Satellite for rendezvous and docking using an electric propulsion thruster

Abstract

A satellite system and method for performing rendezvous and docking between a servicing satellite and an on-orbit satellite, and more particularly, a satellite system and method for performing rendezvous and docking between a servicing satellite and an on-orbit client satellite using an electric propulsion thruster. In one aspect, a servicing satellite having a set of thruster arms each attached to an electric propulsion thruster performs acceleration, deceleration, and steering operations by six-degree-of-freedom positioning of the thrusters, and the same set of thruster arms and thrusters performs position holding of the docked servicing-client satellite system.

Description

Technical Field

[0001] This application is a non-provisional application of U.S. Provisional Patent Application No. 63 / 449,307, filed on March 1, 2023, entitled "SATELLITE FOR RENDEZVOUS AND DOCKING USING ELECTRIC PROPULSION THRUSTERS", and claims the benefit thereof. The entire disclosure of the provisional patent application is incorporated herein by reference.

[0002] The present disclosure generally relates to satellite systems and methods for performing rendezvous and docking between a servicer satellite and an on-orbit client satellite, and more particularly to satellite systems and methods for performing rendezvous and docking between a servicer satellite and an on-orbit client satellite using electric propulsion thrusters.

Background Art

[0003] Conventional servicer satellites use chemical thrusters for rendezvous and docking. However, chemical thrusters are relatively very expensive, inefficient, heavy, and consume large amounts of fuel because of their low thruster fuel efficiency. Typically, at least 16 chemical propulsion thrusters, which can be operated only during rendezvous and docking ("RvD"), are attached to the corners of the service-providing spacecraft. Such chemical thrusters provide thrust along the velocity vector and thrust along the anti-velocity vector (for braking and separation), and also perform out-of-plane maneuvers such that six degrees of freedom (6 DOF or 6 DOF) are achieved. Depending on the servicer satellite, auxiliary electric propulsion thrusters may be used for station keeping after docking by the chemical thrusters. In such a conventional hybrid propulsion system consisting of both chemical propulsion and electric propulsion, the service-providing spacecraft becomes unnecessarily overly expensive, heavy, and complex.

[0004] The electric propulsion servicing satellite system of the present disclosure solves this problem by using an electric propulsion thruster as a dedicated in-orbit propulsion thruster, thereby eliminating chemical thrusters. The electric propulsion servicing satellite system achieves the same maneuverability as a chemical propulsion system used only for RvD, thereby providing a relatively inexpensive, more efficient, lighter, and more environmentally friendly system. The services provided include, but are not limited to, life extension (position holding and attitude control), orbit departure, orbit repositioning, and debris mitigation.

Summary of the Invention

[0005] A satellite system and method of using the same for performing rendezvous and docking between a servicing satellite and an in-orbit satellite are disclosed. The system and method use an electric propulsion thruster exclusively attached to the servicing satellite for performing rendezvous and docking. The servicing satellite includes a set of thruster arms each attached to the electric propulsion thruster, and acceleration, deceleration, and steering are performed by six-degree-of-freedom positioning of the thruster. The same set of thruster arms and thrusters can also perform position holding of the docked servicing-client satellite system.

[0006] Generally, the electric propulsion servicing satellite system, in one embodiment, includes one or more of the following features. 1. Instead of a conventional system that requires 16 chemical thrusters, it is scaled down to a set of four thruster arms with a single electric thruster attached for achieving full directional control 2. Each of the hinge-rotatable thruster arms has a single thruster fixedly attached to the edge of the thruster arm that can generate accelerating and decelerating forces 3.2 pairs of rotatable thruster arms, where the first pair operates to provide forces in the +Z direction and the out-of-plane direction, and the second pair operates to provide forces in the -Z direction and the out-of-plane direction. In other words, the +Z orientation can have an out-of-plane component for steering the spacecraft towards the host while simultaneously controlling the angular momentum (each TA can have its own cant and slew or perform throttling). 4. The thruster is throttlable and gimbaled to adjust or direct the thrust vector during thruster firing. 5. Thrusters can be used for momentum management of servicing and / or docked servicing-client satellite systems.

[0007] For further background and context, and to further meet the written description requirements of 35 U.S.C. § 112, the entire disclosures of the following references are incorporated herein by reference: U.S. Patent No. 10,384,811 to Knirsch; U.S. Patent No. 10,737,807 to Haertel; U.S. Patent No. 11,492,148 to Nicholson; U.S. Patent No. 10,513,352 to Poncet; U.S. Patent No. 7,575,199 to D’Ausilio; U.S. Patent No. 11,292,618 to Weiss; U.S. Patent No. 9,944,412 to Szabo; U.S. Patent No. 10,611,504 to Halsband; and U.S. Patent No. 11,286,061 to Reitman.

[0008] The expressions six (6) degrees of freedom (“DOF”), six degrees of freedom, 6DOF, 6 - DOF, 6 DOF refer to the six mechanical degrees of freedom of the movement of a rigid body in three-dimensional space. Specifically, this rigid body can freely change its position as a combination of translational movements forward and backward, up and down, left and right along three orthogonal axes (e.g., the XYZ axes), and changes in orientation by rotation about three orthogonal axes, often called yaw (the normal axis), pitch (the transverse axis), and roll (the longitudinal axis).

[0009] In one embodiment, a method for performing rendezvous, docking, and position holding between a servicing satellite and a client satellite, the servicing satellite comprising a body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions, a gripping mechanism attached to the body and adapted to be attached to the client satellite, a set of extensible arms connected to respective sets of thrusters, the set of extensible arms comprising: i) a first extensible arm attached to the body at a first position of the first extensible arm and connected to a first thruster at a second position of the first extensible arm; ii) a second extensible arm attached to the body at a first position of the second extensible arm and connected to a second thruster at a second position of the second extensible arm; iii) a third extensible arm attached to the body at a first position of the third extensible arm and connected to a third thruster at a second position of the third extensible arm; and iv) a fourth extensible arm attached to the body at a fourth position of the first extensible arm and connected to a fourth thruster at a fourth position of the second extensible arm, each of the set of extensible arms being 6 of the corresponding set of thrustersA servicing satellite is provided that includes a set of extendable arms configured to achieve (provide) DOF positioning, and a control device configured to perform each injection of a set of thrusters. The method includes deploying the set of extendable arms, accelerating the servicing satellite and forming a first rendezvous thrust vector that acts to reduce the separation distance between the servicing satellite and the client satellite (and in one aspect, additionally or alternatively, to reduce the angular momentum of the servicing satellite to substantially zero) by injecting each of a third thruster and a fourth thruster along the +Z direction, decelerating the servicing satellite and forming a second rendezvous thrust vector that acts to substantially reduce the angular momentum or rotational speed of the servicing satellite to substantially zero by injecting a first thruster and a second thruster along the +X, -X, +Y, -Y, and -Z directions, docking the servicing satellite with the client satellite using a grasping mechanism to form an interconnected unit, and maintaining the interconnected unit in a substantially stationary orbit by operating the first thruster and the second thruster.

[0010] In one aspect, the servicing satellite further includes a set of momentum wheels configured to perform attitude control of the servicing satellite. In another aspect, the first thrust vector of the first thruster and the second thrust vector of the second thruster pass through the center of mass of the servicing satellite. In another aspect, the first position of the first extendable arm is diagonally spaced from the first position of the second extendable arm, and the first position of the third extendable arm is diagonally spaced from the first position of the fourth extendable arm. In another aspect, each of the first thruster, the second thruster, the third thruster, and the fourth thruster is an electric propulsion thruster.

[0011] In another embodiment, a service satellite that performs rendezvous, docking, and position holding with a client satellite, the service satellite comprising: a body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions; a gripping mechanism attached to the body and adapted to form an interconnecting unit attached to the client satellite and comprising the client satellite and the service satellite; a set of extension arms comprising: i) a first extension arm attached to the body at a first position of the first extension arm and connected to a first thruster at a second position of the first extension arm, and ii) a second extension arm attached to the body at a first position of the second extension arm and connected to a second thruster at a second position of the second extension arm, each of the first and second extension arms being configured to achieve 6 DOF positioning of the respective first and second thrusters; a control device configured to: i) cause the first thruster to generate a first thruster thrust vector including a first thruster thrust vector component along at least one of +X and -X directions, at least one of +Y and -Y directions, and at least one of +Z and -Z directions, and ii) cause the second thruster to generate a second thruster thrust vector including a second thruster thrust vector component along at least one of +X and -X directions, at least one of +Y and -Y directions, and at least one of +Z and -Z directions; wherein rendezvous and docking of the service satellite with the client satellite are performed by firing the first and second thrusters, and the interconnected unit is maintained in a substantially stationary orbit by operation of the first and second thrusters. A service satellite is disclosed.

