device includes a main body, a shaft extending outward from the main body and configured to rotate about a first rotation axis, a rotation part attached to the shaft end on the side opposite to the main 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 a target, a braking part provided in 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 generated 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 rotation part fixed to the capture part and the shaft attached to the rotation part can rotate together with the client 5. Next, the braking part gradually suppresses the rotation of the shaft, and the main body rotation suppression part suppresses the rotation of the main body generated when the braking part operates. That is, the capture part rotates integrally with the client 5, for example, and this rotation is suppressed by the braking part, thereby suppressing the rotation of the main body generated when the braking part operates (that is, while maintaining the position of the main body, the angular momentum of the client 5 moves to the main body rotation suppression part and is absorbed there). As a result, the rotational movement of the relatively large (having a large angular momentum) client 5 can be effectively suppressed.

[0040]

[0041] The docking and mooring system on tag 1, service machine 3, and client 5 can be either asexual or sexual depending on the design of the attachment parts. The initial systems for coupling spacecraft were all sexual docking system designs. A sexual design is a form of gender mating where each component of the interface to be connected has a unique design (e.g., a "male" shape or a "female" shape) and plays a specific role in the docking process. This role cannot be swapped, and two spacecraft with the same connection attachment parts cannot be connected. In contrast, an asexual docking (and asexual mooring) scheme uses the same interface on all spacecraft so equipped and can be used to connect spacecraft.

[0041]

[0042] In an embodiment, the client 5 may not include propulsion means, for example, when there is no remaining fuel storage in the client 5. The service machine 3 can induce the movement of the client 5 to facilitate the formation of a composite with the tag 1. The tag 1 and the service machine 3 can be equipped with a computerized control system. This system can use data collected by a LIDAR ranging sensor and an infrared visible light camera when inducing the movement of the client 5.

[0042]

[0043] Figure 4 shows the service machine 3 attached to the client 5 being brought into a lower orbit by the tag 1.

[0043]

[0044] When Tag 1 generates a thrust in a direction opposite to the current direction of motion, Tag 1, Service Machine 3, and Client 5 can fall into a low-energy elliptical transfer orbit. Then Tag 1 can generate a thrust to insert Tag 1, Service Machine 3, and Client 5 into the corresponding low-energy circular orbit. Conversely, when Tag 1 generates a thrust in the same direction as the current direction of motion, Tag 1, Service Machine 3, and Client 5 can rise into a high-energy elliptical transfer orbit. Then Tag 1 can generate a thrust to insert Tag 1, Service Machine 3, and Client 5 into the corresponding high-energy circular orbit.

[0044]

[0045] In an embodiment, the Hohmann transfer orbit can provide a way to move Tag 1, Service Machine 3, and Client 5 into a higher or lower orbit. The Hohmann transfer orbit is a tangent to both the current orbit and the desired orbit of Tag 1, Service Machine 3, and Client 5. To initiate a change in orbit, when moving to a higher orbit, Tag 1, Service Machine 3, and Client 5 are propelled in the direction of motion to accelerate them along the elliptical Hohmann transfer orbit, and when a transition to a lower orbit is desired, they are propelled in a direction opposite to the direction of motion to decelerate them. When Tag 1, Service Machine 3, and Client 5 reach the contact point between the Hohmann transfer orbit and the desired orbit, Tag 1, Service Machine 3, and Client 5 can accelerate and change their motion so that they proceed within the desired orbit.

[0045]

[0046] As used herein, a low orbit means an orbit having a low altitude when measured from the center of the Earth, and a high orbit means an orbit having a high 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 having an altitude of 2,000 km (about 1,200 miles) or less. The high-orbit class includes the medium Earth orbit, sometimes called an intermediate circular orbit (ICO), and also includes the geostationary orbit further above. An object within the geostationary orbit moves at the same angular velocity as the rotation of the Earth, so it remains at a single point in the air when viewed from the Earth's surface. A high-Earth orbit is a geocentric orbit having an overall higher altitude than the geosynchronous orbit (35,786 kilometers (22,236 miles)).

