Abstract:
A spacecraft dependent on a non-intrusive servicing vehicle is provided. The spacecraft has a spacecraft body having at least one propellant tank and an actual orbit. The spacecraft body is located in an actual orbit and is adapted to be moved to an intended orbit. A payload is connected to the spacecraft body. A non-intrusive servicing interface is connected to the spacecraft. The servicing interface is adapted to repeatedly removably connect the spacecraft body to the non-intrusive servicing vehicle. The actual orbit of the spacecraft body is adjusted by the non-intrusive servicing vehicle and depends upon substantially periodic adjustments of the actual orbit by the non-intrusive servicing vehicle in order to maintain the parameters characterizing the actual orbit within dead bands specified for the intended orbit.

Description:
BACKGROUND OF THE INVENTION  
         [0001]    1. Field of the Invention  
           [0002]    The present invention relates to a spacecraft and, more particularly, to a spacecraft dependent on non-intrusive servicing by a servicing vehicle.  
           [0003]    2. Prior Art  
           [0004]    A major impediment to the exploitation of the economic and scientific potential of orbiting spacecraft, such as for example telecommunications satellites, has been the size and weight limitations imposed on a given orbiting spacecraft&#39;s high value added payload. The high value added payload size and weight limit of a given orbiting spacecraft is constrained by propellant tank size within the orbiting spacecraft. The conventional approach has generally been to provide a full-lifetime propellant load for the orbiting spacecraft in order to allow the orbiting spacecraft to adjust its actual orbit in order to maintain the orbital parameters within the specified dead bands for the mission, an example of such a parameter would be to maintain inclination of less than or equal to 0.05 degrees on a typical geosynchronous spacecraft. When the orbital parameters are all within dead band, the spacecraft is within its intended orbit. As a result of this conventional approach, a typical geosynchronous (GEO) orbiting spacecraft is required to employ large 1000-liter, 49-inch diameter tanks for a full-lifetime propellant load. These large tanks in turn severely limit the high value added payload capacity of the orbiting spacecraft due to their size and weight.  
           [0005]    Accordingly, there is a desire to increase the high value added payload weight and size capacity of orbiting spacecraft. The present invention overcomes the problems of the conventional approach as will be described in greater detail below.  
         SUMMARY OF THE INVENTION  
         [0006]    In accordance with a first embodiment of the present invention, a spacecraft dependent on non-intrusive servicing is provided. The spacecraft has a spacecraft body having at least one propellant tank and an actual orbit. The spacecraft body is located in an actual orbit and is adapted to be transferred to its intended orbit. A payload is connected to the spacecraft body. A non-intrusive servicing interface is connected to the spacecraft. The servicing interface is adapted to repeatedly removably connect the spacecraft body to the non-intrusive servicing vehicle. The actual orbit of the spacecraft body is adjusted by the non-intrusive servicing vehicle. The spacecraft body depends upon substantially periodic adjustments of the actual orbit by the non-intrusive servicing vehicle in order to make the actual orbit parameters lie within the dead bands specified for the intended orbit.  
           [0007]    In accordance with a second embodiment of the present invention, a spacecraft dependent on non-intrusive servicing is provided. The spacecraft comprises a spacecraft body adapted to be placed in an orbit. The spacecraft body is further adapted to be coupled to a payload. A propellant container is coupled to the spacecraft body. The propellant container has a maximum propellant holding capacity and is adapted to contain a maximum propellant quantity sufficient for operation of the spacecraft for a period of time which is less than an expected total propellant quantity for an expected full life time of the spacecraft. A re-fueling interface is coupled to the propellant container. The propellant container is adapted to be re-fueled by the non-intrusive servicing vehicle at the re-fueling interface while the spacecraft body is in the orbit. The propellant container is adapted to be re-fueled by the same non-intrusive servicing vehicle at a substantially periodic interval of time.  