[0012] In one aspect, the first deployment arm further comprises a rotatable medial hinge, is rotatably attached to the body at a first position of the first deployment arm, and is fixedly attached to the first thruster at a second position of the first deployment arm. The second deployment arm further comprises a rotatable medial hinge, is rotatably attached to the body at a first position of the second deployment arm, and is fixedly attached to the second thruster at a second position of the second deployment arm. In another aspect, the first position of the first deployment arm is diagonally spaced from the first position of the second deployment arm. In another aspect, the injection of the first thruster and the injection of the second thruster substantially reduce the angular momentum or rotational speed of the servicing satellite to nearly zero. In another aspect, the servicing satellite comprises i) a third deployment arm attached to the body at a first position of the third deployment arm and coupled to a third thruster at a second position of the third deployment arm, and ii) a fourth deployment arm attached to the body at a fourth position of the fourth deployment arm and coupled to a fourth thruster at a fourth position of the fourth deployment arm. Each of the third deployment arm and the fourth deployment arm is configured to realize the respective 6 DOF positioning of the third thruster and the fourth thruster. In another aspect, the control device is further configured to i) cause the third thruster to generate a third thruster thrust vector including a third thruster thrust vector component along the +Z direction, and ii) cause the fourth thruster to generate a fourth thruster thrust vector including a fourth thruster thrust vector component along the +Z direction. (Note that in one aspect, the two diagonal thrusters can additionally or alternatively generate one or more out-of-plane components to enable maneuvering and angular momentum management). In another aspect, the first position of the third deployment arm is diagonally spaced from the first position of the fourth deployment arm. In another aspect, the first thruster thrust vector passes through the center of mass of the servicing satellite, and the second thruster thrust vector passes through the center of mass of the servicing satellite. In another aspect, the servicing satellite further comprises a set of momentum wheels that operate to perform attitude control of the servicing satellite. In another aspect, each of the first thruster, the second thruster, the third thruster, and the fourth thruster is an electric propulsion thruster.

[0013] In yet another embodiment, a service satellite that performs rendezvous, docking, and position holding with a client satellite, the service satellite comprising: a body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions; a gripping mechanism attached to the body and adapted to form an interconnecting unit attached to the client satellite and comprising the client satellite and the service satellite; at least one deployable arm attached to the body at a first position and connected to at least one thruster at a second position, the at least one deployable arm being configured to achieve 6 DOF positioning of the thruster; and a control device configured to fire at least one thruster, wherein the rendezvous and docking of the service satellite with the client satellite is performed by firing at least one thruster to generate: i) a first thrust vector including a first thrust vector component along at least the +Z direction; and ii) a second thrust vector including a second thrust vector component along at least one of +X and -X, at least one of +Y and -Y, and at least one of +Z and -Z. A service satellite is disclosed.

[0014] In one aspect, at least one thruster is an electric propulsion thruster. In another aspect, the service satellite further comprises a set of momentum wheels operative to perform attitude control of the service satellite. In another aspect, at least one deployable arm further comprises a rotatable intermediate hinge and is rotatably attached to the body at the first position and fixedly attached to the first thruster at the second position.

[0015] In yet another embodiment, a method for performing rendezvous, docking, and position holding between a servicing satellite and a client satellite, the servicing satellite comprising a body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions, a gripping mechanism attached to the body and adapted to form an interconnecting unit attached to the client satellite and comprising the client satellite and the servicing satellite, at least one extendable arm attached to the body at a first position and coupled to at least one thruster at a second position, the at least one extendable arm being configured to achieve 6 DOF positioning of the thruster, a control device configured to eject at least one thruster, providing a servicing satellite comprising: deploying at least one extendable arm; ejecting at least one thruster along the +Z direction to form a first rendezvous thrust vector acting to accelerate the servicing satellite and reduce the separation distance between the servicing satellite and the client satellite; ejecting at least one thruster along the +X, -X, +Y, -Y, and -Z directions to form a second rendezvous thrust vector acting to decelerate the servicing satellite and substantially reduce the angular momentum or rotational speed of the servicing satellite to substantially zero; docking the servicing satellite with the client satellite using the gripping mechanism to form an interconnected unit; and maintaining the interconnected unit in a substantially stationary orbit by operating at least one thruster. A method is disclosed that includes the steps of:

[0016] In one aspect, at least one thruster is an electric propulsion thruster. In another aspect, the servicing satellite further comprises a set of momentum wheels operative to perform attitude control of the servicing satellite. In another aspect, at least one extendable arm further comprises a rotatable intermediate hinge and is rotatably attached to the body at a first position and fixedly attached to a first thruster at a second position.

[0017] In another embodiment, a service satellite configured to perform rendezvous, docking, and position holding with a client satellite, the service satellite comprising: a body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions; a gripping mechanism attached to the body and configured to form an interconnected and docked unit attached to the client satellite and comprising the client satellite and the service satellite; a set of extension arms comprising: i) a first extension arm attached to the body at a first position of the first extension arm and connected to a first thruster at a second position of the first extension arm, and ii) a second extension arm attached to the body at a first position of the second extension arm and connected to a second thruster at a second position of the second extension arm, each of the first extension arm and the second extension arm being configured to achieve 6 DOF positioning of the respective first thruster and second thruster; a control device configured to cause: i) the first thruster to generate a first thruster thrust vector including an accelerating force and at least one of +X and -X directions, at least one of +Y and -Y directions, and a first thruster thrust vector component along the +Z direction, and ii) the second thruster to generate a second thruster thrust vector including a decelerating force and at least one of +X and -X directions, at least one of +Y and -Y directions, and a second thruster thrust vector component along the -Z direction; wherein the service satellite is configured such that the separation distance between the service satellite and the client satellite is reduced by a thrust in the +Z direction, and the approaching speed between the service satellite and the client satellite is decelerated by a thrust in the -Z direction, the service satellite performs rendezvous and docking with the client satellite by jetting of the first thruster and the second thruster, and the first thruster and the second thruster jet to maintain the interconnected and docked unit in a geostationary orbit.

[0018] In another embodiment, a service satellite configured to perform rendezvous, docking, and position holding with a client satellite, the service satellite comprising: a body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions; a gripping mechanism attached to the body and configured to be attached to the client satellite to form an interconnected and docked unit comprising the client satellite and the service satellite; a set of extensible arms comprising: i) a first extensible arm attached to the body and connected to a first thruster, and ii) a second extensible arm attached to the body and connected to a second thruster, the first thruster being configured to generate a first thruster thrust vector including an acceleration force and at least one first thruster thrust vector component along at least one of the +X and -X directions, at least one of the +Y and -Y directions, and at least one of the +Z and -Z directions, the second thruster being configured to generate a second thruster thrust vector including a deceleration force and at least one second thruster thrust vector component along at least one of the +X and -X directions, at least one of the +Y and -Y directions, and at least one of the +Z and -Z directions; a control device configured to, by instructions, i) cause the first thruster to eject so as to generate a thrust in the +Z direction, and ii) cause the second thruster to eject so as to generate a thrust in the -Z direction, wherein the thrust in the +Z direction reduces the separation distance between the service satellite and the client satellite, and the thrust in the -Z direction decelerates the approaching speed between the service satellite and the client satellite, and the service satellite performs rendezvous and docking with the client satellite by ejecting the first thruster and the second thruster, and the first thruster and the second thruster eject to maintain the interconnected and docked unit at rest, is disclosed.

[0019] In one aspect, the control device is further configured to, by instructions, perform momentum management of the service satellite.

[0020] In another embodiment, there is a service satellite configured to perform rendezvous, docking, and position holding with a client satellite. The service satellite includes a body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions, and a gripping mechanism attached to the body and configured to form an interconnected and docked unit attached to the client satellite and including the client satellite and the service satellite. The service satellite further includes: i) a first extensible arm attached to the body at a first position of the first extensible arm and connected to a first thruster at a second position of the first extensible arm; and ii) a second extensible arm attached to the body at a first position of the second extensible arm and connected to a second thruster at a second position of the second extensible arm. Each of the first and second extensible arms is configured to achieve 6 DOF positioning of the respective first and second thrusters. The first thruster is configured to generate a first thruster thrust vector including an accelerating force and at least one first thruster thrust vector component along at least one of +X and -X directions, at least one of +Y and -Y directions, and at least one of +Z and -Z directions. The second thruster is configured to generate a second thruster thrust vector including a decelerating force and at least one second thruster thrust vector component along at least one of +X and -X directions, at least one of +Y and -Y directions, and at least one of +Z and -Z directions. The service satellite also includes a control device configured, by instructions, to: i) cause the first thruster to eject so as to generate a thrust in the +Z direction and accelerate the service satellite; and ii) cause the second thruster to eject so as to generate a thrust in the -Z direction and decelerate the service satellite. The service satellite performs rendezvous and docking with the client satellite by ejecting the first and second thrusters, and the first and second thrusters maintain the interconnected and docked unit in a geostationary orbit. A service satellite is disclosed.

