Patent ID: 12202632

DETAILED DESCRIPTION OF EMBODIMENTS

A description of embodiments of the present invention will now be given with reference to the Figures. It is expected that the present invention may be embodied in other specific forms without departing from its spirit or essential characteristics. The described embodiments are to be considered in all respects only as illustrative and not restrictive.

In general, the disclosed system100provides a means for performing on-orbit servicing of spacecrafts. It features mechanisms that enable a host (or servicing) spacecraft (hereafter, S1)102to capture and dock with the spacecraft to be serviced (hereafter, S2)104and a robotic (manipulator) arm108that can provide repairs, change-outs, etc.FIGS.1-16exemplarily illustrates the working of spacecraft servicing system, according to different embodiments of the present invention.

Referring toFIG.1, a nominal configuration of the two spacecrafts (102and104) of the servicing system100, is illustrated. In the figures showing S1102, the manipulator arm108is depicted in a stowed position (it will be recognized that a number of orientations of the arm are possible); in this position, the tool change mechanism will be preloaded to a ground (spacecraft) structure which is not shown.

Referring toFIGS.2and3, an enlarged view of the host spacecraft102and client spacecraft, S2104of the servicing system100respectively, is illustrated. It will be evident that the disclosed system100has the following novel features and advantages over the state-of-the-art each of which will enable it to be simpler and lightweight: The system100is kinematically simple, serially configured, with robot arm. The system100comprises a capture mechanism that is independent of a docking mechanism and vice-versa. The elimination of the need for guidance cones in order to either capture or dock, which take up volume and increase mass. The use of the manipulator arm108to position (align) the captured spacecraft for docking, permits a very flexible, larger, capture envelope and reduces operational complexity. The system100incorporates re-fueling capability into the docking mechanism to eliminate the need for hose management, etc. There is no need for active (motorized) components to capture or dock on the spacecraft to be serviced, S2. The use of the manipulator arm108to deploy the capture/docking boom106eliminates the need for active deployment mechanisms on the boom106. Incorporation of a roll axis about the boom106which increases the work envelope or reach of the manipulator arm108without adding complexity. Further, the servicing vehicle, S1102may be packaged for ride-share capability with the ESPA payload adapter. To emphasize, each of the items above is unique to the disclosed servicing system100and taken together, represents an ambitious and comprehensive approach to on-orbit servicing.

Capture:

In order for any servicing to occur, the host spacecraft, S1102needs to first capture the client spacecraft, S2104, to be serviced. For this purpose, the host spacecraft S1102carries a dedicated boom that is deployed on orbit. At the end of the boom106is the capture mechanism as shown inFIG.4. The capture mechanism is comprised of 3 electromagnets spaced, ideally, 120 degrees apart. The electromagnets are suspended on a frame122that allows for some spherical displacements as a way of compensating for any out of plane misalignments during capture.

The boom106comprises a spherical “rod end” bearing116and a spring preload collar120. A compression spring118is used to lightly preload the electromagnets in their nominal position. Alternatively, three tension springs, attached to the frame122, can be used for the same purpose. Additionally, flex pivots or flexures (not shown in figures) are used to preload the coil124or yoke assembly126of the electromagnets such that, if the assembly approaches the striker plate110(shown inFIG.3) out of plane, the magnetic poles will self-align to be perpendicular to the striker plate110. The striker plate110is located on the client spacecraft, S2104and may be annular in construction; however, to reduce mass, it may simply be discrete pads130as shown inFIG.6A, equal in number to the electromagnets and arrayed similarly. The striker plate (this reference hereafter, refers to the discrete pads or a single annular plate)110may cover an area larger than the footprint of the capture mechanism, that is, the arrayed electromagnets, in order to allow for imprecise or misaligned capturing.

The dimensions of the annular area of the striker plate110are an important factor in providing position information and guidance for capture. The striker plate110is made from a magnetic material such that the energized electromagnets induce an attractive force on the striker plate110and thus the spacecraft to be captured. The boom106may be deployed by active (motorized, springs; not shown in figures) or a passive means. The latter can only occur if the host spacecraft102includes a manipulator arm108, described later, which will be used to deploy the boom106. In all scenarios, once the boom106is deployed, it is locked or latched in position (these features are not depicted in the figures).