           [0008]    In accordance with a third embodiment of the present invention, a spacecraft servicing system for servicing a plurality of spacecraft that have been placed in substantially similar orbits is provided. These spacecraft orbits may be sufficiently similar such that a spacecraft could transfer from the physical position of one spacecraft to the physical position of another spacecraft using a modest amount of propellant in a modest amount of time. Such amounts could be, for example, three kilograms of propellant and three days. The spacecraft servicing system comprises an orbiting depot having propellant and a non-intrusive servicing vehicle having a propellant container. The non-intrusive servicing vehicle is adapted to selectively transfer propellant from the orbiting depot to the propellant container. The non-intrusive servicing vehicle is adapted to either transfer propellant from the propellant container to the spacecraft or adjust the orbit of the spacecraft.  
           [0009]    In accordance with a method of the present invention, a method of servicing spacecraft is provided comprising a step of placing a plurality of spacecraft in orbits. A non-intrusive servicing vehicle is provided. The non-intrusive servicing vehicle is adapted to re-fuel and/or adjust the orbits of the plurality of spacecraft. A step of re-fueling the plurality of spacecraft or adjusting the orbits of the plurality of spacecraft with the non-intrusive servicing vehicle at substanially periodic intervals of time while the plurality of spacecraft are in the orbits is then provided. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0010]    The foregoing aspects and other features of the present invention are explained in the following description, taken in connection with the accompanying drawings, wherein:  
         [0011]    [0011]FIG. 1 is a schematic perspective view of an orbiting spacecraft servicing system incorporating features of the present invention;  
         [0012]    [0012]FIG. 2A is a schematic perspective view of a prior art spacecraft;  
         [0013]    [0013]FIG. 2B is a schematic perspective view of a spacecraft incorporating features of the present invention;  
         [0014]    [0014]FIG. 3 is a schematic perspective view of a spacecraft incorporating features of the present invention;  
         [0015]    [0015]FIG. 4 is a schematic perspective view of a non-intrusive servicing vehicle incorporating features of the present invention; and  
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT  
       [0016]    Referring to FIG. 1, there is shown a schematic perspective view of an orbiting spacecraft servicing system  10  incorporating features of the present invention. Although the present invention will be described with reference to the single embodiment shown in the drawings, it should be understood that the present invention can be embodied in many alternate forms of embodiments. In addition, any suitable size, shape or type of elements or materials could be used.  
         [0017]    Orbiting spacecraft servicing system  10  generally comprises servicing vehicle  14 , client spacecraft  16 ,  18 ,  20  and  22  and orbiting depot  24 . Servicing vehicle  14 , client spacecraft  16 ,  18 ,  20  and  22  and orbiting depot  24  are shown orbiting earth or body  26 . Servicing spacecraft  14 , client spacecraft  16 ,  18 ,  20  and  22  and orbiting depot  24  may have individual orbits that are congruent with orbit  28  or may involve orbital paths not congruent with orbit  28 . Orbit  28  is shown as a substantially circular geosynchronous orbit (GEO), but may alternately be an elliptical orbit  30  or  32  or may alternately be another orbit about earth or body  26 . Servicing vehicle  14  and orbiting depot  24  are adapted to service and support the  4  client spacecraft  16 ,  18 ,  20  and  22  as herein described but may alternately be adapted to support more or less than 4 client spacecraft. Servicing vehicle  14  is shown supported by a single orbiting depot  24  but may alternately be supported by more than one orbiting depot. Although only a single servicing vehicle  14  is shown, multiple servicing vehicles may be used to interchangeably support multiple client spacecraft.  
         [0018]    Orbiting Depot  24  may be placed in orbit by a Launch Vehicle. Orbiting depot  24  may be placed in a substantially circular geosynchronous orbit (GEO) or alternately another orbit. Orbiting depot  24  has large propellant tanks, sufficient to support a single or multiple re-fuelings of the servicing vehicle. Orbiting depot  24  may be re-supplied by subsequent launch vehicles and/or replaced upon depletion of its propellant supply.  
         [0019]    The servicing vehicle  14  may have multiple large propellant tanks, such as 1000-liter 49-inch diameter tanks. The servicing vehicle  14  may have the capacity for enough propellant to participate in the process of achieving GEO. The servicing vehicle has the capacity for enough propellant to participate in significant maneuvers to support each of the client spacecraft multiple times and in a re-usable manner.  