[0021] In another embodiment, there is a service satellite configured to perform rendezvous, docking, and position holding with a client satellite. The service satellite includes a main body having an X-axis with +X and -X directions, a Y-axis with +Y and -Y directions, and a Z-axis with +Z and -Z directions, a gripping mechanism attached to the main body and configured to form an interconnected and docked unit attached to the client satellite and comprising the client satellite and the service satellite, a set of extensible arms including: i) a first extensible arm attached to the main body and connected to a first thruster, and a second extensible arm attached to the main body and connected to a second thruster, wherein at least one of the first thruster and the second thruster is configured to generate an accelerating force, and ii) a third extensible arm attached to the main body and connected to a third thruster, and a fourth extensible arm attached to the main body and connected to a fourth thruster, wherein at least one of the third thruster and the fourth thruster is configured to generate a decelerating force, and a control device configured to, by an instruction, i) cause at least one of the first thruster and the second thruster to eject so as to generate a thrust in the +Z direction, and ii) cause at least one of the third thruster and the fourth thruster to eject so as to generate a thrust in the -Z direction. The separation distance between the service satellite and the client satellite is reduced by the thrust in the +Z direction, and the approaching speed between the service satellite and the client satellite is decelerated by the thrust in the -Z direction. The service satellite performs rendezvous and docking with the client satellite by ejection of at least one of the first thruster and the second thruster and ejection of at least one of the third thruster and the fourth thruster. A service satellite is disclosed.

[0022] The phrases "at least one", "one or more", and "and / or" are open-ended expressions that are both conjunctive and disjunctive. For example, each of the expressions "at least one of A, B, and C", "at least one of A, B, or C", "one or more of A, B, and C", "one or more of A, B, or C", and "A, B, and / or C" means only A, only B, only C, A and B together, A and C together, B and C together, or A, B, and C together.

[0023] The term "one (a)" or "one (an)" entity refers to one or more of that entity. Thus, the terms "one (a)" (or "one (an)"), "one or more", and "at least one" may be used interchangeably herein. It should also be noted that the terms "comprising", "including", and "having" may be used interchangeably.

[0024] The term "automatic" and variations thereof, as used herein, refer to any process or operation that occurs without significant human input during the execution of the process or operation. However, a process or operation may be automatic if the input, whether significant or not by human standards, is received prior to the execution of the process or operation. Human input is considered significant if such input affects the manner in which the process or operation is executed. Human input that merely authorizes the execution of the process or operation is not considered "significant".

[0025] "Determine", "calculate", and "compute", and variations thereof, as used herein, are used interchangeably and include any type of methodology, process, mathematical operation, or technique.

[0026] As used herein, the term "means" shall be given its broadest possible interpretation in accordance with 35 U.S.C. § 112, paragraph 6. Accordingly, claims that recite the term "means" shall be construed to cover all structures, materials, or acts specified in this specification, and equivalents thereof. Further, structures, materials, or acts, and equivalents thereof, shall include all those described in the summary of the invention, brief description of the drawings, detailed description, abstract, and claims themselves.

[0027] Various embodiments or portions of the manufacturing method may also or alternatively be implemented in part by software and / or firmware, such as by analysis of codes. This software and / or firmware can take the form of instructions contained in or on a non-transitory computer-readable storage medium. These instructions can then be read and executed by one or more processors to enable performance of the operations described herein. The instructions can be in any suitable form, such as source code, compiled code, interpreted code, executable code, static code, dynamic code, etc., but are not limited thereto. Such computer-readable media can include, but are not limited to, any tangible non-transitory medium for storing information in a form readable by one or more computers, such as read only memory (ROM), random access memory (RAM), magnetic disk storage media, optical storage media, flash memory, etc.

[0028] The foregoing is a simplified summary of the present disclosure provided to facilitate understanding of some aspects of the present disclosure. This summary is neither an extensive overview nor a comprehensive overview of the present disclosure and its various aspects, embodiments, and / or configurations. This summary is not intended to identify key or important elements of the present disclosure nor to clarify the scope of the present disclosure, but rather to present selected concepts of the present disclosure in a simplified form as an introduction to the more detailed description presented below. As will be understood, other aspects, embodiments, and / or configurations of the present disclosure may utilize one or more of the features defined above or described in detail below, either alone or in combination. Also, although the present disclosure is presented from the perspective of exemplary embodiments, it should be understood that individual aspects of the present disclosure may be separately claimed.

[0029] The present disclosure will be readily understood by the following detailed description in connection with the accompanying drawings, in which like reference numerals are assigned to like elements. The elements of the drawings are not necessarily drawn to scale with each other. Wherever possible, the same reference numerals are used to designate the same features common to each figure.

Brief Description of the Drawings

[0030]

Figure 1A

Figure 1B

Figure 1C

Figure 2

Figure 3A

Figure 3B

Figure 3C

Figure 3D

Figure 3E

Figure 4A

Figure 4B

Figure 5

[0031] Reference is made in detail to representative embodiments below. The following description is not intended to limit the embodiments to one preferred embodiment. On the contrary, alternative forms, modifications, and equivalents that may be included within the spirit and scope of the described embodiments, as defined, for example, by the appended claims, are intended to be included.

[0032] The disclosed apparatus, system, and method of use will be described with reference to FIGS. 1A-5. Generally, a system and method for rendezvous and docking between a satellite and an orbiting spacecraft are described. An electric propulsion servicing satellite system may be referred to as a "servicing satellite system," a "servicing satellite," a "servicing system," or simply a "system." The servicing satellite system may also be referred to as "LEXI." A method of using the electric propulsion servicing satellite system may be referred to as a "method of the servicing satellite system," a "method of the servicing satellite," a "method of the servicing system," a "method of the service," or simply a "method."

[0033] This disclosure relates to a method by which LEXI performs rendezvous and docking with a client satellite or "host" satellite using only electric propulsion. Previous or conventional rendezvous and docking (RvD) of client satellites has been performed using chemical thrusters attached to the corners of a service-providing (or "servicing") satellite, and a set of chemical thrusters provides acceleration in all six directions of movement of the servicing satellite (+Z, -Z, +X, -X, +Y, -Y). To achieve such acceleration on all three sides of X, Y, and Z, a conventional servicing satellite requires 16-24 chemical thrusters (4 thrusters at each corner of each of the 6 panels). Chemical thrusters are used only during the RvD phase. On the other hand, for on-orbit service provision of the host satellite after LEXI docks, 4 electric propulsion thrusters, one attached to each of 4 thruster arms, are used. (For an explanation of 4 thruster arms each having one thruster, see, for example, U.S. Pat. Nos. 11,286,061, 11,117,683, and 10,625,882. All of these are incorporated by reference in their entirety.)

[0034] By including a chemical propulsion system only for RvD, the satellite becomes more expensive and heavier, requires a significant amount of fuel (since chemical thrusters are less fuel-efficient than electric thrusters), the design and integration become more complex, and thus chemical propulsion is a relatively very inefficient and cumbersome system.

[0035] If the RvD phase can be properly executed while achieving the same accuracy and control as chemical propulsion using only an electric propulsion system, significant benefits can be obtained, such as relatively low cost, weight reduction of the satellite, and reduction of system complexity.

[0036] Figure 1A is a prior art Figure 10 showing notations related to the location and orientation of a satellite. A geostationary satellite, such as a communication satellite, is depicted orbiting in an orbital trajectory 7000 substantially above a location L on the Earth's surface. The plane of orbit 7000 is parallel to the Earth's equatorial plane. Satellite 700 is oriented such that its direction of movement 710 substantially points towards location L. Direction 710 coincides with the designated longitude of satellite 700, also known as the Nadir direction, or orbital slot, and the longitudinal axis of satellite 700, also known as the Zenith-Nadir direction. In an external reference frame, the axis passing through satellite 700 and parallel to the north-south of the Earth is denoted as the N-S axis of the satellite, while the axis passing through satellite 700 and perpendicular to arrow 710 and the N-S axis of the satellite is denoted as the E-W axis of the satellite, where the east direction points to the east of the Earth and the west direction points to the west of the Earth. Thus, the E-W axis of the satellite is substantially within the plane of orbit 7000. Such notations and reference frames will be referred to hereinafter with respect to the electric propulsion service satellite system and method of use.