The electromagnet-based capture mechanism/system relies on state-of-the-art sensing systems to rendezvous the spacecrafts and to bring them into a close enough alignment and proximity for the capture system to be effective. Once the spacecrafts (102and104) are in close enough proximity to enable capture, it is anticipated that the subsequent operation can be done automatically and without imparting any disorienting impulses on either spacecraft.

In order to reduce power consumption, the function of the electromagnets for and during capture may be replaced by permanent magnets; however, the electromagnets will be used to enable the release and separation of the two spacecrafts (102and104) after servicing. During capture, the electromagnets are also used to induce out of plane positional corrections, in either scenario. Regardless of the configuration used, permanent or electromagnet, the magnetic capture system may include a means for sensing the flux density flowing through the yoke126of the magnetic circuit. Normally, these sensors (for example, hall effect sensors) will all have identical measurements; when brought into near contact (near capture) with striker plate110, unless the magnetic poles are normal to the striker plate110, the sensed flux density values will differ. These measurements can be used in conjunction with other alignment sensors to globally adjust the position of the host spacecraft, S1102with respect to the striker plate110of the client spacecraft, S2104or, they can be used to make adjustments to the capture assemble locally (that is, with respect to the boom). In the latter case, increasing and/or decreasing the power into the three electromagnets, independently, will produce the desired adjustments and this can be done automatically by using a feedback loop.

Capture Interface:

As noted above, capturing of the client spacecraft, S2104by host spacecraft102is accomplished by an interface112(shown inFIG.3) located on client spacecraft, S2104. This interface112could be made into a standardized feature on spacecrafts in order to allow for the possibility of on-orbit servicing. The client spacecraft, S2104further includes one or more possible attachment points114to the manipulator arm108. In one embodiment, the interface112includes three discrete pads130(shown inFIG.6A) that are ferromagnetic, spaced in an array that nominally matched the spacing on the electromagnets. The dimensions of the pads130are determined so as to allow capturing with the maximum possible misalignment between the two spacecrafts (102and104), that is, within the capture envelope or zone. The capture envelope, in turn, is determined by the sensor system used to rendezvous and bring the spacecrafts (102and104) together for proximity operations, and the ability to control movements of the servicing spacecraft, S1102.

The pads130may be offset to project from the plane of the spacecraft's structure128, distance “g”, they are mounted on; in this configuration, the gap, g, between the pad130and the structure creates another interface to which docking, and capturing, is possible. Specifically, a three-pronged grapple can be inserted into the space between the pads130, once they are below the capture surface and in gap g, between the pads130and spacecrafts, rotating these prongs will essentially capture the spacecraft. Additional rotation of the grappling prongs, will preload the connection between the two spacecrafts. This feature may be particularly useful to permanently attach a payload, an avionics box, for example, to client spacecraft, S2104. If the capturing mechanisms are grappling, rather than electromagnet, then there is no need for a ferromagnetic, striker plate110, and the only features that matter are those that will enable the grappling mechanism to interface and lock.

In one embodiment, a further modification to the striker plate110, whether an annular plate or discrete pads130, may be instituted in order to achieve rotational alignment between the two spacecrafts (102and104). This embodiment will include projections from the striker plate110that may be considered “teeth” as in the salient poles of an electric motor. Similarly, the electromagnet may be considered as having two salient poles; thus, energizing the electromagnet and moving it into close proximity with the striker plate will produce a torque on the client spacecraft. S2104, that will seek to align the poles on the two spacecrafts (102and104). This feature will, therefore, automatically compensate for rotational misalignments during capture.

Another arrangement of the striker plate110is that in which it is embedded to the flush with, or recessed into, the surface of the plane on which it is mounted. While this configuration eliminates the possibility of using grappling prongs to dock and capture the spacecraft, the ability to have the salient poles described above will remain.

Docking:

In addition to the capture mechanism, also located on the boom106is a mechanism that enables host spacecraft, S1102to dock with client spacecraft, S2104; this is separate from the capture mechanism. Docking is accomplished when the boom106on host spacecraft, S1102is engaged with a threaded interface138on the client spacecraft, S2104; this interface may be concentrically located with the center of the striker plate110used for capture but an offset is also possible. Detail B,FIG.6, shows a cross-section of the docking elements on the client spacecraft, S2104. The docking mechanism may have four degrees of freedom and is normally recessed below the plane of the electromagnets of the capture mechanism in the stowed position. The first degree of freedom is an axial translation, along the length of the boom106, the second, which may be eliminated, is a rotation or roll about the same axis and the third and fourth are orthogonal translational movements in a plane (X, Y) that is perpendicular to the roll axis. The axial translational motion drives the docking or coupling mechanism into a position beyond the capture mechanism, as shown inFIG.7, in which it can engage with the other half of the docking mechanism on the client spacecraft, S2104.