[0046]

[0047] Objects within a low-altitude orbit (less than about 500 km) are affected by atmospheric drag. Since atmospheric drag reduces the kinetic energy of an object in orbit, the object accordingly descends in altitude until it re-enters the atmosphere. Thus, atmospheric drag moves an object from its orbit without human intervention. Objects within a low orbit are greatly affected by atmospheric drag and as a result have a fast decay. The decay life of a space object depends on its altitude, the level of solar activity, and the mass relative to the cross-sectional area. An object with a large mass-to-area ratio has a small resistance effect and thus a slow decay. High solar activity increases the atmospheric density and atmospheric drag in the low-Earth orbit. For an object on an orbit at a relatively low altitude, if the operation of raising the orbit is not sometimes performed, since the atmospheric drag is strong enough, it may cause re-entry before the intended end of the mission. On average, if client 5 is initially within an orbit at an altitude of 300 km, client 5 has a decay life of only a few months. If client 5 is within a 500 km orbit, its life is about 10 years. If client 5 is at an altitude of 1,000 km, it can remain in orbit for thousands of years without the intervention of an external force acting on client 5.

[0047]

[0048] In some embodiments, instead of moving the client 5 to a low orbit, the client 5 is moved to a high orbit known as a graveyard orbit, junk orbit, or disposal orbit by the complex formed by the tag 1 and the service machine 3. For satellites in geostationary orbit and geosynchronous orbit, the graveyard orbit is several hundred kilometers above the operating orbit.

[0048]

[0049] In FIG. 5, the reentry shepherd 7 has been launched.

[0049]

[0050] In an embodiment, the reentry shepherd 7 can be designed to provide guidance, control, and / or thrust to debris objects for a specific purpose of targeted atmospheric reentry. Therefore, the reentry shepherd 7 can be equipped with a chemical or electric propulsion system. In some embodiments, the reentry shepherd 7 may be equipped with an attachment that can be used for the client to increase atmospheric drag. Resistance devices expand the cross-sectional area of the satellite, allowing the atmosphere to slow down the satellite and lower its altitude, but these devices may not guarantee a specific reentry location such as the uninhabited area of the South Pacific.

[0050]

[0051] The mass of the reentry shepherd 7 can be in the range from several hundred kg to 2 - 3 tons. For example, for a debris object of about 3 tons, a smaller reentry shepherd 7 than that for an 8 - ton debris object is required. The design of the reentry shepherd 7 is assumed to be based on the kick stage of a launch vehicle such as Rocket Lab's Photon or Soyuz's Fregat. The final stage of a typical launch vehicle must impart a large amount of impulse quickly, so it has a large thrust (several hundred Newtons) useful for reentry combustion. The final operation usually has to lower the perigee from about 180 km to 50 km to succeed, so multiple burns are not possible in the final reentry process. Usually, with the perigee at these altitudes, control is lost due to aerodynamic disturbances, but the mission fails because atmospheric reentry does not occur immediately. The large thrust helps ensure that a single burn imparts enough impulse to sufficiently change the perigee.

[0051]

[0052] The reentry shepherd 7 usually has a mass between the tag and the service vehicle. The reentry shepherd 7 does not require the agility of the service vehicle 3, nor the long - life and fuel - efficiency of the tag 1. The reentry shepherd 7 only requires a large impulse and thrust. The reentry shepherd 7 is considered to be an extended kick stage of the launch vehicle. In some embodiments, all three spacecraft, namely the tag 1, the service vehicle 3, and the reentry shepherd 7, can be launched together, but in other embodiments, they are all launched separately. The architecture is flexible. Also, one can wait for the optimal time to lower the client 5 from a high altitude to a low altitude and then launch the reentry shepherd 7 several months after the tag and the service vehicle. Since deploying multiple spacecraft simultaneously is time - consuming, this allows for some flexibility in the development, deployment, and operation time - line. Also, the flexibility in the choice of launch vehicle is obtained due to the architecture that can be configured in various ways.