[0037] Figure 1B is a prior art Figure 12 regarding the notation and location when a conventional thruster used for rendezvous and docking is attached to a satellite body. The conceptual satellite body is shown as a rectangular block consisting of six faces or panels. A set of thrusters is shown attached to four faces or panels of the satellite body (e.g., SE if on the south - east panel, NW if on the north - west panel), and the direction of thruster jet is shown as +Z or -Z. The notations X, Y, Z indicate on which panel the thruster is located. The notations N, S, E, W specify in which part of the north, south, east, or west of the panel the thruster is. Also, the notation Na, Z specifies whether the thruster is on the bottom or top of the panel. The thruster arrangement in Figure 1B is such that the thruster is directly attached or mounted (specifically not on an extended arm) to the satellite body, which is the previous or conventional attachment method for thrusters for rendezvous and docking, and usually uses chemical thrusters.

[0038] Figure 1C is a prior art figure of an existing service satellite 16 designed specifically to perform position - holding of an interconnected service - client satellite system. The service satellite has a set of four thrusters 36, 38, 40, 42 attached to two yoke arms. In particular, note the right - hand side three - axis X - Y - Z reference system of the interconnected service - client. Such a reference system is consistent with those in Figures 1A - 1B.

[0039] FIG. 2 is a block diagram showing one embodiment of the electric propulsion servicing satellite system 200 of the present disclosure. The servicing satellite system 200 includes a servicing satellite 210, and the servicing satellite 210 includes a servicing body 202, a control device 204, a gripping mechanism 206, an RvD sensor 208 used to enable rendezvous and docking with a client satellite, a momentum wheel 209 used to enable attitude control (i.e., movement with respect to the center of mass ( "CM") of the servicing satellite 210), and a set of extension arms 210, 220, 230, 240 attached to, connected to, or engaged with the corresponding thrusters 211, 221, 231, and 241 of a set of thrusters. (Note that in some embodiments of the electric propulsion servicing satellite system, the "set" of extension arms may include a different number of extension arms other than four, for example, a single extension arm or a pair (i.e., two) of extension arms. Also, in some embodiments of the electric propulsion servicing satellite system, the set of thrusters is other than a set consisting of four thrusters, for example, two thrusters, a single thruster configured to inject in two or more directions, etc. Such embodiments will be described in detail below).

[0040] The servicing satellite 210 operates to rendezvous and dock with a client satellite 100 including a client body 102 and an interface ring 106 attached to the client body 102. The gripping mechanism 206 attached to the servicing body 202 is adapted to be attached to the client satellite 100 or dock with the client satellite 100 via the interface ring 106. In some embodiments, the gripping mechanism 206 is configured to be attached to any external component of the client 100, such as a client nozzle, a flange of the client body 102 or a flat surface of the client body, a docking plate such as a magnetic docking plate, etc., as is known to those skilled in the art.

[0041] The control device 204 is configured to operate each of the set of thrusters to fire, to operate each of the set of deployment arms, to operate the momentum wheel 209, and / or to operate the RvD sensor 208. The control device 204 may comprise a set of control devices. For example, one control device controls aspects of operations such as the electric propulsion RvD, position holding, etc. of the system, and another control device processes aspects of the previous system such as momentum management. The set of thrusters is configured to perform any of several functions including rendezvous, docking, proximity operations (e.g., to a client satellite), tandem position holding (e.g., of interconnected service - client units), and debris removal.

[0042] The service satellite 210, also known as LEXI, includes several deployable mechanical thruster arms (also referred to as "extension arms") or a combination of such mechanical thruster arms. Further, LEXI includes a set of small thrusters, each attached to the edge of a deployable mechanical thruster arm (see FIG. 3A). By way of example, four thrusters can be attached in a generally symmetric rectangular arrangement. The four thrusters include a northeast thruster N1 connected to the deployable mechanical thruster arm via a first thruster positioning device, a northwest thruster N2 connected to the deployable mechanical thruster arm via a second thruster positioning device, a southeast thruster S1 connected to the deployable mechanical thruster arm via a third thruster positioning device, and a southwest thruster S2 connected to the deployable mechanical thruster arm via a fourth thruster positioning device (FIGS. 3A, 3B, 3D, and 3E). These thrusters can be of any type available in the commercial market and / or known to those skilled in the art, such as, for example, chemical propulsion thrusters or electric propulsion thrusters, among others, bi-propellant thrusters, Hall effect thrusters, Gridded Ion Thrusters, arcjets, resistojet thrusters, plasma propulsion engines, and the like. Each deployable mechanical thruster arm has a positioning device incorporating two hinges rotatable at two angles Θ1, Θ2 about a single axis (see FIG. 3B and related H hinge and F hinge). Other configurations of hinges and / or rotational components are possible that provide the kinematic ability to position the thrusters attached or disposed at the distal end or portion of the thruster arm such that 6 DOF positioning of the thrusters is achieved. For example, ball and socket joints, four-bar link mechanisms, etc., as known to those skilled in the art, can be used.

[0043] LEXI has a thruster arm ("TA") design with two rotatable arms (F hinge and H hinge). By rotating these arms, each of the four thrusters can produce forces about all axes. Also, the thrust vector may pass through the center of mass (C.O.M. or COM) of LEXI so that no induced torque occurs.

[0044] The thruster arms can be gimballed at various positions on the central satellite body to generate forces acting in any in-plane direction, thereby improving maneuverability and controllability. There are at least two ways to distinguish the resulting forces in each plane. One way is to arrange each of the thruster arms at different cant and slew angles (see Figure 3C), thereby creating different force components (of each thrust vector) that pass through the C.O.M. The second way is to position two thruster arms symmetrically with respect to the C.O.M. in a diagonal orientation and use the electric propulsion throttling function to throttle each of the electric propulsion thrusters to create a difference in the forces passing through the C.O.M. without inducing torque. Throttling (by which a selectable thrust vector magnitude is obtained) may be performed in any implementation of an electric propulsion servicing satellite system, and it should be noted that this includes implementations of an electric propulsion servicing satellite system having two thruster arms with the servicing satellites positioned in a diagonal orientation.

[0045] More generally, by selecting and controlling the thrust vector of one or more thrusters of the servicing satellite of an electric propulsion servicing satellite system, movement of the servicing satellite along any or all of the X, Y, and Z axes becomes possible. In other words, by selecting and controlling the thrust vector of one or more thrusters of the servicing satellite of an electric propulsion servicing satellite system, control of the servicing satellite in any of the ±X directions, ±Y directions, and ±Z directions, as well as rotation (or non-rotation) about the three axes of the servicing, becomes possible (see FIG. 3A).

[0046] To perform the RvD operation phase, four thruster arms (see FIG. 3A) are positioned in two fundamental orientations while providing the necessary maneuverability and controllability.

[0047] Two diagonal thrusters (e.g., N1 and S2) are arranged parallel to the positive Z axis of the satellite (see FIG. 3D), enabling the satellite to accelerate in the positive Z direction (i.e., the +Z direction), follow a trajectory such that the orbital parameters of the two satellites match, and reduce the distance between LEXI and the host satellite. Each of these N1 and / or S2 thruster arms can also change their steering angle so as to produce X and / or Y components as well. Any resulting unwanted induced torque can be absorbed by the reaction wheels. The other two diagonal thruster arms (e.g., N2 and S1) are rotated in opposite directions such that the thrust vectors of both thrusters pass through the C.O.M. of LEXI, during which five-directional forces of negative Z (deceleration), positive and negative X, and positive and negative Y are generated (see FIG. 3E).

[0048] For each thruster arm, by imparting an inclination angle (measured from the Y-axis) and a rotation angle (measured from the Z-axis) with respect to the C.O.M. (see Fig. 3C), a thrust in the negative Z-direction is accumulated. By gimbal supporting each of the thruster arms (or otherwise providing an equivalent set of rotations thereto), the resulting X and Y components can be made zero, or these components can generate a net component that is non-zero in either the positive or negative direction for both the X and Y components (the term "net" or "Net" refers to the synthesis of one or more vectors, such as thrust vectors, so that a resulting composite or combined vector is obtained. The addition of such vectors is known to those skilled in the art). Another technique for generating a non-zero net thrust vector (as described above) is to throttle each of the thrusters to generate different forces, thereby generating a non-zero net component in either the positive or negative direction for both the X and Y components.

[0049] Figures 3A - 3E depict additional details of the servicing satellite 310 of the servicing satellite system.