Referring toFIG.8, the two spacecrafts (102and104) are in docked position, is illustrated. In this embodiment, the docking mechanism140is in extended position, which creates a gap144between the electromagnets' yoke and the striker plate110. To enable docking, the electromagnets of the capture mechanism will be energized (or alternatively, when it is included, the manipulator arm108on the host spacecraft, S1102may position the client spacecraft, S2104for docking, after it has been captured); once docking is completed, the electromagnets are de-energized. There are a number of ways in which the vehicles can be docked or held together, for example, using spring loaded detents or latches that, potentially, interface138with the gap g, ofFIG.6Cor any of the peripheral surfaces of the striker plate. The preferred approach, however, is to use a bayonet-styled coupling system (as in a BNC connector)136. The bayonet connection136can also be designed to permit the transfer of torques to roll S2about the longitudinal axis of the boom106, if this feature is desired. This, in turn, will permit the boom106to position the periphery of S2104in any location to facilitate a repair task by the manipulator arm108. The X, Y movements allow for compensation of misalignments in the plane of the striker plate110, for example, when the capture mechanism is not aligned with the docking feature on the client spacecraft, S2104. X, Y movements are accomplished by mounting the docking mechanism on two linear slides, or stage, and driving (motorizing) each independently. Another possible configuration of the various axis is one in which only one of the X or Y axis is present; in this case, a roll axis about the boom axis must be present.

Internal to the bayonet-styled connection136are additional mechanisms; specifically, rotation of the S1coupling half during docking will produce motion internal to the coupling half of S2which could be a cam/follower action or more simply, a nut and threaded stud engagement as shown inFIG.66. These relative motions will be used to, potentially, mate connectors but more specifically, to open port(s) and a fully mated, leakproof, connection that enables refueling of the client spacecraft, S2104. One way to transfer fuel across the connection would be by centrally locating, in the docking mechanism, a means of connecting to a refuel supply line on the client spacecraft, S2104. This line will be flexible (a hose, for example) or otherwise be articulated in order to enable the boom to be stowed. In this arrangement, there will be no need to separately manage a refueling hose or the like which further reduces complexity and mass.

Manipulator Arm:

FIGS.9-11exemplarily illustrates the operation of the manipulator arm108, is illustrated. In the simplest configuration of the host spacecraft, S1102, a manipulator arm108is included.FIG.9shows a possible configuration of a six degree of freedom arm with arrows indicating the possible motions (146,148,150, and152) at the six joints (it should be understood that the arrows indicate movement in both directions about the identified axes). The manipulator arm108is equipped with various collision avoidance sensors and may incorporate hardstop features to prevent it from damaging its host spacecraft, S1102.

As depicted inFIG.9, the manipulator arm108has the familiar, conventional, layout of a robot and as such, its kinematics is well known and defined for simple operations; again, reducing complexity. This is in direct contrast to the state-of-the-art. Clearly many variations on the geometric layout and location of the arm on S1102are possible. Ultimately, the manipulator arm108is paired with a number of end effectors each of which is designed to perform specific repair tasks. These end effectors are stored in a suite of holsters from which they can be retrieved and stowed by the manipulator arm108. In one embodiment, the manipulator arm108is equipped with a tool-change mechanism that interfaces with the various end effectors each of which has a common interface for compatibility. Located in various positioned on the client spacecraft, S2104are interfaces112(depicted inFIG.3) that are identical to that on the end effectors; which enables the manipulator arm108on the host spacecraft, S1102to “grab” or hold the client spacecraft, S2104; if needed.