         [0020]    The servicing vehicle  14  is first inserted into a transfer orbit (GTO), typically by a launch vehicle (LV) from which it can reach GEO using a quantity of propellant, about 1000 kg of propellant or more. The servicing vehicle  14  then transfers from GTO to GEO using its own propellant. Alternately, the servicing vehicle  14  is first inserted directly into or close to GEO, typically by a launch vehicle (LV). The servicing vehicle  14  then transfers to its own GEO using its own propellant. The servicing vehicle  14  may selectively maneuver between the 4 client spacecraft  16 ,  18 ,  20  and  22  and the orbiting depot  24 .  
         [0021]    The client spacecraft  16 ,  18 ,  20  and  22  have either no propellant tanks or small propellant tanks, but would not include large 1000-liter 49-inch diameter tanks like existing geosynchronous (GEO) spacecraft. As a result, these client spacecraft do not have enough propellant to provide a major fraction of the propellant expended in the process of achieving GEO, and depend on either a launch vehicle and/or another external vehicle, such as the servicing vehicle  14  to reach their own GEO.  
         [0022]    The client spacecraft  16 ,  18 ,  20  and  22  may be inserted directly into GEO using a launch vehicle (LV) that has the capability to inject significant spacecraft mass directly into GEO. The LV takes a client spacecraft directly to or close to GEO and separates it there. Each client spacecraft may be injected into a near-circular orbit very near GEO where a small quantity, about 10 kg or less, of propellant is needed for injection into GEO. This injection could be assisted by the servicing vehicle as follows. The servicing vehicle adjusts its orbit, docks with the client spacecraft and acts as a space tug to transfer the client spacecraft to GEO. This eliminates the need for apogee maneuvering firing (AMF) or any major orbit raising activity by each of the client spacecraft. As a result, large client spacecraft tanks are reduced or eliminated from the client spacecraft, thus freeing up considerable volume within the client spacecraft for various uses, including extending the payload from the earth deck toward the anti-earth panel. The need to design the client spacecraft structure to handle the coupled loads from large quantities of fluid within the launch environment is eliminated resulting in lower cost and reduced weight. Deployments and other client spacecraft operations take place in GEO, eliminating outages during Geosynchronous Transfer Orbit (GTO) perigee crossings and the need for the client spacecraft&#39;s earth sensor to operate with an earth of variable apparent size. These advantages, that translate into reduced cost, schedule, complexity and spacecraft development risk, are traded against the significant loss in spacecraft mass capability at beginning of life for each LV for GEO direct insertion of client spacecraft.  
         [0023]    An alternative to the client spacecraft  16 ,  18 ,  20  and  22  being inserted directly into GEO is provided by first inserting each client spacecraft into a transfer orbit from which it can reach GEO using a quantity of propellant, about 1000 kg of propellant or more. This transfer orbit is known as a geosynchronous transfer orbit (GTO) and has apogee or maximum altitude at about the same altitude as GEO and perigee or minimum altitude typically at an altitude of about 200 km. As a result, GTO is a lot easier for a launch vehicle to reach when ferrying each of the client spacecraft as compared to a GEO direct insertion of the client spacecraft. GTO is highly eccentric or elliptical. Once the client spacecraft is inserted into GTO, the servicing vehicle  14  lowers its orbit all the way down to GTO and acts as a space tug to transfer each client spacecraft from GTO to GEO. This eliminates the need for apogee maneuvering firing (AMF) or any major orbit raising activity by each of the client spacecraft alone. As a result, large spacecraft tanks are eliminated from the client spacecraft, thus freeing up considerable volume within the client spacecraft for various uses, including extending the payload from the earth deck toward the anti-earth panel. The need to design the client spacecraft structure to handle the coupled loads from large quantities of fluid within the launch environment is also eliminated resulting in lower cost and reduced weight. These advantages translate into reduced cost, schedule, complexity and spacecraft development risk.  