[0050] Referring to Fig. 3A, the electric propulsion servicing satellite 310 having a body 302 is shown having a set of four deployable arms 210, 220, 230, 240 in the deployed position. The body 302 includes a body axis 303 centered on the center of mass (COM or CM) of the body 302, and the body axis 303 includes an X-axis having +X and -X directions, a Y-axis having +Y and -Y directions, and a Z-axis having +Z and -Z directions.

[0051] The electric propulsion service satellite 310 comprises a set of four deployable arms 210, 220, 230, 240 attached, connected, or coupled to respective thrusters 211, 221, 231, 241. Generally, each of the four thruster arms comprises two rigid body parts also referred to as a base link and a row, and involves a linkage mechanism that enables or realizes the six-degree-of-freedom (DOF) positioning of the thruster disposed at the distal portion of the deployable arm.

[0052] The first deployable arm 210 is attached to the main body 302 at the first position 212 of the first deployable arm and is connected to the first thruster 211 (also called the "S1" thruster as it is disposed at the position of the first south panel) at the second position 213 of the first deployable arm. The first deployable arm 210 comprises a base link 217 of the first deployable arm. The base link 217 of the first deployable arm is connected to the main body 302 by a first hinge 214 of the first deployable arm at the first position 212 of the first deployable arm and is connected to a row 218 of the first deployable arm by a second hinge 216 of the first deployable arm. The first thruster 211 is rigidly or fixedly attached to the first deployable arm 210 at the second position 213 of the first deployable arm.

[0053] The second deployable arm 220 is attached to the main body 302 at the first position 222 of the second deployable arm and is connected to the second thruster 221 (also called the "N2" thruster as it is disposed at the position of the second north panel) at the second position 223 of the second deployable arm. The second deployable arm 220 comprises a base link 227 of the second deployable arm. The base link 227 of the second deployable arm is connected to the main body 302 by a first hinge 224 of the second deployable arm at the first position 222 of the second deployable arm and is connected to a row 228 of the second deployable arm by a second hinge 226 of the second deployable arm. The second thruster 221 is rigidly or fixedly attached to the second deployable arm 220 at the second position 223 of the second deployable arm.

[0054] The third deployment arm 230 is attached to the main body 302 at the first position 232 of the third deployment arm and is connected to the third thruster 231 (also called the "S2" thruster because it is located at the location of the second south panel) at the second position 233 of the third deployment arm. The third deployment arm 230 includes a base link 237 of the third deployment arm. The base link 237 of the third deployment arm is connected to the main body 302 by a first hinge 234 of the third deployment arm at the first position 232 of the third deployment arm and is connected to a rod 238 of the third deployment arm by a second hinge 236 of the third deployment arm. The third thruster 231 is rigidly or fixedly attached to the third deployment arm 230 at the second position 233 of the third deployment arm.

[0055] The fourth deployment arm 240 is attached to the main body 302 at the fourth position 242 of the fourth deployment arm and is connected to the fourth thruster 241 (also called the "N1" thruster because it is located at the location of the first north panel) at the fourth position 243 of the fourth deployment arm. The fourth deployment arm 240 includes a base link 247 of the fourth deployment arm. The base link 247 of the fourth deployment arm is connected to the main body 302 by a first hinge 244 of the fourth deployment arm at the first position 242 of the fourth deployment arm and is connected to a rod 248 of the fourth deployment arm by a second hinge 246 of the fourth deployment arm. The fourth thruster 241 is rigidly or fixedly attached to the fourth deployment arm 240 at the second position 243 of the fourth deployment arm.

[0056] As is known to those skilled in the art, other configurations of the telescoping arm and / or hinge and / or thruster attachment portion are possible to achieve 6 DOF positioning of the thruster located at the distal end of the arm. For example, if the thruster is gimbal supported rather than fixedly attached to the end of the telescoping arm, one or both of the hinges described above can reduce their respective degrees of freedom. In one embodiment, the telescoping arm to which the thruster is attached is the same as or very similar to that described in and illustrated in FIG. 1E of Reitman U.S. Patent No. 11,286,061 entitled "Service Satellite for Providing In-Orbit Services using Variable Thruster Control".

[0057] In some embodiments of the electric propulsion service satellite system, the service satellite has three or fewer deployable arms, such as a single thruster arm or two thruster arms. For example, the service satellite can have a single deployable arm that is attached to the main body at a first (proximal) position and attached to a single thruster at a second (distal) position, and the deployable arm is configured to achieve 6 DOF positioning of the thruster. The thruster can be controlled to eject or generate a thrust vector having a thrust vector component along the +Z direction, another thrust vector having a thrust vector component along the -Z direction, and yet another thrust vector having a thrust component along one or more of the +X, -X, +Y, and -Y directions. Similar embodiments may have a single deployable arm, but two thrusters may be attached to the distal end of the deployable arm. Another embodiment may include a single deployable arm having a single thruster capable of ejecting (and thus generating a thrust vector) from multiple sides of the thruster, for example, a thruster that can eject in the +Z direction in some cases and eject in the opposite -Z direction without rotating or moving. In one embodiment, the service satellite has two deployable arms, each deployable arm includes at least one thruster, and one or both thrusters operate to accelerate the service satellite to reduce the separation distance between the service satellite and the client satellite, and one or more thrusters operate to decelerate or slow down or brake the service satellite as the service satellite approaches the client satellite (such one or more thrusters can also perform attitude control or momentum management). In one embodiment, the service satellite performs one or both of the rendezvous operation and the proximity operation, but does not perform the docking operation.

[0058] Characteristics of the deployable arm with a thruster attached (also referred to as "TA" for thruster arm) include · TA has a wide range of TA tilt positions and rotational positions by using two rotational hinges · It is possible to use two diagonally oriented TAs for acceleration while using the other two TAs for deceleration ·The out-of-plane force components (X, Y) also become part of the thrust vector, enabling orbit correction. Either a certain thrust level or throttling can be utilized. ·A momentum management function (such as known to those skilled in the art) is incorporated into the system. ·Due to the use of EP in the RvD phase, ΔV is small but very accurate and the orbit deviation is very small and easily corrected, allowing for continuous injection of the EP thruster. is included.

[0059] Figure 3B is an enlarged view of the fourth thruster stem 240 fixed to the fourth thruster 241 (N1 thruster). As described above, the fourth deployment arm 240 is connected to the main body 302 by the first hinge 244 of the fourth deployment arm at the first position 242 of the fourth deployment arm, and is connected to the lower part 248 of the fourth deployment arm by the second hinge 246 of the fourth deployment arm, and includes a base link 247 of the fourth deployment arm. The fourth thruster 241 is rigidly or fixedly attached to the fourth deployment arm 240 at the second position 243 of the fourth deployment arm.

[0060] The first hinge 244 of the fourth deployment arm is an "F hinge", providing rotation about two axes at the attachment point to the main body 302 (the first position 242 of the fourth deployment arm). (The "F hinge" has a structure and operation as known to those skilled in the art.) In the stowed state, the H angle and the F angle are zero. When deployed, the F hinge angle is measured with respect to the north or south panel plane. The H angle is measured from a line extending from the base link. When the H angle is zero, the lower part extends to the base link, and as a result, the two overlap.

[0061] The second hinge 246 of the fourth deployment arm is an "H hinge", providing rotation about one axis at the connection between the lower part 248 of the fourth deployment arm and the base link 247 of the fourth deployment arm. (The "H hinge" has a structure and operation as known to those skilled in the art.)

[0062] Each of the first and second hinges of the other three extension arms 210, 220, 230 has a hinge set similar to that of the fourth extension arm 240 in FIG. 3B.

[0063] In one embodiment, each of the hinges is a dual hinge rotatable connection. In one embodiment, one or more of the hinges are gimbal connections. In one embodiment, one or more of the extension arms are attached to the service body by a gimbal connection. In one embodiment, one or more of the extension arms are attached to one or more thrusters by a gimbal connection, such as an H hinge or an F hinge, or a rotatable connection.

[0064] FIG. 3C is a diagram 303 of notations and directions related to thruster operation, particularly the cant angle and the slew angle.

[0065] FIG. 3D is a diagram of an embodiment of the electric propulsion service satellite 310 of FIG. 3A, focusing on two thruster arms attached to the corresponding thrusters N1 and S2, showing thruster jets along the +Z axis. The fourth extension arm 240 having the fourth thruster 241 (N1 thruster) that is jetting to generate the fourth thrust vector 249 in the +Z direction is shown. Similarly, the third extension arm 230 having the third thruster 231 (S2 thruster) that is jetting to generate the third thrust vector 239 in the +Z direction is shown. Note that the first position of the third extension arm is diagonally spaced from the first position of the fourth extension arm on the common body panel.