In one embodiment, a sensor and vision system on the manipulator arm108will be used to locate potential attachment joints on the client spacecraft, S2104. If the alignment of the mating elements on the spacecrafts (102and104), after capture, does not permit their engagement, the client spacecraft, S2104can be re-positioned by the manipulator arm108in order to achieve alignment and therefore, docking. This repositioning will follow a sequence in which the manipulator arm108on the host spacecraft, S1102takes hold of the client spacecraft, S2104in a hand off from the capture mechanism. The deployed boom106is at a fixed reference with respect to the manipulator arm108; consequently, the manipulator arm108is able to position the client spacecraft, S2104in an optimum alignment location to permit docking. This operational feature may be useful in order to reduce complexity associated with the initial alignment of the two vehicles and the capture of the client spacecraft, S2104. This attribute of the system may be useful in that it greatly expands the capture zone (defined by the striker plate) by simply increasing the size of the striker plate110.

FIG.10depicts the manipulator arm108in the deployed (that is, not stowed) position in which it is ready to perform any repair task; an end effector is not shown.FIG.11shows the possible work envelope for this specific configuration. To reiterate, the docking mechanism may include the ability to roll S2about the boom106. This feature dramatically increases the reach of the manipulator arm to any surface on the periphery of the spacecraft without increasing the complexity of the manipulator arm108itself. If need be, the translational axis degree of freedom of the boom106may be implemented to further increase the reach of the manipulator arm108but this may be redundant in that the several geometries of the manipulator arm108and boom106can be optimized to eliminate this need. However, it is obviously possible to add further degrees of freedom (pitch or yaw, for example) and complexity to the boom106in order to fully maximize robot reach and operations.

As noted, it is clear from the preceding description that the entire on-orbit servicing system100may be configured and deployed in different configurations. A first configuration, will simply be with a deployable boom106that features capture, docking and refueling mechanisms. In this configuration, the boom106will necessarily be actively deployed (springs or motorized). Another configuration which includes the manipulator arm108will enable other repairs or activities. However, the docking mechanism may include an electrical connection132to the client spacecraft, S2104that can be used to perform diagnostic tests, extract or exchange data, etc. Inclusion of the manipulator arm108will, necessarily, include a suite of tools; this will enable full functionality for the execution of repairs on the client spacecraft, S2104.

The following is a description, with related figures, of one specific embodiment of the capturing, and docking systems, reduced to practice and sized for an ESPA-class payload. It should be noted that this disclosure does not limit the invented systems to ESPA-class payload compatibility, and as such, the dimensions noted in the related figures are all subject to change. Each element may be sized according to the requirements of a particular mission and compatibility with the contemplated launch vehicle.

Referring toFIGS.12-16, various views of the spacecraft servicing system100, are illustrated. The spacecraft servicing system100uses at least three mounting brackets (156,157, and159) for demonstration. The capture mechanism154is comprised of three electromagnets assembly174suspended on a frame that is mechanically grounded to the servicing spacecraft, S1102. Each electromagnet has a number of turns of magnet wire (typically, copper) wrapped around a yoke. These are attached to an intermediate frame/plate178that is separated from a ground frame/plate176by three conical (tapered) coil springs or separation springs172. The intermediate frame/plate178is free to move axially with respect to the ground frame/plate176, comprising the conical springs172when it does. The strength of the electromagnet is obviously a function of distance from the striker plate110; each is designed to hold 3 lb. at capture (that is, a total of 9 lbs at distance zero). The entire capture system is mechanically grounded to the host spacecraft102via three struts164attached to the ground plate176.

The docking system154is mounted on an X, Y stage162; each of these axes is driven independently. The stage162permits final alignment of the male docking probe (FIG.12) with the female receptacle, shown in detail inFIG.6.FIG.6also shows the developed design for a “Standard Client Interface” (SCI). The interface includes the BNC-styled receptacle142that mates with the docking probe, an electrical connector half132that is coupled with the opposite half on the docking interface plate (seeFIG.16), the striker plates for the capture system, and alignment holes that guide the alignment pins170during docking in order to compensate for any misalignments along the roll axis.

The docking system154has an x-axis actuator166and z-axis actuator160which drives a leadscrew158. After capture, docking is accomplished when the docking probe168is inserted by the z-axis (vertical) actuation mechanism160, into the receptacle and rotated. Because of the threaded interface138between the receptacle and the housing, rotating the former advances it into the housing by distance, x. Distance x is also the nominal spacing between the intermediate178and ground plates176; thus, the conical springs are compressed by this same distance during docking. The movement, x, is also intended to produce internal actuations on the client spacecraft, S2104, specifically, opening of a port(s) in order to form a continuous, leakproof, path for refueling (note that the docking probe has a through hole or alignment pin hole134that can form a part of the delivery path). Further, the rear side of the docking system154includes a docking interface plate180having a D-sub 9 pin connector182which is merely representative of an electrical connector, and one or more alignment pins170.