[0052]

[0053] Figure 6 shows the docking of the reentry shepherd 7 with the client 5 maintained in a stable state by the service vehicle 3 attached to the tag 1.

[0053]

[0054] The client 5 can be kept in a stable state by magnetic force, mechanical force, or other forces. The service machine 3 can include, for example, a permanent magnet to facilitate retention after successful docking. Using superconducting wires cooled to extremely low temperatures, magnetic force can be generated in the service machine 3. The client 5 can have a built-in magnet intended to adjust the orientation of the client 5 using the Earth's magnetic field. This built-in magnet can be utilized by the service machine 3 to attract or repel the client 5 or to shift the orbit of the client 5. Also, the service machine 3 may be equipped with other mechanical means for holding the robotic arm or the client 5.

[0054]

[0055] Docking between the service machine 3 and the client 5 can be difficult, but after the service machine 3 takes control, subsequent docking operations become easier. The service machine 3 can grip a small part of the client 5 and leave a large area for another spacecraft such as the reentry shepherd 7 to approach and grip the same interface. This can be achieved by extending the robotic arm of the service machine 3 to avoid another approaching spacecraft, as shown in FIG. 6. In this operation, the service machine 3 can also provide optical guidance to the reentry shepherd 7 by a retroreflector or LED positioned on the robotic arm or end effector of the service machine 3. Since the client 5 may not have an optical reference to assist docking, the ability of the service machine 3 to provide this guidance to the reentry shepherd 7 is a significant advantage.

[0055]

[0056] FIG. 7 shows the separation of the service machine 3 attached to the tag 1 from the client 5 attached to the reentry shepherd 7. The tag 1 and the service machine 3 accelerate in the direction of the next client object. The reentry shepherd 7 uses resistance to lower the orbit of the client 5 until the orbit is low enough for a direct reentry burn.

[0056]

[0057] Ideally, the client 5 completely evaporates while passing through the atmosphere. As the altitude of the client 5 decreases, the heat due to the friction of the gases in the Earth's atmosphere burns up the client 5. As a rough rule of thumb, the temperature in Kelvin units 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, the temperature can reach a height of 7800 K. However, in embodiments, some parts of the client 5 can withstand this high temperature and reach the Earth's surface, creating a risk of damaging people and property on the ground.

[0057]

[0058] Some components of the client 5 (particularly parts made of heat-resistant materials such as titanium) may withstand atmospheric re-entry and fall to the ground. In that case, it is desirable for the re-entry shepherd 7 to take the client 5 to a position that causes re-entry over an unpopulated area of the Earth. In an embodiment, the re-entry shepherd 7 can target a specific atmospheric entry point to reduce the likelihood of a major accident to people or property on the Earth's surface. If the client 5 has operating capabilities and there is fuel remaining at the end of its life, the client 5 can be positioned to re-enter over a large body of water. If the client 5 has no operating capabilities and there is no fuel remaining, a new re-entry shepherd must be used for each client 5 that requires a direct atmospheric entry to provide the guidance and control of the client 5 necessary during re-entry.

[0058]

[0059] If the client 5 does not require guidance and control after being placed on the desired reentry trajectory across the Earth, the reentry shepherd 7 may be reusable. In some embodiments, the reentry shepherd 7 provides a reentry burn at apogee and is then separated from the client 5 and can quickly increase altitude before reaching perigee. This operation process can be completed in just a fraction of the orbital period and may need to be automated. The reentry shepherd 7 can use chemical propulsion for this operation, most likely a bipropellant with at least 400 N of thrust. When the reentry shepherd 7 is separated from the client 5, high accuracy is required for the reentry burn to ensure hitting the target reentry trajectory within the Pacific Ocean. After recovery, the reentry shepherd 7 will probably remain in low Earth orbit and wait for the next client. The reentry shepherd 7 may be carried to different altitudes by tag 1.

[0059]

[0060] Tag 1 and the service vehicle 3 are designed to be reused multiple times for multiple clients, and since some clients may not require direct atmospheric reentry, they can be used with or without the reentry shepherd 7. This reusability of both these highly capable platforms helps reduce the disposal cost per object.