[0066] The injection of the fourth thruster 241 and the injection of the third thruster 231 are performed so as not to induce torque or rotation of the main body about the main body COM. Such a configuration is used to accelerate the electric propulsion service satellite 310 toward the client satellite during the rendezvous operation and the docking operation, as will be discussed in more detail with respect to FIGS. 4A-4B, for example. In other words, the injection of each of the third thruster and the fourth thruster along the +Z direction acts to accelerate the service satellite and to reduce the separation distance between the service satellite and the client satellite, forming a (synthetic or net) first rendezvous thrust vector.

[0067] The features of the configuration of FIG. 3D include · The distance between the LEXI and the target client spacecraft is reduced by the thrust vector in the +Z direction, which is achieved by two diagonal TAs parallel to the C.O.M. · The two diagonal TAs can have inclination and rotation angles such that the thrust components in the +X, -X, +Y, and -Y directions, which enable the maneuvering of the LEXI while keeping the angular momentum under control, can also be achieved. · The same TA orientation can be used during orbit raising and position holding (failure mode). · The combustion commands are generated on the ground and executed on board based on sensor measurements. · Two thrusters can be continuously injected for a long time. · Using sensors and reaction wheels, it is possible to keep the attitude of the spacecraft within the required range by the on-board computer. · The required momentum management is minimal, and the off-plane correction requirements are minimal or not required at all. · The thrust vector is parallel to the C.O.M., and thus no torque is generated. · The momentum management follows the same method as in the case of the position holding method. For example, the reaction wheels are maintained within boundaries using the installed torque table. are included.

[0068] Figure 3E is a view of an embodiment of the electric propulsion servicing satellite of Figure 3A, focusing on the two thruster arms attached to the corresponding thrusters N2 and S1, and shows a thruster jet passing through the center of mass of the electric propulsion servicing satellite that generates a thrust vector component along the -Z axis.

[0069] A first extended arm is shown having a first thruster 211 (S1 thruster) that is ejecting to generate a first thrust vector 219 having components in the -Z direction, +X direction, and +Y direction. A second extended arm is shown having a second thruster 221 (N2 thruster) that is ejecting to generate a second thrust vector 229 having components in the -Z direction, -X direction, and -Y direction. It should be noted that the thrust vectors of each thruster in Figure 3E are inclined and rotated with respect to the Y axis (see Figure 3C), generating three-component vectors, and these inclination angles and rotation angles are directly related to the H angle and F angle (see Figure 3C). (Note that the coordinate axis 303 is offset from the COM of the main body 302 for clarity.) It should be noted that the first position of the first extended arm is diagonally separated from the first position of the second extended arm on the common body panel.

[0070] The configuration of Figure 3E can also achieve the maneuvering of the main body by intentionally shifting the two vectors so that a thrust vector having a non-zero value in one or both of the X and Y axes is ultimately generated (such relative orientation and / or maneuvering is depicted by the rotating part 309). In other words, the first thruster and the second thruster are ejected along the +X, -X, +Y, -Y, and -Z directions so that a second (composite or net) rendezvous thrust vector is formed that acts to decelerate the servicing satellite and substantially reduce the angular momentum or rotational speed of the servicing satellite to nearly zero.

[0071] The features of the configuration of Figure 3D include · The thrust vector in the Z direction (deceleration) can be achieved by directing the other two diagonal TAs to pass through the C.O.M. · Most of the net thrust can be used for deceleration by linking it to a complete out-of-plane correction. · Momentum management may follow the same concept as in the case of the position-holding method, and the reaction wheel is maintained within limits using the installed torque table. · It is possible to operate one or two TAs while managing momentum with the position-holding flight software. · The thrust vector passes through the C.O.M. is included.

[0072] Figures 4A and 4B are respectively a rear view and a side view of a schematic depiction of the combined operation of a set of four thrusters of the embodiment of the electric propulsion servicing satellite of FIG. 3A. As described above, the set of four thrusters can operate in pairs, particularly to impart motion to the servicing satellite along the +Z axis or -Z axis.

[0073] Referring to FIGS. 4A - 4B, thruster 1 411 attached to arm 410 and thruster 3 421 attached to arm 420, which are a pair of two diagonally positioned thrusters, can be oriented or directed parallel to the +Z axis of the servicing satellite 402 and fired simultaneously. This pair of thrusters, thruster 1 411 and thruster 3 421, generates a combined thrust vector that does not generate a torque about the center of mass 403 of the servicing satellite 402. That is, thruster 1 411 generates vector 419 of thrust 1, thruster 3 421 generates vector 429 of thrust 3, and when these are combined, the servicing satellite combined thrust vector 409 is obtained. The servicing satellite 402 will be moved or steered towards the target client satellite or host satellite by the servicing satellite combined thrust vector 409. The two diagonal thrusters can have tilts and rotation angles such that thrust components in the +X, -X, +Y, and -Y directions, which enable the steering of the LEXI while keeping the angular momentum under control, can also be achieved.

[0074] Continuing to focus on FIGS. 4A-4B, the thrusters 2 431 attached to the arm 430 and the thruster 4 441 attached to the arm 440, which are the other pair of thrusters, can be oriented or directed such that their corresponding thrust vectors pass through the center of mass 403 of the servicing satellite 402. Such orientation of each of the thrusters 2 431 and 4 441 is achieved by the gimbaled movement of their respective arms 430 and 440.

[0075] This pair of thrusters, thruster 2 431 and thruster 4 441, generates a combined thrust vector that does not generate a torque about the center of mass 403 of the servicing satellite 402 when fired simultaneously. That is, thruster 2 431 generates the vector 439 of thrust 2, and thruster 4 441 generates the vector 449 of thrust 4, and when these are combined, a servicing satellite combined thrust vector in the -Z direction (i.e., the direction opposite to the thrust vector 409) is obtained. Such a combined thrust vector acts to slow down, brake, or disengage a target client satellite or host satellite.

[0076] FIG. 5 is a flowchart of one embodiment of a method of using the electric propulsion servicing satellite system of FIG. 3A. Note that depending on the embodiment, some functions (e.g., any of the identified elements numbered 508-540) may be absent, occur in a different order, or include other functions or steps not shown (e.g., if the client satellite is pre-identified by a ground control device or the like, step 512 may be omitted). Next, the method of use 500 depicted in FIG. 5 will be described with reference to the aspects and features of the systems of FIGS. 1-4 described above.

[0077] Starting at step 504, method 500 proceeds to step 508. At step 508, an electric propulsion servicing satellite system is provided. After completion of step 508, method 500 proceeds to step 512.

[0078] In step 512, the client satellite is identified. There can be various missions in which the client satellite is involved. For example, the mission may include closely approaching the client satellite and observing its condition, may require rendezvous and docking, and / or may include the client's orbital departure. Depending on the characteristics of the client satellite, certain docking requirements may be determined, such as docking by the client's interface ring, docking by the client's nozzle, or docking by the client's panel / body, etc. After completion of step 512, method 500 proceeds to step 516.

[0079] In step 516, a set of four deployable arms of the servicing satellite is deployed. Each of the four deployable arms is attached to a thruster and is configured to orient its respective thruster in six degrees of freedom (i.e., all three attitude arrangements and all three rotational directions). After completion of step 516, method 500 proceeds to step 520.

[0080] In step 520, an inquiry is made to determine whether movement along the +Z axis of the servicing satellite is necessary. Such movement is necessary to reduce the separation distance between the servicing satellite and the client satellite. If the response to the inquiry is YES, method 500 proceeds to step 524. If the response to the inquiry is NO, method 500 proceeds to step 528.

[0081] In step 524, as illustrated in FIG. 3D, both the third thruster (S2) and the fourth thruster (N1) are fired. These thrusters are fired such that no rotation of the servicing satellite about the center of mass of the main body 302 occurs due to the net thrust vector (i.e., the combination of thrust vector 249 and thrust vector 239). (Note that if an out-of-plane component is required, the resulting torque may be absorbed by the reaction wheels of the servicing satellite.) After completion of step 524, method 500 proceeds to step 528.

[0082] In step 528, an inquiry is made to determine whether movement along the -Z axis of the servicing satellite is required and / or whether movement about the X axis and / or Y axis is required. Such -Z movement is required to decelerate the approach speed between the servicing satellite and the client satellite. Such movement about the X and Y axes is required for steering control of the servicing satellite with respect to the client satellite. If the response to the inquiry is YES, method 500 proceeds to step 532. If the response to the inquiry is NO, method 500 proceeds to step 536.