To unlock or separate the two spacecrafts (102and104) following on OOS activity, the docking probe is rotated in the reverse direction which resets the internal mechanisms on the client spacecraft, S2104, this is the position shown in detail B ofFIG.6. The compressed conical springs exert a force on the client spacecraft, S2104, thus, enabling the physical separation of the two spacecrafts (102and104) as distance x is traversed. Finally, the docking probe is retracted to its stowed position by using the mechanisms of the z-axis actuator160.

In one embodiment, the system100that utilizes capture and docking mechanisms for spacecrafts. The capturing mechanism and docking mechanism are explained in detail as follows:

Capture:

In order for any servicing to occur, the host spacecraft S1102needs to first capture the client spacecraft, S2104. For this purpose, the host spacecraft102carries a deployable appendage that may be configured as a boom106, with a few degrees of freedom, or a robotic arm108that has many (6 or more) degrees of freedom. In a first embodiment, the capture and docking mechanisms are integrated near the end of the boom106which means that the robotic arm108may or may not be present on the host spacecraft S1102. If the robotic arm108is present, a scaled-down version of the capture mechanism200may be used as an end-effector; this will enable grasping and manipulation of the client spacecraft, S2104, or elements of it. In a second embodiment, the boom106is boom but it only contains or carries the docking mechanisms. The capture mechanism200will be an end-effector which requires a robotic arm108. In a third embodiment, both the capture and docking mechanisms (200and214) form an end-effector; this will also necessarily require a robotic arm108.

Referring toFIG.17, a schematic layout of a capture mechanism or capture system200and docking mechanism214that are mounted on an X, Y stage212, is illustrated. The capture mechanism200is provided at the end of the boom106. The capture mechanism200is comprised of a plurality or array of capture arms202attached to a grounding structure204. In one embodiment, the capture mechanism200may include at least three arms202to capture a circular object. In one embodiment, the capture mechanism200may include at least four arms202to capture both circular and square objects. The arms202are free to rotate with respect to the grounding structure204; this may be accomplished by a single actuator or motor driving a gear that, in turn, drives smaller gears that are coupled to the arm shafts206. This configuration synchronizes rotation of the arms202such that for any angular displacement of a driver gear208, the driven gears and arm shafts206, and thus the arms202, form a circle that is concentric with the driver gear208. The driver gear208may be internal or external, and of any of the standard types, that is, spur, helical, bevel, a worm and gear, or a combination of these types. The driver gear208need not be fully toothed around its pitch diameter, depending on the gear ratios used in the system100, only toothed segments may be needed.

Furthermore, all or some of the gears may be replaced by mechanisms that achieve the same kinematic link107between the driven elements and the driver; for example, sprockets and chain, pulleys and belts, etc. All subsequent references to “gears” are intended to include any form of achieving the kinematic link107. In a second embodiment, the arms202may be independently driven, each by its own actuator/motor (160,166). In a third embodiment, the arms202may be driven by a combination of the aforementioned. As an example of the latter, third, embodiment, if there are three arms202, two may be driven synchronously by a driver gear208and the third arm202may be driven independently by its own actuator/motor (160,166). When driven independently, rotation of the arms202may or may not be synchronized; the latter may be preferred for grasping irregularly-shaped structures.

In a fourth embodiment, at least one arm202is spring-loaded to deploy and hard stop in a predetermined position, such that it acts as a human thumb in a grasping configuration. Finally, for the permanent installation of one object on another, for example, an avionics box on a spacecraft, the capture arms202may simply be spring-loaded; when released from the stowed position, the spring loading will drive the arms202to grasp the object. The arms202may be released or unlatched actively, pyrotechnically, for example, or passively, by contact with an external feature(s).