         [0024]    The servicing vehicle  14  is capable of acting as a space tug for captive-carry through maneuvers and/or as a short-term/just-in-time propellant supplier for maneuvers for each of the client spacecraft to be serviced. The servicing vehicle  14  re-usably supports the client spacecraft without either the client spacecraft or the servicing vehicle  14  being manned. Both the servicing spacecraft  14  and the client spacecraft are un-manned. The servicing vehicle  14  is capable of re-fueling at orbiting depot  24  in order to re-usably support the propellant requirements of servicing multiple client spacecraft. The servicing vehicle  14  is capable of supporting each of the client spacecraft for all significant maneuvers such as atmospheric drag compensation for low earth orbits (LEO), orbit raising maneuvers and north-south station keeping or east-west stationkeeping in geosynchronous orbits (GEO) and attitude control. The servicing vehicle  14  rendezvous and docks with each of the client spacecraft for a few hours each week or each month as appropriate, and may be capable of servicing a dozen or so spacecraft. The servicing vehicle  14  does not perform intrusive servicing, such as equipment change-out or repair. Instead, servicing is limited to non-intrusive activities such as refueling, captive-carry through orbit adjust maneuvers, power transfer to the client spacecraft for battery re-conditioning, operational monitoring, and deployment of stowed equipment. Examples of deployment of stowed equipment include deployment of solar arrays, antenna reflectors, magnetometer booms, solar radiation covers or shields. The client spacecraft being serviced would be entirely dependent upon a servicing vehicle  14  for frequent support in at least some of these activities to perform its mission. The launch mass and therefore cost of the client spacecraft is reduced since it need not carry all propellant for a lifetime. Because it need not carry all propellant for a lifetime, the design of the client spacecraft is simplified resulting in additional available volume at the geometrical center of the client spacecraft body which may be used for high value added payloads or other equipment eliminating the need for group payload equipment near the earth deck or other outer decks of the client spacecraft.  
         [0025]    Table 1 summarizes the cost benefit associated with one non-intrusive servicing operational scenario according to the present invention.  
                       TABLE 1                       Consideration   Client Spacecraft   Servicing Vehicle                   Mode of servicing   Non-Intrusive. No materials   Acts as a space tugboat, may           exchanged   transfer power and signal,               assists with deployments       Frequency of servicing   Monthly   Shuttles between twelve               clients, services 1 client every               2-3 days       External fluid re-supply   None, no fluids resupplied   Obtains propellant from orbiting               depot. Depot resupplied with               propellant carried by Aquarius               or other low cost launch               vehicles and ferried to GEO by               appropriate means, e.g. Orbital               Transfer Vehicles       North-South Stationkeeping   Does not perform   Supports client.       East-West Stationkeeping   Cold-gas   Does not support client       Attitude control   Cold-gas, no momentum or   Momentum wheels, bi-           reaction wheels   propellant, takes over attitude               control of client when docked       Orbit raising   Early spacecraft carried directly   Same as client spacecraft           to GEO by new large           launchers. Later spacecraft           carried to LEO by launcher and           ferried to GEO by Orbital           Transfer Vehicles       Internal configuration   No large tanks, volume   No communications payload           available for thermal control   includes large tanks for           also to minimize spacecraft   maneuvers.           dimensions       Cost relative to current GEO   Lower, plus shorter schedule   Lower, no comm payload                  
 
         [0026]    Referring now to FIG. 2A, there is shown a prior art spacecraft  40 . Prior art spacecraft  40  has a spacecraft body  42 , and at least one propellant tank  44 ,  46 . The spacecraft body is adapted to be coupled to a payload  48 , typically on or adjacent to the earth face  50  of spacecraft body  42 . Payload  48  may be equipment for communications or imaging or other purposes typically found in orbiting unmanned spacecraft. The size of payload  48  is constrained principally by the size of propellant tanks  44  and  46 . Solar arrays  52  and  54  provide power to batteries which in turn may support controllers, attitude and orbital control, momentum wheels, communications and other power based functions typical of unmanned spacecraft. Thrusters  56 ,  58  and  60  provide thrust for all major maneuvers including north south station keeping, east west station keeping, attitude control and orbit raising. Thrusters  56 ,  58  and  60  are supplied propellant from propellant tanks  44  and  46 . The propellant may be cold gas such as pressurized helium or nitrogen or monopropellant hydrazine or other propellant sufficient for spacecraft propulsion. Propellant Tanks  44  and  46  provide propellant sufficient for the lifetime of the spacecraft  40 . The size of propellant tanks  44  and  46  as a result is constrained by the expected propellant consumption over the lifetime of the spacecraft which in turn limits the volume and mass available for the high value added payload  48  that may be coupled to spacecraft  40 .  