[0083] In step 532, both the first thruster (S1) and the second thruster (N2) are fired. These two thrusters can be fired in either of two modes or manners. In one manner, as depicted in FIG. 3E, the two thrusters are fired such that each of the thrust vectors 219 and 229 passes through the COM of the body 302, thereby imparting no torque or rotation to the servicing satellite. Each of the thrusters has its own corresponding inclination, rotation, and magnitude of thrust that are equal to or different from each other. Each of the thrust vectors passes through the COM and generates a vector having three components. These three components are in the -Z direction, in the +X or -X direction, and in the +Y or -Y direction. By the three components of each thrust vector, it becomes possible to steer the satellite such that only force is generated without generating torque or rotation. For example, by firing the two thrusters simultaneously, a combined (or net) force is generated for each of the three components, making it possible to change the orbital parameters and guide the servicing satellite. For example, when the signs of the inclination and rotation of each of the thrust vectors are opposite to each other and the magnitudes of the thrusts are the same, the Y and X components of each of the thrust vectors cancel each other out such that the force in the Y direction acting on the satellite becomes zero and the force in the X direction acting on the satellite becomes zero, and in this case only a resultant force in the -Z direction is generated. In another mode, for example, in a situation where the on-board computer requires offloading the momentum stored in the reaction wheel or momentum wheel, one or both of the thrust vectors 219 and 229 are fired at their respective inclination angles and rotation angles, and a net thrust vector is generated that is directed at a selectable inclination / rotation such that it does not pass through the COM of the body 302 but instead imparts rotation or momentum to the servicing satellite. This momentum enables offloading of the reaction / momentum wheel and at the same time enables the reaction / momentum wheel to generate a force for steering the servicing satellite. After completion of step 532, method 500 proceeds to step 536.

[0084] In step 536, an inquiry is made to determine whether a unit in which the service satellite is docked with the client satellite and interconnected is formed. If the response to the inquiry is YES, method 500 proceeds to step 540. If the response to the inquiry is NO, method 500 proceeds to step 520.

[0085] In step 540, the geostationary orbit of the interconnected unit (of the service satellite and the client satellite) is maintained by a set of four thruster arms and the associated four thrusters. After completion of step 540, the method proceeds to step 544 and the method ends.

[0086] Generally, the RvD operation phase includes propelling along the positive Z-axis to steer LEXI towards the host satellite. The satellite-borne computer uses the on-board sensors to calculate the thrust profile (acceleration and deceleration) that LEXI should execute. The on-board computer commands the thrusters to fire at the appropriate time while arranging the thruster arms in the appropriate positions and commands the magnitude of the thrust of each of the thrusters. Two diagonal thrusters (e.g., N1 and S2) are continuously fired based on the trajectory profile calculated by the LEXI on-board computer and the guidance control system. While acceleration is required, the two diagonal thrusters operate to reduce the distance between LEXI and the host satellite. When the trajectory requires deceleration, reaching a complete stop of acceleration, or an out-of-plane correction in the X or Y direction, the two diagonal thrusters in the positive Z direction (e.g., N1 and S2) are turned off. The other two diagonal thrusters (e.g., N2 and S1) are used such that the thruster vector is in the negative Z direction to decelerate LEXI. These two thrusters also have the ability to steer the LEXI satellite in the other two planes (X, Y) and perform control to correct the trajectory and remove the disturbance of momentum at the same time. This process continues until LEXI achieves the appropriate distance and accuracy for docking with the host satellite.

[0087] In one embodiment of the operation of the RvD stage, the system control device issues commands to generate thrust in the +Z axis and / or in any other in-plane direction by at least one thruster (including those by the first pair of diagonal thrusters, e.g., N1 and S2), and / or commands to generate thrust in the -Z axis and / or in any other in-plane direction by at least one thruster (including those by the second pair of diagonal thrusters, e.g., N2 and S1). More generally, the system control device can command the firing of one or more thrusters to generate thrust in any selectable direction or orientation and thrust value, i.e., any selectable thrust vector.

[0088] After LEXI has completed its on-orbit service, the two diagonal thrusters are rotated to positions that produce a force in the negative Z direction, safely pushing LEXI away from the host (client) satellite to a safe distance and maneuvering LEXI into its parking orbit.

[0089] In one embodiment, one or more deployable arms are deployable extension arms.

[0090] Note that the electric propulsion service satellite system and method of use have focused on the space environment, but the system and method can also be applied in other environments. For example, the system can be applied in an underwater environment.

[0091] The exemplary systems and methods of the present disclosure are described in the context of the space environment. However, to avoid unnecessarily obscuring the present disclosure, some known structures and devices, as well as other applications and embodiments, have been omitted from the above description. This omission should not be construed as limiting the scope of the claims. Specific details are set forth so that the present disclosure may be understood. However, of course, the present disclosure may be practiced in various ways that do not remain limited to the specific details described herein.

[0092] Furthermore, of course, the various links connecting the elements can be wired or wireless links or any combination thereof, or any other known or later developed element capable of supplying and / or communicating data with the connected elements. These wired or wireless links can also be secure links and can communicate encrypted information. The transmission medium used as the link can be any suitable carrier for electrical signals including, for example, coaxial cables, copper wires, and optical fibers, and can also take the form of sound waves or light waves such as those generated during radio waves and infrared data communication.

[0093] Also, although the method has been discussed and described in relation to a particular sequence of events, of course, changes, additions, and omissions to this sequence can be made without significantly affecting the operation of the disclosed embodiments, configurations, and aspects.

[0094] Some modifications and variations of the present disclosure can be used. It may be possible to provide some features of the present disclosure without providing other features.

[0095] In the present disclosure, the components and functions implemented in the aspects, embodiments, and / or configurations are described with reference to specific standards and protocols, but those aspects, embodiments, and / or configurations are not limited to such standards and protocols. There are other similar standards and protocols not mentioned in this specification, and they are considered to be included in the present disclosure. Furthermore, the standards and protocols mentioned in this specification, as well as other similar standards and protocols not mentioned in this specification, can be replaced at any time with faster or more effective equivalents having essentially the same functions. Such replacement standards and protocols having the same functions are considered to be equivalents included in the present disclosure.

[0096] In various aspects, embodiments, and / or configurations, the present disclosure includes components, methods, processes, systems, and / or apparatuses substantially as depicted and described herein, including various aspects, embodiments, embodiments of the configurations, sub-combinations, and / or subsets thereof. Those skilled in the art will understand how to make and use the disclosed aspects, embodiments, and / or configurations upon understanding the present disclosure. In various aspects, embodiments, and / or configurations, the present disclosure includes providing devices and processes in the absence of things not depicted and / or described herein or in its various aspects, embodiments, and / or configurations, including, for example, the absence of things that may have been used in previous devices or processes for purposes such as performance improvement, ease of implementation, and / or reduction of implementation costs.

[0097] The foregoing has been presented for purposes of illustration and description. It is not intended to limit the present disclosure to one or more forms disclosed herein. For example, in the above "Modes for Carrying Out the Invention", for the purpose of simplifying the present disclosure, various features of the present disclosure are grouped together in one in one or more aspects, embodiments, and / or configurations. The features of the aspects, embodiments, and / or configurations of the present disclosure may be combined in alternative aspects, embodiments, and / or configurations other than those described above. The method of this disclosure should not be construed as intending that the claims require more features than are expressly recited in each claim. Rather, as reflected in the following claims, aspects of the invention consist of features that do not meet all of the above single aspects, embodiments, and / or configurations disclosed. Accordingly, the following claims are incorporated herein in this "Modes for Carrying Out the Invention", and each claim stands on its own as an individual preferred embodiment of the present disclosure.

[0098] Furthermore, this specification includes descriptions of one or more aspects, embodiments, and / or configurations as well as specific changes and modifications, but other changes, combinations, and modifications are within the scope of the present disclosure, for example, they may be within the skill and knowledge of those skilled in the art after the present disclosure has been understood. Whether or not alternative, interchangeable, and / or equivalent structures, functions, scopes, or steps are disclosed herein for the claimed structures, functions, scopes, or steps, it is intended to obtain the right to include such alternative, interchangeable, and / or equivalent structures, functions, scopes, or steps in alternative aspects, embodiments, and / or configurations to the extent permitted, and it is not intended to publicly disclose any patentable subject matter.