The length of the capture arms202is sized based on the dimensions of the object to be grabbed, and the amount of gripping force to be exerted on it, for a given motor/actuator. Increasing the length of the arms202obviously increases the size of objects that can be grasped. The capture arms202are provided with one or more features that directly interface or contact with the object to be captured or grasped; these are depicted in figures following as rollers210but need not be as the features simply need to come into contact with the object and, in some instances, may conform to the surface being grabbed. Additionally, the features may be designed such that they are capable of grasping an object both externally and internally (as shown inFIGS.21-22andFIGS.28-30). The end of arm features may be compliant to allow for preloaded grasping although this may be inherently designed into the arm202, itself. Nevertheless, rollers210are the preferred contacting feature as they provide the flexibility and versatility of grasping both inside and outside of an object, such as a capture ring216(as shown inFIG.19). Furthermore, the rollers210allow for rolling contact as opposed to sliding contact with the object being grasped. The rollers210, however, may be mounted such that upon application of the grasping force, their center of rotation is offset or made eccentric such that rolling is no longer possible.

In a first embodiment, an assembly is located between the boom106and the spacecraft structure that allows for motion in an XY (Cartesian) plane that is perpendicular to the main or longitudinal axis of the boom106, referring to the schematic ofFIG.17, in which the X-axis is left to right and the Y axis is in and out of the page. This configuration will be preferred when the docking features on the client spacecraft, S2104are concentric with the capture feature, for example, a capture ring216, as shown inFIG.3. The purpose of this assembly is to compensate for misalignments in the XY plane during capturing. This assembly may be what is commonly referred to as a stage but it is essentially an assembly of motors/actuators, linear motion components and support structure, that provides two independent degrees of freedom that are perpendicular to each other. An alternate method of achieving this compensation capability is by arranging the two axes in an r, 0 (polar) coordinate system such that there is a roll (0) axis about the boom's longitudinal axis and a radial (r) translational axis. When the capture system200is used as a basic gripper, that is, an end-effector at the end of the robotic arm108, it may or may not include this misalignment compensation capability. In the Cartesian configuration, motions in X, and Y, may be effected by small rocket motors instead of electrical motors.

An alternate configuration of the misalignment compensation system is that in which the stage is located such that only the docking mechanisms214are mounted on it; this is depicted schematically inFIG.18. This means that post-capturing, only the docking system214will be aligned with features on the client spacecraft, S2104. This arrangement will be preferred for configurations in which the docking elements or interface are not concentric with the feature that forms the capture interface. For example, referring toFIG.3, in which the features are concentric, if the docking interface was radially offset from the center of the ring, this would require the misalignment compensation system to re-position only the docking system214in order to enable docking. In this alternate configuration, the host spacecraft S1102will have to be pre-positioned within the capture zone in order to enable capturing as there will be no misalignment compensation system. Obviously, a second XY compensation system may be used to re-position the capture mechanisms200as described below; however, this will increase the complexity and mass of the system.

Referring toFIG.18, a schematic layout of the capture mechanism200and docking mechanism214, wherein the docking mechanism214alone in mounted to an X, Y stage212, is illustrated. The capture system200, when used to capture a spacecraft, operates as follows; the sequence (which also includes docking, described further below) is depicted inFIGS.19-32, following. It is assumed that the state-of-the-art sensor suite allows for rendezvous and proximity positioning of the two spacecrafts in order to enable the following sequence to occur, and that the capture and docking systems214are concentric (that is, both can be moved in the XY plane). It is further assumed that the alignment along the roll axis, for both spacecrafts, is within +/−5 degrees or better; again, given the state-of-the-art sensing systems. Also, it will be known beforehand what spacecraft is being grasped, the location, and the physical and geometrical properties of the surface or component to be grabbed. As shown in the figures, the host spacecraft S1102approaches the client spacecraft, S2104with the capture arms202in the stowed position. A vision system, located on the boom106(not shown in figures) may be used to locate fiducial marks on S2; the vision system, and markings, will be located in specific quadrant(s) of the planar surface containing the capture ring216. Once there is no relative motion between the two spacecrafts and they are located within the capture zone, the vision system will record an image of the capture ring216. This image will be processed in order to determine where the center of the ring216is located relative to the center of the capture system200; this difference will represent the amount of misalignment between the two spacecrafts (102and104). This misalignment information will then be sent to the misalignment compensation system which utilizes an algorithm to move the axes in the system in order to center the boom106and capture system200on the capture ring216. This process can be done iteratively, and automatically, until the centers of the two systems align, that is, concentric. After this planar alignment is accomplished, the capture arms202will be driven to encircle and close around the capture ring216; thus, fully executing capture. This process ensures that no disorienting impulses are imparted to either spacecraft. However, if this is not a concern, then there is no need for the alignment process to occur, and capture may be executed as long as the capture arms202overlap the capture ring216.