         [0027]    Referring now to FIG. 2B, there is shown a client spacecraft  16 , typical of client spacecraft  16 ,  18 ,  20  and  22 , according to one embodiment of the present invention. Client spacecraft  16  has a spacecraft body  60 , and at least one propellant tank  62 ,  64 . The spacecraft body is adapted to be coupled to a payload  66 , typically on or adjacent to the earth face  68  of spacecraft body  60 . Payload  66  may be equipment for communications or imaging or other purposes typically found in orbiting unmanned spacecraft. The size of payload  66  is constrained principally by the size of propellant tanks  62  and  64 . Solar arrays  70  and  72  may provide power to batteries which in turn may support controllers, attitude and orbital control, momentum wheels, communications and other power based functions typical of unmanned spacecraft. Thrusters  74 , and  76  may provide thrust for minor maneuvers including minor north south station keeping, minor east west station keeping, and attitude control. Thrusters may be propellant based thrusters, electric thrusters or other types of thrusters. Thrusters  74  and  76  are supplied propellant from propellant tanks  62  and  64 . The propellant may be cold gas or monopropellant hydrazine and may be accompanied by electrical power for electric thrusters or other propellant sufficient for spacecraft propulsion. In an alternate embodiment where client spacecraft  16  would rely on servicing vehicle for substantially all significant maneuvers, thrusters  74  and  76  may not be provided. In a further alternate embodiment where client spacecraft  16  would rely on servicing vehicle for substantially all significant maneuvers, thrusters  74  and  76  may be provided for attitude control or for minor station keeping and used in conjunction with momentum wheels for attitude control, or momentum wheels may not be included on the spacecraft at all where momentum control is performed by the thrusters. Propellant Tanks  62  and  64  need not provide propellant sufficient for the lifetime of the spacecraft  16  as is required by the prior art spacecraft  40 . The size of propellant tanks  62  and  64  is constrained by the expected propellant consumption over the period of time where the servicing vehicle  14  is not available to support client spacecraft  16 , not by the lifetime of spacecraft  16  as in prior art spacecraft  40 . This in turn increases the volume and mass available for the high value added payload  66  that may be coupled to spacecraft  16  as compared to prior art spacecraft  40 . Client spacecraft has interfaces  78  shown located on the anti earth panel  82 . Although interfaces  78  are shown located on the anti earth panel  82 , they may alternately be located on any suitable panel or location. Interfaces  78  are adapted to re-usably interface with the servicing spacecraft  14  to enable the non-intrusive servicing and captive-carry maneuvers by the servicing vehicle  14 . Interface  78  may actually be a single interface point or multiple interface points.  
         [0028]    Referring now to FIG. 3 is a schematic perspective view of a client spacecraft  16  incorporating features of the present invention. Client spacecraft  16  has a spacecraft body  60 , and at least one propellant tank  62 . The spacecraft body is coupled to a payload  66  on the earth side  68  of spacecraft body  42 . Solar arrays  70  and  72  are provided. Thrusters  74 , and  76  may be provided. Client spacecraft  14  has interfaces  78  shown located on the anti earth panel  82 . Interfaces  78  may comprise coupling points  90 ,  92  and  94 , propellant interface  96 , power interface  98 , data communication interface  100  and/or stowed equipment deployment interface  102 . Interfaces  78  are adapted to re-usably interface with the servicing vehicle  14  to enable the non-intrusive servicing and captive-carry maneuvers by the servicing vehicle  14 . Although interfaces  78  are shown located on the anti earth panel  82 , any or all of the interfaces  78  may be located on one or more other panels or locations relative to client spacecraft  16 . Although interfaces  78  are broken down as shown, more or less interfaces between the client spacecraft  16  and the servicing vehicle  14  may be provided.  