Claims

1. 1. A method for rendezvous, docking, and station keeping between a servicer satellite and a client satellite, comprising: a servicer satellite, the servicer satellite comprising: a body having an X-axis having a +X-direction and a -X-direction, a Y-axis having a +Y-direction and a -Y-direction, and a Z-axis having a +Z-direction and a -Z-direction; a gripping mechanism attached to the body and adapted to be attached to the client satellite; a set of extension arms coupled to respective sets of thrusters, the set comprising: i) a first extension arm attached to the body at a first position of the first extension arm and coupled to a first thruster at a second position of the first extension arm; ii) a second extension arm attached to the body at a first position of the second extension arm and coupled to a second thruster at a second position of the second extension arm; iii) a third extension arm attached to the body at a first position of the third extension arm and coupled to a third thruster at a second position of the third extension arm; and iv) a fourth extension arm attached to the body at a fourth position of the first extension arm and coupled to a fourth thruster at a fourth position of the second extension arm, each of the sets of extension arms configured to provide 6 DOF positioning of a respective set of thrusters; a controller configured to provide a burn for each of said sets of thrusters; deploying the set of extension arms; firing each of the third thruster and the fourth thruster along the +Z direction to form a first rendezvous thrust vector acting to accelerate the servicer satellite and to reduce a separation between the servicer satellite and the client satellite; firing the first thruster and the second thruster along the +X, -X, +Y, -Y, and -Z directions to provide a second rendezvous thrust vector that acts to decelerate the servicer satellite and to substantially reduce the angular momentum or rotational rate of the servicer satellite to approximately zero; docking said servicer satellite with said client satellite using said gripping mechanism to form an interconnected unit; maintaining the interconnected units in a substantially geostationary orbit by operating the first thrusters and the second thrusters. A method comprising:

2. 10. The method of claim 1, wherein said servicer satellite further comprises a set of momentum wheels operative to provide attitude control of said servicer satellite.

3. 2. The method of claim 1, wherein a first thrust vector of the first thruster and a second thrust vector of the second thruster pass through a center of mass of the servicer satellite.

4. 2. The method of claim 1, wherein the first position of the first extension arm is diagonally spaced from the first position of the second extension arm, and the first position of the third extension arm is diagonally spaced from the first position of the fourth extension arm.

5. 10. The method of claim 1, wherein each of the first thruster, the second thruster, the third thruster, and the fourth thruster is an electric propulsion thruster.

6. A servicer satellite that performs rendezvous, docking, and position maintenance with a client satellite, a main body having an X-axis having a +X-direction and a -X-direction, a Y-axis having a +Y-direction and a -Y-direction, and a Z-axis having a +Z-direction and a -Z-direction; a gripping mechanism attached to the body and adapted to be attached to a client satellite to form an interconnection unit comprising the client satellite and the servicer satellite; a set of extension arms comprising: i) a first extension arm attached to the body at a first position of the first extension arm and coupled to a first thruster at a second position of the first extension arm; and ii) a second extension arm attached to the body at a first position of the second extension arm and coupled to a second thruster at a second position of the second extension arm, each of the first extension arm and the second extension arm configured to provide 6 DOF positioning of the first thruster and the second thruster, respectively; a controller configured to i) fire the first thruster to generate a first thruster thrust vector having first thruster thrust vector components along at least one of the +X and -X directions, the +Y and -Y directions, and at least one of the +Z and -Z directions, and ii) fire the second thruster to generate a second thruster thrust vector having second thruster thrust vector components along at least one of the +X and -X directions, the +Y and -Y directions, and at least one of the +Z and -Z directions, rendezvous and docking of the servicer satellite with the client satellite is performed by firing the first thruster and the second thruster; A servicing satellite, wherein said interconnected units are maintained in a substantially geostationary orbit by operation of said first thrusters and said second thrusters.

7. the first extension arm further comprises a rotatable intermediate hinge rotatably attached to the body at a first location of the first extension arm and fixedly attached to the first thruster at a second location of the first extension arm; 7. The servicer satellite of claim 6, wherein the second extension arm further comprises a rotatable intermediate hinge rotatably attached to the body at a first position of the second extension arm and fixedly attached to the second thruster at a second position of the second extension arm.

8. 8. The servicer satellite of claim 7, wherein the first position of the first extender arm is diagonally spaced from the first position of the second extender arm.

9. 9. The servicing satellite of claim 8, wherein the firing of said first thruster and the firing of said second thruster substantially reduces the angular momentum or rotational rate of said servicing satellite to approximately zero.

10. 10. The servicer satellite of claim 9, further comprising: i) a third extension arm attached to the body at a third extension arm first location and coupled to a third thruster at a third extension arm second location; and ii) a fourth extension arm attached to the body at a fourth extension arm fourth location and coupled to a fourth thruster at a fourth extension arm fourth location, each of the third extension arm and the fourth extension arm configured to provide 6 DOF positioning of the third thruster and the fourth thruster, respectively.

11. 11. The servicer satellite of claim 10, wherein the controller is further configured to: i) fire the third thruster to generate a third thruster thrust vector having a third thruster thrust vector component along the +Z direction; and ii) fire the fourth thruster to generate a fourth thruster thrust vector having a fourth thruster thrust vector component along the +Z direction.

12. 12. The servicer satellite of claim 11, wherein the first position of the third extender arm is diagonally spaced from the first position of the fourth extender arm.

13. 13. The servicing satellite of claim 12, wherein the first thruster thrust vector passes through a center of mass of the servicing satellite and the second thruster thrust vector passes through a center of mass of the servicing satellite.

14. 14. The servicing satellite of claim 13, further comprising a set of momentum wheels operable to provide attitude control of the servicing satellite.

15. 15. The servicer satellite of claim 14, wherein each of the first thruster, the second thruster, the third thruster, and the fourth thruster is an electric propulsion thruster.

16. A servicer satellite that performs rendezvous, docking, and position maintenance with a client satellite, a main body having an X-axis having a +X-direction and a -X-direction, a Y-axis having a +Y-direction and a -Y-direction, and a Z-axis having a +Z-direction and a -Z-direction; a gripping mechanism attached to the body and adapted to be attached to a client satellite to form an interconnection unit comprising the client satellite and the servicer satellite; at least one extension arm attached to the body at a first location and coupled to at least one thruster at a second location, the at least one extension arm configured to provide 6 DOF positioning of the thruster; A controller configured to fire the at least one thruster; and Equipped with rendezvous and docking of the servicer satellite with the client satellite by firing the at least one thruster to generate: i) a first thrust vector having a first thrust vector component along at least the +Z direction; and ii) a second thrust vector having a second thrust vector component along at least one of the +X direction and the -X direction, at least one of the +Y direction and the -Y direction, and along the -Z direction; A servicer satellite, wherein said interconnected units are maintained in a substantially geostationary orbit by firing of said at least one thruster.

17. 17. The servicer satellite of claim 16, wherein said at least one thruster is an electric propulsion thruster.

18. 17. The servicing satellite of claim 16, wherein said servicing satellite further comprises a set of momentum wheels operable to provide attitude control of said servicing satellite.

19. 17. The servicing satellite of claim 16, wherein the at least one extension arm further comprises a rotatable intermediate hinge and is rotatably attached to the body at the first location and fixedly attached to the first thruster at the second location.

20. 1. A method for rendezvous, docking, and station keeping between a servicer satellite and a client satellite, comprising: a servicer satellite, the servicer satellite comprising: a body having an X-axis having a +X-direction and a -X-direction, a Y-axis having a +Y-direction and a -Y-direction, and a Z-axis having a +Z-direction and a -Z-direction; a gripping mechanism attached to the body and adapted to be attached to a client satellite to form an interconnection unit comprising the client satellite and the servicer satellite; at least one extension arm attached to the body at a first location and coupled to at least one thruster at a second location, the at least one extension arm configured to provide 6 DOF positioning of the thruster; providing the servicer satellite, the servicer satellite comprising a controller configured to fire at least one thruster; deploying the at least one extension arm; firing the at least one thruster along the +Z direction to provide a first rendezvous thrust vector acting to accelerate the servicer satellite and to reduce a separation between the servicer satellite and the client satellite; firing the at least one thruster along the +X, -X, +Y, -Y, and -Z directions to provide a second rendezvous thrust vector that acts to decelerate the servicer satellite and to substantially reduce the angular momentum or rotational rate of the servicer satellite to approximately zero; docking said servicer satellite with said client satellite using said gripping mechanism to form an interconnected unit; and maintaining the interconnected units in a substantially geosynchronous orbit by operating the at least one thruster.

21. 21. The method of claim 20, wherein the at least one thruster is an electric propulsion thruster.

22. 21. The method of claim 20, wherein the servicer satellite further comprises a set of momentum wheels operative to provide attitude control of the servicer satellite.

23. 21. The method of claim 20, wherein the at least one extension arm further comprises a rotatable intermediate hinge and is rotatably attached to the body at the first location and fixedly attached to the first thruster at the second location.

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