The boom106may be deployed by active (motorized, springs; not shown in figures) or a passive means. The latter can only occur if the servicing craft includes a manipulator arm, described later, which will be used to deploy the boom106. In all scenarios, once the boom106is deployed, it is locked or latched in position (these features are not depicted in the figures).

To complement the vision system, the capture arms202may include proximity sensors that are capable of detecting the edges of the capture ring216. When the two spacecrafts are in the capture zone, the capture arms202(proximity sensors included) can be used to sweep across the plane of the capture ring216, detecting the edges of the ring, as they pass over them. This information can then be used to help map or confirm the location of the center of the capture ring216, relative to the center of the boom106. Additionally, the capture arms202may be instrumented with strain gages in order to provide force/torque feedback to S1controllers.

Docking:

In the preferred embodiment, in addition to the capture mechanism200, also located on the boom106is a mechanism that enables the host spacecraft, S1102to dock with the client spacecraft, S2104; this is separate from the capture mechanism200. Docking is accomplished when the boom106on the host spacecraft, S1102is engaged with an interface on S2; this interface may be concentrically located with the center of the capture ring216but an offset220is also possible, as previously noted. The docking mechanism214may have up to five actuators and is normally recessed below the plane of the capture arms202of the capture mechanism200, in the stowed position. There is an axial translation motion, along the length of the boom106(a Z-axis), the second, which may be eliminated, is a rotation or roll about the same axis. The third and fourth are orthogonal translational movements in a plane (X, Y) that is perpendicular to the roll axis; if the system is configured as such (disclosed in the Capture section). The fifth actuator or motion rotates the host spacecraft, S1102coupling half during the docking process. Post-capture, the axial translational motion drives the docking or coupling mechanism into a position beyond the plane of the capture mechanism200, as shown in the figures, in which it engages with the other half of the docking mechanism on the client spacecraft, S2104. When it is included, the manipulator arm on the host spacecraft, S1102may position the client spacecraft, S2104for docking, after it has been captured; this assumes that the system is configured such that the docking mechanism214is independent of the capture mechanism200. There are a number of ways in which the vehicles can be docked or held together, for example, using spring loaded detents or latches that, potentially, also interface with the gap g, ofFIG.36or any of the peripheral structures of the capture plate. The preferred approach, however, is to use a bayonet-styled coupling system (as in a BNC connector). The bayonet connection can also be designed to permit the transfer of torques to roll the client spacecraft, S2104about the longitudinal axis of the boom106, if this feature is desired. This, in turn, will permit the boom106to position the periphery of the client spacecraft, S2104in any location to facilitate a repair task by the manipulator arm.

Internal to the bayonet-styled connection are additional mechanisms; specifically, rotation of the host spacecraft, S1102coupling half during docking will produce motion internal to the coupling half of the client spacecraft, S2104which could be a cam/follower action or more simply, a nut and threaded stud engagement; see detail “B”,FIG.39. These relative motions will be used to, potentially, mate connectors, but more specifically, to open port(s) and a fully mated, leakproof, connection that enables refueling of the client spacecraft, S2104. One way to transfer fuel across the connection would be by centrally locating, in the docking mechanism, a means of connecting to a refuel supply line on the client spacecraft, S2104. This line will be flexible, for example, a hose, or otherwise be articulated in order to enable the boom106to be stowed. In this arrangement, there will be no need to separately manage a refueling hose or the like which further reduces complexity and mass. A possible implementation of the docking mechanism214, on the host spacecraft, S1102as shown inFIG.29.

Standard Capture/Docking Interface:

Referring toFIGS.33-43, various views of the docking interface plate215, are illustrated. As noted above, capturing of the client spacecraft, S2104by the host spacecraft, S1102is accomplished by an interface located on the client spacecraft, S2104. This interface could be made into a standardized feature on spacecrafts in order to allow for the possibility of on-orbit servicing. In one embodiment, the interface includes three discrete pads with an axial offset220“g” as shown inFIG.36, equally spaced in an array and concentric with the docking interface. The pads may be ferromagnetic in order to be compatible with an electromagnetic capturing system200. The dimensions of the pads are determined so as to allow capturing with the maximum possible misalignment between the two spacecrafts, that is, within the capture envelope or zone. The capture envelope, in turn, is determined by the sensor system used to rendezvous and bring the spacecrafts together for proximity operations, and the ability to control movements of the servicing spacecraft, S1102.