         [0029]    Referring also to FIG. 4 is a schematic perspective view of a servicing vehicle  14  incorporating features of the present invention. Servicing vehicle  14  has a spacecraft body  120 , and at least one propellant tank  122 ,  124 . Servicing vehicle  14  is re-usably adapted for non-intrusive servicing of client spacecraft, and as a result, may utilize a large portion of its payload capability for propellant storage and/or stowed equipment deployment storage and/or other non-intrusive service related equipment storage. Solar arrays  126  and  128  are provided. Thrusters  130 ,  132  and  134  are provided. In an alternate embodiment, one or more thrusters used in combination with momentum wheels may be provided. Thrusters  130 ,  132  and  134  may be propellant based thrusters, electric thrusters or propulsion devices otherwise. Thrusters  130 ,  132  and  134  may be supplied propellant from propellant tanks  122  and  124 . The propellant may be cold gas or propellant such as monopropellant hydrazine and may involve electrical power for electric thrusters or other propellant sufficient for spacecraft propulsion. The servicing vehicle  14  utilizes thrusters  130 ,  132  and  134  to support each of the client spacecraft while it is docked with the client spacecraft for any or all significant maneuvers such as atmospheric drag compensation for low earth orbits (LEO), orbit raising maneuvers and north-south station keeping or east-west stationkeeping in geosynchronous orbits (GEO) and attitude control while the client spacecraft is docked to the servicing vehicle  14 . The servicing vehicle  14  may also reposition and/or reorient the client spacecraft  16  as required while docked. Repositioning and/or reorienting may be for the purposes of testing or other purposes such as antenna pattern characterizing as part of In Orbit Testing (IOT) or other IOT such as imaging mapping or other payload pattern mapping and characterizing or other wise. Where the servicing vehicle  14  is utilized, momentum wheels may not be included on the client spacecraft  16  at all where momentum control is performed by the thrusters on client spacecraft  16  or where the client spacecraft  16  relies on the servicing vehicle  14  for attitude control. Servicing vehicle  14  has interfaces  140  shown located on the earth side panel  142 . Interfaces  140  may comprise coupling points  144 ,  146  and  148 , propellant interface  150 , power interface  152 , data communication interface  154  and/or stowed equipment deployment interface  156 . Imaging equipment  160  may be included to visually inspect the client spacecraft  16 . Interfaces  140  are adapted to re-usably interface with the client spacecraft  16 ,  18 ,  20  and  22  to selectively enable the non-intrusive servicing and captive-carry maneuvers by the servicing vehicle  14 . Interfaces  140  is adapted to interface with interface  78  of the client spacecraft  16 . Interfaces  140  may also be adapted to re-usably interface with the orbiting depot  24  for re-fueling and servicing the servicing vehicle in order to re-usably support the propellant and non-intrusive servicing requirements of servicing multiple client spacecraft. In an alternate embodiment, a separate interface, in whole or in part, may be provided to allow the servicing vehicle  14  to interface with the orbiting depot  24 . Although interfaces  140  are shown located on the earth side panel  142 , any or all of the interfaces  140  may be located on one or more other panels or locations relative to servicing vehicle  14 . Although interfaces  140  are broken down as shown, more or less interfaces between the servicing spacecraft  14  and the client spacecraft  16  may be provided. Imaging equipment  160  are included to visually inspect the client spacecraft  16 .  
         [0030]    Coupling points  90 ,  92  and  94  of client spacecraft  16  are adapted to re-usably interface and couple with the coupling points  144 ,  146  and  148  respectively of servicing vehicle  14  to enable the captive-carry maneuvers by the servicing vehicle  14 . The coupling points may also be used to allow client spacecraft  16  to dock with servicing vehicle  14  during other non-intrusive servicing activities. In an alternate embodiment, client spacecraft  16  would not couple with servicing vehicle  14  during other non-intrusive servicing activities. The coupling points  90 ,  92  and  94  may selectively be coupled or de-coupled from coupling points  144 ,  146  and  148  respectively such as by a ball and socket arrangement where the socket may be articulated to selectively latch to the ball or release from the ball. In an alternate embodiment, any suitable coupling may be used to rigidly or semi rigidly couple servicing vehicle  14  to client spacecraft  16 . Although three coupling points are shown, more or less coupling points may be provided.  