[0060]

[0061] FIG. 8 is a flowchart showing the steps in a computer-implemented method of designing a spacecraft useful for rendezvous, capture, and disposal of an object in orbit, which is the subject of one embodiment. In the embodiment, many variations of this method are possible, and in fact, the order of the steps may vary from embodiment to embodiment.

[0061]

[0062] One embodiment is a computer-implemented method for calculating parameters of a spacecraft. The method includes inputting attributes of one or more target clients via one or more devices (101), and calculating, by one or more processors, one or more tags and / or desired properties of one or more service machines 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 can be various design and aviation parameters related to the predicted performance of tags and / or service machines. Therefore, the output of the desired properties can provide useful information regarding the manufacturing process of tags and service machines, including the materials used, the sizes of components and assemblies, and the shapes of components and assemblies. Also, a reentry shepherd can be constructed according to the desired properties output from the embodiment.

[0063]

[0064] The desired properties can be output in various data formats, including various spreadsheet and word processing programs, and various file formats usable in computer-aided design.

[0064]

[0065] The previously defined ranking can add different amounts of weight according to the received data regarding the client, the rendezvous with that client, and the mission objectives related to its disposal, to different design and aviation parameters. The previously defined ranking can be input via an I / O device such as a keyboard, a mouse, or a memory.

[0065]

[0066] Other embodiments may include calculating the target weight of the fuel of one or more of a tag, a service machine, and a reentry shepherd (105). Other embodiments may include calculating the target time of one or more of launch, rendezvous, and disposal of one or more of a tag, a service machine, a reentry shepherd, and a client (107).

[0066]

[0067] Also, in order to calculate (109) the desired properties of the re-entry shepherds according to the previously defined ranking, computer hardware such as a processor useful for implementing the embodiments can be provided. In an embodiment, 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 executed simultaneously with steps 101 and 103, or step 109 may be executed later and then steps 111 and 113 may be executed.

[0067]

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

Claims

1. A method for performing a rendezvous with an object in orbit, comprising: launching a tag and a service vehicle into a client orbit; separating the service vehicle from the tag; docking the service vehicle with the client; docking a reentry shepherd with the client; separating the service vehicle from the client docked with the reentry shepherd; and a method comprising the steps of:

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

3. The method according to claim 1 or 2, wherein the tag docks with the service vehicle while the service vehicle maintains holding of the client.

4. The method according to any one of claims 1 to 3, further comprising placing the client on a desired reentry trajectory crossing the Earth by the reentry shepherd.

5. The method according to any one of claims 1 to 4, further comprising guiding and controlling the client during atmospheric reentry by the reentry shepherd.

6. The method according to any one of claims 1 to 5, further comprising separating the reentry shepherd from the client.

7. The method according to any one of claims 1 to 6, further comprising refueling the client.

8. The method according to any one of claims 1 to 7, further comprising repairing the client.

9. The method according to any one of claims 1 to 8, further comprising changing the orbit of the client.

10. The method according to any one of claims 1 to 9, further comprising connecting one or more components to the client.

11. The service vehicle closes a high-bandwidth data link to the tag using an omnidirectional antenna; The tag acts as an intermediate communication relay to enable a high-bandwidth link to the ground by a directional antenna on the tag; and The method according to any one of claims 1 to 10, further comprising:

12. A system for performing a rendezvous with an object in orbit, comprising: a first spacecraft including a tag capable of towing a second spacecraft; a reentry shepherd; and the second spacecraft is a service vehicle configured to dock with a tumbling in-orbit client object. The re-entry shepherd is configured to dock with the on-orbit client object docked to the service machine. The service machine is configured to separate from the on-orbit client object docked to the re-entry shepherd, a system.

13. The service machine is equipped with an omnidirectional antenna. The system according to claim 12, wherein the tag is equipped with a directional antenna.

14. The re-entry shepherd is configured to guide the on-orbit client object on a desired trajectory, the system according to claim 12 or 13.

15. The re-entry shepherd is reusable, the system according to any one of claims 12 to 14.

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