The pads, or annular ring, will be offset220to project from the plane of the spacecraft's structure, distance “g” inFIG.36, they are mounted on; in this configuration, the gap “:g” between the pad and the structure creates an interface for capturing. The annular ring may be the separation ring that is typically located on spacecrafts to allow separation from a launch vehicle system. Specifically, the disclosed capture mechanism200, with rollers at the end of the capture arms202, can be inserted into the space between the pads and the spacecraft structure, once they are below the capture surface and in gap g, between the pads and spacecraft, rotating the arms202will essentially capture the spacecraft. Additional rotation of the arms202, will preload the connection between the two spacecrafts. This feature may be particularly useful to permanently attach a payload, an avionics box, for example, to spacecraft S2104. If the capturing mechanism200is to be only compatible with the disclosed mechanism, then there is no need for a ferromagnetic, striker plate (that is, pads), and the only features that matter are those that will enable the grappling mechanism to interface and lock.

The features ofFIGS.33-43form the design for a “Standard Client Interface” (SC)215. The interface215includes: the BNC-styled receptacle that mates with the docking probe, an electrical connector half that is coupled with the opposite half on a docking interface plate215. The docking interface plate215includes the capture ring or pads216, a plurality of interface components218, and one or more alignment holes that guide the alignment pins during docking in order to compensate for any misalignments along the roll axis; the several, opposite, interfaces for the host spacecraft, S1102as shown, in one possible configuration. Detail B, ofFIG.39, shows a cross-section of the docking elements on the client spacecrafts, S2104. After capture, docking is accomplished when the probe is inserted by the Z-axis actuation mechanism, into the receptacle, and rotated. Because of the threaded interface between the receptacle and the housing, rotating the former advances it into the housing by distance, x; this is the distance that draws the two spacecrafts together. The movement, x, is also intended to produce internal actuations on the client spacecraft, S2104, specifically, opening of a port(s) in order to form a continuous, leakproof, path for refueling. The docking probe has a through hole that can form a part of the delivery path.

To unlock or separate the two spacecrafts following an OOS activity, the docking probe is rotated in the reverse direction which resets the internal mechanisms on S2, this is the position shown in detail B ofFIG.39. Springs, mounted on the host spacecraft, S1102but in contact with the client spacecraft, S2104, may be used to exert a force on the client spacecraft, S2104, thus, aiding the physical separation of the two spacecrafts as distance x is traversed. Finally, the docking probe is retracted to its stowed position by using the mechanisms of the Z-axis.

According to the present invention, the disclosed capture system200and docking system214include, but not limited to the following advantages. The systems do not use any permanent magnets or electromagnets. They are actively and possibly, passively (with compliance) self-aligning. The active self-alignment does not impart any disorienting forces to either spacecraft. They have possible wider range of capture/grab diameters. They do not require any guidance cones. They can capture/grab on internal or external surfaces. They will grab the nozzle of a thruster engine. They can capture/grab irregular shaped objects if capture arms202are driven independently or in a combination of dependent and independent drives. They can be configured in a few different ways, including as an end-effector. In addition, the misalignment compensation system may be polar coordinates based rather than cartesian; the former may be better when the client spacecraft docking interface218is not concentric with the capture ring216. The system may include no docking mechanism214which could be end-effector when a robot arm is available. Further, spring-loaded capture arms202may be used for permanent installation of an object, for example, avionics box, to spacecraft structure.

Preferred embodiments of this invention are described herein, including the best mode known to the inventors for carrying out the invention. It should be understood that the illustrated embodiments are exemplary only and should not be taken as limiting the scope of the invention.

The foregoing description comprise illustrative embodiments of the present invention. Having thus described exemplary embodiments of the present invention, it should be noted by those skilled in the art that the within disclosures are exemplary only, and that various other alternatives, adaptations, and modifications may be made within the scope of the present invention. Merely listing or numbering the steps of a method in a certain order does not constitute any limitation on the order of the steps of that method. Many modifications and other embodiments of the invention will come to mind to one skilled in the art to which this invention pertains having the benefit of the teachings in the foregoing descriptions. Although specific terms may be employed herein, they are used only in generic and descriptive sense and not for purposes of limitation. Accordingly, the present invention is not limited to the specific embodiments illustrated herein.