         [0031]    Propellant interface  96  of the client spacecraft  16  is adapted to re-usably interface with the propellant interface  150  of the servicing vehicle  14  to enable the non-intrusive servicing related to propellant transfer from the servicing vehicle  14  to the client spacecraft  16 . In the embodiment shown, propellant interface  96  is coupled to tank  62  and propellant interface  150  is coupled to tank  122 . Propellant interface  96  may be adapted to re-usably couple with the propellant interface  150  and to make a positive seal to prevent leakage of propellant. In one embodiment, propellant interface  96  and  150  may be adapted to transfer propellant such as the monopropellant hydrazine. In an alternate embodiment, propellant interface  96  and  150  may be adapted to transfer any type of propellant, fuel or power suitable for use by the client spacecraft  16 .  
         [0032]    Power interface  98  of the client spacecraft  16  is adapted to re-usably interface with the power interface  152  of the servicing vehicle  14  to enable the non-intrusive servicing related to power transfer from the servicing vehicle  14 . In typical operation, power may be transferred to the client spacecraft  16  for battery re-conditioning, battery charging or simple power up in the case of the client spacecraft  16  not having a working battery on board, may not have a battery on board or may not have sufficient battery capacity on board for startup or otherwise. Power interface  98  and  152  may comprise electrical contacts allowing power such as electricity to pass through the electrically conductive contacts. In alternative embodiments, power may be transferred inductively, optically or otherwise.  
         [0033]    Imaging equipment  160  is used to visually inspect client spacecraft  16  to monitor its condition. Imaging equipment  160  may incorporate a camera, such as a CCD camera, memory to store images or image streams and a data transmission device to transmit the images to be viewed and analyzed. Imaging equipment  160  may further include a actuator driven mount or linkage to allow the attitude and/or the position of the camera to be changed relative to servicing vehicle  14  while docked with client spacecraft  16  or otherwise. Although imaging equipment  160  is shown as a single imaging source or interface, more or less such interfaces may be provided in alternate locations.  
         [0034]    Data communication interface  100  of the client spacecraft is adapted to re-usably interface with data communication interface  154  the servicing vehicle  14  to enable communication with the servicing vehicle  14 . In typical operation, data may be transferred to or from the client spacecraft  16  for operational monitoring, data down load or data upload to or from the client spacecraft. Data communication interfaces  100  and  154  may comprise electrical contacts allowing signals such as digital or analog electrical signals to pass through the electrically conductive contacts. In alternative embodiments, data may be transferred inductively, optically or otherwise. Data transfer may alternately take place when the servicing vehicle  14  and the client spacecraft  16  are physically separated by a large or short distance. Such a modest short distance could be, for example, approximately 100 meters. Although data communication interface  100  and data communication interface  154  are shown as a single interface, more or less such interfaces may be provided. Such an additional interface may include, for example, a communication interface for communicating data or otherwise with other spacecraft or earth and base control communications as in prior art spacecraft.  
         [0035]    Stowed equipment deployment interface  102  of the client spacecraft  16  is adapted to re-usably interface with the stowed equipment deployment interface  156  of the servicing vehicle  14  to enable the non-intrusive servicing associated with the deployment of stowed equipment. The stowed equipment that may be deployed may include solar arrays, magnetometer booms, solar radiation covers or shields, deployable radiators, antenna reflectors or other equipment that may be stowed on client spacecraft  16 . Although one stowed equipment deployment interface  102  is shown, there may be more or less stowed equipment deployment interfaces. Although one stowed equipment deployment interface  156  is shown, there may be more or less stowed equipment deployment interfaces. Deployment of stowed equipment typically consists of no material being exchanged between the client spacecraft  16  and the servicing vehicle  14 . Servicing vehicle  14  provides electrical power and/or torque and/or force to assist with the mechanical motion and mechanical power required drive the mechanisms and linkages associated with the deployment of stowed equipment on client spacecraft  16 . The deployment may be monitored using imaging equipment  160 . The use of the stowed equipment deployment interface enables the client spacecraft  16  to reduce the cost, volume and mass associated with deployment actuators that would have been on board client spacecraft  16  were there no reliance on servicing vehicle  14 .  
         [0036]    It should be understood that the foregoing description is only illustrative of the invention. Various alternatives and modifications can be devised by those skilled in the art without departing from the invention. Accordingly, the present invention is intended to embrace all such alternatives, modifications and variances which fall within the scope of the appended claims.